So I propose that we actually refuel the MCT in LEO (something we would already routinely do) for partially propulsive Earth Entry.
As Jim said earlier somewhere, Dragon is too small to have a rover.
From the last thread:So I propose that we actually refuel the MCT in LEO (something we would already routinely do) for partially propulsive Earth Entry.
Your fuel needs to get into LEO. If other MCTs on BFRs are launching fuel to LEO depots, then, pretty much by definition, those MCTs will need to be able to handle a full LEO reentry, so the TPS problem would need to be solved before you can use this to solve the program that you are trying to solve by using this. (So to speak.)
Why bother with the cargo?From the last thread:So I propose that we actually refuel the MCT in LEO (something we would already routinely do) for partially propulsive Earth Entry.
Your fuel needs to get into LEO. If other MCTs on BFRs are launching fuel to LEO depots, then, pretty much by definition, those MCTs will need to be able to handle a full LEO reentry, so the TPS problem would need to be solved before you can use this to solve the program that you are trying to solve by using this. (So to speak.)
Only if one tries to imagine that MCT is used as a Tanker to LEO, which is very silly and wasteful. Tankers will be a stretched upper stage of the BFR without any cargo on top and will use its low ballistic coefficient and retro-propulsion with residual propellents and likely some parachutes to perform re-entry and landing, all while delivering far MORE propellents.
Lots of people have been pushing this idea of MCT is the ONLY thing that BFR will ever have placed on top of it and that is must do EVERYTHING we want done from LEO all the way to Mars, this is completely unrealistic and dose not save any money as the MCT would be 10x harder to design and build when it has so many requirements put on it.While MCT will most likely be supplanted by a dedicated reusable tanker in the long run, in the short run its capabilities make it good enough for the task. Or a really simple, cheap, disposable tanker stage (a glorified fuel tank with a docking port and a single raptor engine).
Dang it, we really need one spot that has all known information about MCT that comes from SpaceX. I'm not sure if he said 80-100 or 50-100. Not that it makes an enormous difference, but it's annoying.
Dang it, we really need one spot that has all known information about MCT that comes from SpaceX. I'm not sure if he said 80-100 or 50-100. Not that it makes an enormous difference, but it's annoying.
MCT and BFR are supposed to be based on the same kind of platform. MCT will HAVE to have similar mass efficiency of an upper stage (in fact Musk has said MCT needs to be capable of Mars surface to earth in a SINGLE stage, though with far less payload), and essentially that's what it is. So whether you call the tanker a modified MCT or a stretched BFR upper stage may be a distinction without much difference.
MCT and BFR are supposed to be based on the same kind of platform. MCT will HAVE to have similar mass efficiency of an upper stage (in fact Musk has said MCT needs to be capable of Mars surface to earth in a SINGLE stage, though with far less payload), and essentially that's what it is. So whether you call the tanker a modified MCT or a stretched BFR upper stage may be a distinction without much difference.
Show me the quote for this, because I've never heard any such thing. Rather I think direct single stage Earth return is rather a possible (and the most aggressive possible) interpretation of some of Musks statements but it is far from set in stone.
And even if Elon had said this was his goal we should have SERIOUS doubts if such a goal would survive contact with real engineering as the vehicle capable of doing all that would put a single stage to Earth orbit vehicle
It will have to be a Mars SSTO in any case, which means 4.5 km/s of delta V minimum. That's very much in 2nd stage territory.MCT and BFR are supposed to be based on the same kind of platform. MCT will HAVE to have similar mass efficiency of an upper stage (in fact Musk has said MCT needs to be capable of Mars surface to earth in a SINGLE stage, though with far less payload), and essentially that's what it is. So whether you call the tanker a modified MCT or a stretched BFR upper stage may be a distinction without much difference.
Show me the quote for this, because I've never heard any such thing. Rather I think direct single stage Earth return is rather a possible (and the most aggressive possible) interpretation of some of Musks statements but it is far from set in stone.
All this talk about avoiding costs by not developing a 2nd stage are silly, SpaceX MUST have a use for the BFR other then launching for Mars related travel. The rocket would be completely useless for any other purpose if it's payloads were volumetricly constrained by needing to be inside a MCT cargo-hold which is likely no more then 500 m^3, SLS should have a payload fairing in excess of 2000 m^3.But what will use that excessive volumetric capability? The BA2100? Who would use a BA2100 and why? Would you ever need to launch multiple BA2100s? If so, for what reason? If you need something big put in space, wouldn't you rather design it according to the volumetric constraints of the vehicle you'll be using instead of the other way around?
A second attempt at a Mars Colonial Transporter. Not a first generation ship, but perhaps second or third generation of 10m core rockets. Youtube playlist should show 18 little videos covering the whole trip. A few very speculative items have crept in, for fun.
For some unknown reason, you need to restart the playlist after the first video. Sorry.
Michel Lamontagne
All this talk about avoiding costs by not developing a 2nd stage are silly, SpaceX MUST have a use for the BFR other then launching for Mars related travel. The rocket would be completely useless for any other purpose if it's payloads were volumetricly constrained by needing to be inside a MCT cargo-hold which is likely no more then 500 m^3, SLS should have a payload fairing in excess of 2000 m^3.
The quote is in the thread the moderators deleted for some reason.MCT and BFR are supposed to be based on the same kind of platform. MCT will HAVE to have similar mass efficiency of an upper stage (in fact Musk has said MCT needs to be capable of Mars surface to earth in a SINGLE stage, though with far less payload), and essentially that's what it is. So whether you call the tanker a modified MCT or a stretched BFR upper stage may be a distinction without much difference.
Show me the quote for this, because I've never heard any such thing. Rather I think direct single stage Earth return is rather a possible (and the most aggressive possible) interpretation of some of Musks statements but it is far from set in stone.
And even if Elon had said this was his goal we should have SERIOUS doubts if such a goal would survive contact with real engineering as the vehicle capable of doing all that would put a single stage to Earth orbit vehicle to shame. You can hand wave away the incredible difficulty and mass costs of EDL on Mars and Earth and the costs of keeping a vehicle alive during interplanetary transit.
All this talk about avoiding costs by not developing a 2nd stage are silly, SpaceX MUST have a use for the BFR other then launching for Mars related travel. The rocket would be completely useless for any other purpose if it's payloads were volumetricly constrained by needing to be inside a MCT cargo-hold which is likely no more then 500 m^3, SLS should have a payload fairing in excess of 2000 m^3.
With suitable mission kits MCT could perform the following missions:
- Tanker flights to LEO
- Propellant depot
- Satellite and space station (up to BA 2100 size at least) delivery
- Tourist launch to LEO (~300 passengers)
- Cargo/crew delivery to space stations anywhere in cis-lunar space
- Moon landings
- NEO visits.
MCT may not be the most efficient system to perform such missions, but it is capable enough.
What difference is there between MCT and a reusable upper stage? Just the habitable portion on top. They will need similar performance (~6.5-7km/s). Both need reentry and landing capability (legs, etc).
The problem with Shuttle is there were only a few of them made, no custom ones. With MCT, thousands will be made, so no problem making some that lack the habitable portion or that act as tankers or that are only used for cargo. The requirements for these things are similar but the MCTs can be modified to fit the purpose instead of having one vehicle type do everything at once.
Yes, there is a spreadsheet here:A second attempt at a Mars Colonial Transporter. Not a first generation ship, but perhaps second or third generation of 10m core rockets. Youtube playlist should show 18 little videos covering the whole trip. A few very speculative items have crept in, for fun.
For some unknown reason, you need to restart the playlist after the first video. Sorry.
Michel Lamontagne
A lot of thought and work has gone into this. The engineering of an MCT like this would be formidably difficult, but it addresses concerns about zero gravity and abort that other conceptual designs do not.
One improvement might be to have the capsule part of the MCT nominally land attached to the cargo/transit part. Then it can perform an abort during landing.
Do you have mass estimates to go with the animations?
What difference is there between MCT and a reusable upper stage? Just the habitable portion on top. They will need similar performance (~6.5-7km/s). Both need reentry and landing capability (legs, etc).
The problem with Shuttle is there were only a few of them made, no custom ones. With MCT, thousands will be made, so no problem making some that lack the habitable portion or that act as tankers or that are only used for cargo. The requirements for these things are similar but the MCTs can be modified to fit the purpose instead of having one vehicle type do everything at once.
I don't know about THOUSANDS of MCT's being built, dozens and maybe hundreds, but by that time, I expect newer designs will superceed the current design scheme.
All this talk about avoiding costs by not developing a 2nd stage are silly, SpaceX MUST have a use for the BFR other then launching for Mars related travel. The rocket would be completely useless for any other purpose if it's payloads were volumetricly constrained by needing to be inside a MCT cargo-hold which is likely no more then 500 m^3, SLS should have a payload fairing in excess of 2000 m^3.
I disagree. Estimates in other threads have about 22 m^3 per person of pressurised volume, so for 100 passengers that is 2200 m^3. If the crew accommodations are payload to MCT, this means that the payload volume would need to be 2500 m^3 or above.
50 reuses is far too many. My guess is the MCTs may last about 3 decades, one reuse every ~2 years (every synod), so only 12-15 reuses is practical.What difference is there between MCT and a reusable upper stage? Just the habitable portion on top. They will need similar performance (~6.5-7km/s). Both need reentry and landing capability (legs, etc).
The problem with Shuttle is there were only a few of them made, no custom ones. With MCT, thousands will be made, so no problem making some that lack the habitable portion or that act as tankers or that are only used for cargo. The requirements for these things are similar but the MCTs can be modified to fit the purpose instead of having one vehicle type do everything at once.
I don't know about THOUSANDS of MCT's being built, dozens and maybe hundreds, but by that time, I expect newer designs will superceed the current design scheme.
Sharing the development costs between a large number of ships is a key requirement. A 10 billion dollar development cost over 100 ships is 'only' 100 million dollars per ship, and if each ship can do 50 trips, then it's 2 million $ per trip, and a small portion of the 50 million dollar fare per ship. So perhaps 100 ships per design generation?
I was thinking of reuse for the first stage and cargo modules to Earth orbit. I agree the MCT itself will not have so many runs. Thousands of ships is fine with me, the important concept is the reduction of cost by using large production numbers.50 reuses is far too many. My guess is the MCTs may last about 3 decades, one reuse every ~2 years (every synod), so only 12-15 reuses is practical.What difference is there between MCT and a reusable upper stage? Just the habitable portion on top. They will need similar performance (~6.5-7km/s). Both need reentry and landing capability (legs, etc).
The problem with Shuttle is there were only a few of them made, no custom ones. With MCT, thousands will be made, so no problem making some that lack the habitable portion or that act as tankers or that are only used for cargo. The requirements for these things are similar but the MCTs can be modified to fit the purpose instead of having one vehicle type do everything at once.
I don't know about THOUSANDS of MCT's being built, dozens and maybe hundreds, but by that time, I expect newer designs will superceed the current design scheme.
Sharing the development costs between a large number of ships is a key requirement. A 10 billion dollar development cost over 100 ships is 'only' 100 million dollars per ship, and if each ship can do 50 trips, then it's 2 million $ per trip, and a small portion of the 50 million dollar fare per ship. So perhaps 100 ships per design generation?
Musk has said 80,000 people per year (and ten times as many cargo shipments), which is 1000 Passenger MCTs at once, plus 10,000 cargo MCTs (or actually, there ways around this, but it remains to be seen if they're worth it). So yeah, at any one time, there would need to be thousands of MCTs.
Lots of people have been pushing this idea of MCT is the ONLY thing that BFR will ever have placed on top of it and that is must do EVERYTHING we want done from LEO all the way to Mars, this is completely unrealistic and dose not save any money as the MCT would be 10x harder to design and build when it has so many requirements put on it.While MCT will most likely be supplanted by a dedicated reusable tanker in the long run, in the short run its capabilities make it good enough for the task. Or a really simple, cheap, disposable tanker stage (a glorified fuel tank with a docking port and a single raptor engine).
The idea that early MCTs be used as both a tanker vehicle and an MCT stems from a very real and present fact that SpaceX does not have infinite money and thus cannot really afford development and manufacturing of multiple different reusable, earth-landable and rapidly reusable vehicle designs.
Just make one that is good enough and build as many as you can. An MCT without the cargo will do just fine for refuelling. Not perfect, but good enough.
All this talk about avoiding costs by not developing a 2nd stage are silly, SpaceX MUST have a use for the BFR other then launching for Mars related travel. The rocket would be completely useless for any other purpose if it's payloads were volumetricly constrained by needing to be inside a MCT cargo-hold which is likely no more then 500 m^3, SLS should have a payload fairing in excess of 2000 m^3.
What difference is there between MCT and a reusable upper stage? Just the habitable portion on top. They will need similar performance (~6.5-7km/s). Both need reentry and landing capability (legs, etc).
The problem with Shuttle is there were only a few of them made, no custom ones. With MCT, thousands will be made, so no problem making some that lack the habitable portion or that act as tankers or that are only used for cargo. The requirements for these things are similar but the MCTs can be modified to fit the purpose instead of having one vehicle type do everything at once.
I do expect the tanker to be different. No payload or crew quarters. Just stretched main tanks. That's a lot more mass efficient. But early on for the first few missions or in a test phase they may use MCT for that purpose too.
Stretching probably does make sense. At least, once we're talking hundreds of MCTs per synod.
Stretching probably does make sense. At least, once we're talking hundreds of MCTs per synod.
Do we have any authoritative optimal fuel:oxidizer mass ratio estimates for FFSC methalox?Stretching probably does make sense. At least, once we're talking hundreds of MCTs per synod.
The volume of the fuel for a given mass is much less than the volume required for the crew, so a tanker stage should always be much shorter than an MTC stage.
Liquid methane is only 2.3 m3 per tonne and oxygen is 0,84 m3 per tonne, so 150 tonnes of propellant in a 3.4 ratio is 110 tons of oxygen (50 m3) and 40 tons of methane (100 m3) or about 150 m3 total. For a 10m core, 5^2xpi= 78m3 per m of length, it's barely 2 meters of length. Of course other types of cargo can take up more volume.
But what will use that excessive volumetric capability? The BA2100? Who would use a BA2100 and why? Would you ever need to launch multiple BA2100s? If so, for what reason? If you need something big put in space, wouldn't you rather design it according to the volumetric constraints of the vehicle you'll be using instead of the other way around?
In any case, the bread and butter of most commercial launch service companies is and has always been communication satellites.
A cargo bay which can hold 100 tonnes of cargo for mars can most definitely hold a comms satellite, and the excessive delta V that a MCT is required to pull (even/especially if it doesn't act as its own 2nd stage or do a one burn from Mars to Earth since it still needs at least 4.5 km of delta V to rendezvous with a transfer tug in LMO) make it more than capable of acting as a GTO delivery vehicle.
The volume constraint argument is a red herring. If something big enough to fill the volume constraints of a SLS fairing comes along and requires a launch it might as well get a stage specifically designed for it or even the SLS, if that ever goes into commercial launches. You design things based on the constraints you are given, not the other way around.
I do expect the tanker to be different. No payload or crew quarters. Just stretched main tanks. That's a lot more mass efficient. But early on for the first few missions or in a test phase they may use MCT for that purpose too.
Why stretched main tanks?
The lighter that configuration of MCT would be without any crew accomodations, the more propellant it has left when arriving in LEO. No need to stretch the tanks. Plus better to keep the basic tank/skin/structure common over all variants, I would think.
Why do you differentiate between Earth reentry and Mars reentry? I like most people assume it is one and the same. Some disagree, I know.
Why do you differentiate between Earth reentry and Mars reentry? I like most people assume it is one and the same. Some disagree, I know.
No these aren't comparable at all, in fact their are perhaps 6 different entry scenarios. http://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20040086716.pdf
Mars from LMO 3.5 kms
Mars from Interplanetary Slow 6 kms
Earth from LEO 7.5 kms
Mars from Interplanetary Fast 8 kms
Earth from Interplanetary Slow 12 kms !
Earth from Interplanetary Fast 14 kms !!
The hardest reentry is back to earth. If it can do that, Mars speed is no problem.
At these kinds of packaging densities you would need 1200-5000 m^3 to use that mass effectively for launching satellites. But their is no way the vehicle can have such a huge cargo hold, it would make the overall vehicle too large and require too much structural mass to make it survive re-entry.This is an entirely wrong assumption. An Ariane V, per your words, has ~390 cubic meters of volume and can manage to bring two satellites into GTO and costs about $200 million per launch, at a price of $100 million per satellite. A falcon 9 has ~275 cubic meters and can get a single satellite into GTO in fully expendable mode at a price of $90 million per launch/satellite.
Commercial communication satellites are LOW DENSITY, just look at the size of current payload fairing and you can see that their is no way you could put 100 mT of satellites into the kind of volumes were looking at for a MCT cargo-hold.You would never pack 100 metric tonnes' worth of satellites in a MCT.
Falcon 9 payload fairing has a volume of ~275 m^3 and it launches only 5 mT to GTO, Ariane 5 has ~390 m^3 and launches a maximum of 12 mT to GTO. Shuttle had ~300 m^3 payload bay and could carry 24 mT to LEO.
At these kinds of packaging densities you would need 1200-5000 m^3 to use that mass effectively for launching satellites. But their is no way the vehicle can have such a huge cargo hold, it would make the overall vehicle too large and require too much structural mass to make it survive re-entry.
Volume is VERY important, MANY space launch systems face volume limitations, Dragon capsule for example is volume rather then mass limited for most cargoes that need to be launched to ISS.
But it CAN'T that's the rub, the vehicle can't be the large low density tank people are imagining, it would be crushed. The Apollo heat shield alone was 15% of the mass, the structure was 27% and this was to for a compact and easy to protect shape. So we can't just wave our arms and say MCT will be able to do this.
How much retro-propulsion do you think we can do upon return to Earth? Any propellent to do this with is added to our DeltaV from Mars surface which is already 6-7 kms for a direct Earth return. The propellent fraction is already near the limits of credibility.
How much retro-propulsion do you think we can do upon return to Earth? Any propellent to do this with is added to our DeltaV from Mars surface which is already 6-7 kms for a direct Earth return. The propellent fraction is already near the limits of credibility.
Using a BFR and MCT to launch satellites is comparable to using the Queen Mary as an ore barge.I'm sure you would be better off using the Queen Mary to haul ore than throwing away part of an ore ship each time. You can always build a freighter version, just like the cargo 747 or whatever.
We should also acknowledge that and move on.
It will also refuel at least twice, so Apollo-like 'missions' to Mars are off the table... for SpaceX anyway.
This is why it is not valid to design spacecraft by only looking at Delta-V and tank sizes and imagining that a giant 2nd stage can do the job of direct Earth return from Mars surface just because it can hold the propellents to launch to Earth. It would literally be crushed like an empty beer can against ones forehead when it hits the Earth's atmosphere.
Try crushing a beer can containing several bars of pressure against your forehead.
No. Retropulsion before reentry on Earth isn't going to happen with chemical propulsion. It's always better to improve the heatshield.
Mass penalty is way too high. Increasing the required delta-v for the Mars stage from 7-7.5km/s to ~10km/s would double the required mass, and that's assuming no increase in tankage mass.
Just no.
PICA-X is crazy awesome stuff. Just use a little more of it.
If the MCT is completely refueled on Mars for launch back to earth, all of the fuel will not be needed to get the TEI burn. There will be quite a bit of fuel left. So the MCT could fire and slow down before earth reentry, or it could slow down using aerocapture for a few orbits to slow down. Also, there will probably be only a crew return, not 100 people.
Yeah but,..The wear on the PICA-X is infinitesimal compared to the propellant needed to prevent it. If you already have a heatshield, you should be maximizing it's use when possible. For Mars entry, you'll still need retropropulsion while supersonic, but for Earth you should be subsonic before you start the landing burn.
Any chance they would use both. We have seen the results of firing an engine in a braking burn. (F-9 booster recovery attempts) Also they showed computer modeling of hypersonic reentry with engine firing. ISTM the results show the engine pushes the superheated plasma from reentry, away from the spacecraft.
So say they use an appropriately sized engine to push the plasma away from the spacecraft, but the majority of the braking is done by the atmosphere. Would lead to less wear on the pica-X heat shield and is reusable by refueling the engine.
Try crushing a beer can containing several bars of pressure against your forehead.
Or better - don't. :)
The hardest reentry is back to earth. If it can do that, Mars speed is no problem.
But it CAN'T that's the rub, the vehicle can't be the large low density tank people are imagining, it would be crushed. The Apollo heat shield alone was 15% of the mass, the structure was 27% and this was to for a compact and easy to protect shape. So we can't just wave our arms and say MCT will be able to do this.
How much retro-propulsion do you think we can do upon return to Earth? Any propellent to do this with is added to our DeltaV from Mars surface which is already 6-7 kms for a direct Earth return. The propellent fraction is already near the limits of credibility.
I think that is what he was referring to, and empty MCT. Empty of fuel and cargo, thus being much lighter and easier to crush when slamming into the earths atmosphere, even with a good heat shield. Thus the idea of a retro burn to slow down before entering earths atmosphere. Also, coming from Mars one MCT at a time, an orbit or two would help in enabling a more precise landing at the launch site, instead of the middle of the ocean or the jungle somewhere.
If the MCT is completely refueled on Mars for launch back to earth, all of the fuel will not be needed to get the TEI burn. There will be quite a bit of fuel left. So the MCT could fire and slow down before earth reentry, or it could slow down using aerocapture for a few orbits to slow down. Also, there will probably be only a crew return, not 100 people.
Crew?
Why?
MCT should be able to return to Earth empty. (And as needed provide occasional return transport for humans needing to return)
On the way out assuming several 10s of passengers it would be astounding if there were not several engineers capable of specialized training as "flight engineers" to repair anything repairable by humans. No astronaut corps test pilots needed, just FEs similar to on the shuttle.
Crew mass & life support is wasted mass and money.
Atlas stages aren't weak when empty, they're weak when not pressurized. If MCT isn't pressurized, there would be other, bigger problems to worry about.I think that is what he was referring to, and empty MCT. Empty of fuel and cargo, thus being much lighter and easier to crush when slamming into the earths atmosphere, even with a good heat shield. Thus the idea of a retro burn to slow down before entering earths atmosphere. Also, coming from Mars one MCT at a time, an orbit or two would help in enabling a more precise landing at the launch site, instead of the middle of the ocean or the jungle somewhere.
Being empty will not make the structure weaker. MCT isn't an Atlas fuel tank.
While machines are pretty good at doing their jobs, nobody will want to risk a several billion dollar investment on the possibility that a IC chip will fry because of a stray cosmic ray and send the craft wandering off into space or worse, come c rashing down on Earth at 6 to 7 KMS.
If MCT isn't pressurized, there would be other, bigger problems to worry about.
The hardest reentry is back to earth. If it can do that, Mars speed is no problem.
But it CAN'T that's the rub, the vehicle can't be the large low density tank people are imagining, it would be crushed. The Apollo heat shield alone was 15% of the mass, the structure was 27% and this was to for a compact and easy to protect shape. So we can't just wave our arms and say MCT will be able to do this.
How much retro-propulsion do you think we can do upon return to Earth? Any propellent to do this with is added to our DeltaV from Mars surface which is already 6-7 kms for a direct Earth return. The propellent fraction is already near the limits of credibility.
It would literally be crushed like an empty beer can against ones forehead when it hits the Earth's atmosphere.
At these kinds of packaging densities you would need 1200-5000 m^3 to use that mass effectively for launching satellites. But their is no way the vehicle can have such a huge cargo hold, it would make the overall vehicle too large and require too much structural mass to make it survive re-entry.This is an entirely wrong assumption. An Ariane V, per your words, has ~390 cubic meters of volume and can manage to bring two satellites into GTO and costs about $200 million per launch, at a price of $100 million per satellite. A falcon 9 has ~275 cubic meters and can get a single satellite into GTO in fully expendable mode at a price of $90 million per launch/satellite.
Therefore, the only thing a MCT needs to do in order to be competitive as a satellite delivery platform is lower the cost of getting satellites into GTO, not increase the amount of satellites into GTO and since the goal of the MCT is to be rapidly reusable, a "gas & go" type of system, the only costs involved with the launch would be processing, operation and the cost of the methalox rather than building an entire vehicle, the MCT is cheaper and if the internal volume of its cargo bay is only equal to that of an Ariane V fairing, it can still get two satellites into GTO so the only way a MCT is not competitive is if the cost of its launch approaches the $200 million mark.
And given the MCT's much touted price tag of $50 million per 100 people ($500 000 per passenger) I just don't see that as a likely event.Commercial communication satellites are LOW DENSITY, just look at the size of current payload fairing and you can see that their is no way you could put 100 mT of satellites into the kind of volumes were looking at for a MCT cargo-hold.You would never pack 100 metric tonnes' worth of satellites in a MCT.
Falcon 9 payload fairing has a volume of ~275 m^3 and it launches only 5 mT to GTO, Ariane 5 has ~390 m^3 and launches a maximum of 12 mT to GTO. Shuttle had ~300 m^3 payload bay and could carry 24 mT to LEO.
At these kinds of packaging densities you would need 1200-5000 m^3 to use that mass effectively for launching satellites. But their is no way the vehicle can have such a huge cargo hold, it would make the overall vehicle too large and require too much structural mass to make it survive re-entry.
Volume is VERY important, MANY space launch systems face volume limitations, Dragon capsule for example is volume rather then mass limited for most cargoes that need to be launched to ISS.
Mostly because an MCT without refuelling can do much less than that in tonnes to GTO. (I remember that a number of 10-15 tonnes was mentioned somewhere in the previous topic, which amounts to 2 conventional satellites or 3-4 SEP ones, so about an Ariane 5 worth of payload to GTO). The reason why you'd want to do this without refuelling is to lower operating costs and complexity of the mission.
So you pack it with as many satellites as you can given the volume and mass constraints and that still makes the volume of the MCT a non-issue because it's still cheaper than the alternatives and has enough of a capability for at least 2 satellites.
The only time when the volume of a MCT becomes an issue is if you consider a commercial depot stationed in LEO, to which the MCT could deliver its full payload in mass and from which tugs would take over. But even then it'd still be cheaper
Using a BFR and MCT to launch satellites is comparable to using the Queen Mary as an ore barge.
We should also acknowledge that and move on.
It will also refuel at least twice, so Apollo-like 'missions' to Mars are off the table... for SpaceX anyway.
That brutal 12-14 kms entry to Earth is something everyone who is talking about this direct Earth return is glossing over, that is beyond Apollo speeds, the only thing that can survive that kind of heat, dynamic pressure and g-force is a dense capsule with thick heavy ablatives.
This is why it is not valid to design spacecraft by only looking at Delta-V and tank sizes and imagining that a giant 2nd stage can do the job of direct Earth return from Mars surface just because it can hold the propellents to launch to Earth. It would literally be crushed like an empty beer can against ones forehead when it hits the Earth's atmosphere.
A 2nd stage that can return from Earth orbit is a vastly simpler thing to do because the speed is half (and the energy is a quarter), and it is fairly easy to slow down the 2nd stage by several kms with residual propellents, and to employ disposable things like parachutes because it only needs to perform ONE landing before servicing rather then two, and lastly it can be made much less reliable in landing because it's unmanned, no one dies horribly if it crashes or burns up on reentry unlike MCT.
50 reuses is far too many. My guess is the MCTs may last about 3 decades, one reuse every ~2 years (every synod), so only 12-15 reuses is practical.What difference is there between MCT and a reusable upper stage? Just the habitable portion on top. They will need similar performance (~6.5-7km/s). Both need reentry and landing capability (legs, etc).
The problem with Shuttle is there were only a few of them made, no custom ones. With MCT, thousands will be made, so no problem making some that lack the habitable portion or that act as tankers or that are only used for cargo. The requirements for these things are similar but the MCTs can be modified to fit the purpose instead of having one vehicle type do everything at once.
I don't know about THOUSANDS of MCT's being built, dozens and maybe hundreds, but by that time, I expect newer designs will superceed the current design scheme.
Sharing the development costs between a large number of ships is a key requirement. A 10 billion dollar development cost over 100 ships is 'only' 100 million dollars per ship, and if each ship can do 50 trips, then it's 2 million $ per trip, and a small portion of the 50 million dollar fare per ship. So perhaps 100 ships per design generation?
Musk has said 80,000 people per year (and ten times as many cargo shipments), which is 1000 Passenger MCTs at once, plus 10,000 cargo MCTs (or actually, there ways around this, but it remains to be seen if they're worth it). So yeah, at any one time, there would need to be thousands of MCTs.
Do we have a propellant depot in LEO?
I also take it as a given that at least initially many MCTs would be staying on (or at) Mars at least for a few years. They might be a source of spare engines too if some didn't pass inspection on Mars.
This would suggest that although we intend to ramp up cargo and personnel flights to Mars with each successive launch season, we could do it with a steady state of MCT production, say one a month or less. That makes planning production simple.
Now as soon as we have a few spare MCT's hanging around in LEO waiting for launch season, someone else might want to lease a few to start up lunar ISRU for fuel and maybe something else.
Do we have a propellant depot in LEO?
It is part of the plan laid out by Elon Musk. However for the first few conjunctions when only 2-4 flights go to Mars, it may be easier to just refuel directly in LEO. When the number of flights increases, depots will soon become necessary.
Edit: I could imagine that a manned MCT would be refuelled draining a full cargo MCT that would then be fuelled up for a second time. That may count as a kind of depot and avoid a delay for the crew.
I also take it as a given that at least initially many MCTs would be staying on (or at) Mars at least for a few years. They might be a source of spare engines too if some didn't pass inspection on Mars.
Nothing to have a disagreement on but I believe only the first 2 or 3 would stay, probably forever. I believe it is safer to send them back after unloading and only a few weeks stay on Mars rather than having them there for a full synod and then relying on their continued function and safety without means for a thorough inspection.
This would suggest that although we intend to ramp up cargo and personnel flights to Mars with each successive launch season, we could do it with a steady state of MCT production, say one a month or less. That makes planning production simple.
Now as soon as we have a few spare MCT's hanging around in LEO waiting for launch season, someone else might want to lease a few to start up lunar ISRU for fuel and maybe something else.
Sounds good. I would guess though that it will be some time until production rate is ramped up to one a month. Depending on funds.
This discussion gets me to another thought. I had anticipated cargo flights might not be pressurized. But they will likely need to be pressurized, not only for the benefit of the cargo but to give them stability.
Getting slightly OT, I wonder if equipment will have to be pressurized all the way, which would make unloading quite difficult. Or if it could be exposed to the near vacuum of Mars for a short time during unloading. Some equipment would be designed to work on the surface, no problem there. But a lot of stuff would go into habitats.
If the MCT is completely refueled on Mars for launch back to earth, all of the fuel will not be needed to get the TEI burn. There will be quite a bit of fuel left. So the MCT could fire and slow down before earth reentry, or it could slow down using aerocapture for a few orbits to slow down. Also, there will probably be only a crew return, not 100 people.
Crew?
Why?
MCT should be able to return to Earth empty. (And as needed provide occasional return transport for humans needing to return)
On the way out assuming several 10s of passengers it would be astounding if there were not several engineers capable of specialized training as "flight engineers" to repair anything repairable by humans. No astronaut corps test pilots needed, just FEs similar to on the shuttle.
Crew mass & life support is wasted mass and money.
You may be correct aboutthe need for a crew, but people are still a bit primitive. They'd be more than a bit nervous to trust their lives to nothing more than machines. I'm pretty sure that they'd want at least a minimal crew orf pilot, navigator/copilot and at least one engineer. (A dedicated doctor/medic would also likely be a good idea). If nothing else, to keep the passengers calm during the flight.
While machines are pretty good at doing their jobs, nobody will want to risk a several billion dollar investment on the possibility that a IC chip will fry because of a stray cosmic ray and send the craft wandering off into space or worse, come c rashing down on Earth at 6 to 7 KMS.
At the end of they day, F9R and FHR will be doing the vast bulk of the sat launches.
At the end of they day, F9R and FHR will be doing the vast bulk of the sat launches.
Surely once a reusable BFR is flying, F9/FH will be retired? Simplify to one engine line, one tank line, one type of launch infrastructure, etc. Reduces cost. (Especially if BFR is a single core and cheaper to integrate than a triple core FH.)
[I would think the customers would end up forcing the decision. In much the same way that few were interested in F1 when they could fly as a secondary payload on F9 for half the price.]
That brutal 12-14 kms entry to Earth is something everyone who is talking about this direct Earth return is glossing over, that is beyond Apollo speeds, the only thing that can survive that kind of heat, dynamic pressure and g-force is a dense capsule with thick heavy ablatives.
This is why it is not valid to design spacecraft by only looking at Delta-V and tank sizes and imagining that a giant 2nd stage can do the job of direct Earth return from Mars surface just because it can hold the propellents to launch to Earth. It would literally be crushed like an empty beer can against ones forehead when it hits the Earth's atmosphere.
A 2nd stage that can return from Earth orbit is a vastly simpler thing to do because the speed is half (and the energy is a quarter), and it is fairly easy to slow down the 2nd stage by several kms with residual propellents, and to employ disposable things like parachutes because it only needs to perform ONE landing before servicing rather then two, and lastly it can be made much less reliable in landing because it's unmanned, no one dies horribly if it crashes or burns up on reentry unlike MCT.
The thing is....this is all a moot point to your argument. Whether MCT is it's own 2nd stage to LEO or it sits atop a dedicated unique 2nd stage won't change the fact that MCT will need to have -large- tanks. It will be mostly propellant tank by volume just to do the TMI burn and Mars EDL retropropulsion....and to get itself off the surface of Mars, even if it were only going to LMO before getting refueled there rather than all the way back in one shot.
It will be mostly a large propellant tank, with some legs, engines, and some cargo or hab internal volume.
Having MCT be it's own 2nd stage rather than having a separate dedicated 2nd stage won't change that. MCT can't then become just a simple larger Dragon capsule.
With a dedicated reusable 2nd stage, then it's just a giant 3rd stage, rather than a giant 2nd stage. Maybe a little smaller. Not much else changes. So it's a bit of a moot argument.
Yes, designing a vehicle that is returning just form LEO is vastly more simple than designing one that's coming back from interplanetary speeds. You are correct. But, there's not an either/or option. SpaceX must figure out how to get a large rocket stage back from Mars and land it on Earth. They already need to solve that long pole issue. A vehicle they design to handle that, can already return to Earth from LEO easily enough, without the [easy] development of a separate LEO only vehicle even necessary.
Surely once a reusable BFR is flying, F9/FH will be retired? Simplify to one engine line, one tank line, one type of launch infrastructure, etc. Reduces cost. (Especially if BFR is a single core and cheaper to integrate than a triple core FH.)
....I can't imagine they have any plans of retiring FH or F9 once MCT starts flying. For several reasons.
When you have a hammer, everything looks like a nail. But not everything is a nail. it's...a matter of the right tool for the job...It's hard to imagine them launching a big Saturn V size (or larger) LV for such a comsat, even if it's fully reusable.
Surely once a reusable BFR is flying, F9/FH will be retired? Simplify to one engine line, one tank line, one type of launch infrastructure, etc. Reduces cost. (Especially if BFR is a single core and cheaper to integrate than a triple core FH.)[...] Semi trucks haul mail cross country because that's the most cost efficient truck for that task. Jeep sized vehicles deliver to local mailboxes because that is the most cost efficient vehicle for that task. You don't use a maul when driving a finish nail. You use a finish nail hammer. Vice versa when demolishing a wall. You use the right tool for the job, and there is no single tool that does every job.
I can't imagine a client wanting to use a semi-expendable rocket when a cheaper secondary/tertiary payload slot is available on a fully reusable HLV. People seem hung up over the size, all that matters is the price.
(Those semis will often carry many small packages because the per-item cost is lower than carrying them individually in a smaller vehicle.)
500 m^3 is a reasonable cargo-hold but their would be no integral habitat as many have speculated, it makes much mores sense to load a large module into the cargo-hold which can be removed and left on the Mars surface to minimize the return mass. This also has the advantage of eliminating separate crew and cargo variants.
It is also extremely inefficient in structural mass. MCT is all about efficiency in structural mass.
The most mass efficient thing is to NOT make it integral to the MCT. The time when structural mass efficiently maters most is take off, and if we indent to me offloading people and not taking them back to Earth then all that habitat mass would be pure dead-weight on take off.
A single MCT won't be billions of dollars of investment or the $500k per passenger figure will be impossible. An order of magnitude less like $100-400 million. Also, electronics can easily be made reliable enough. We have lots an lots of experience running spacecraft for years at a time without maintenance. 6-9 months won't be a challenge for a company that will launch thousands of satellites into LEO.
Of course, transit time is just ~3 months, not 8, so that's a big difference right there. Additionally, they likely did not sleep in shifts like you would on MCT colonization runs.
MCT is a radical design. Acknowledge that and move on.
Musk’s $500,000 ticket price for a Mars trip was derived from what he thinks is affordable.
"The ticket price needs to be low enough that most people in advanced countries, in their mid-forties or something like that, could put together enough money to make the trip," he said, comparing the purchase to buying a house in California. [Photos: The First Space Tourists]
He also estimated that of the eight billion humans that will be living on Earth by the time the colony is possible, perhaps one in 100,000 would be prepared to go. That equates to potentially 80,000 migrants.
Musk figures the colony program — which he wants to be a collaboration between government and private enterprise — would end up costing about $36 billion. He arrived at that number by estimating that a colony that costs 0.25 percent or 0.5 percent of a nation’s gross domestic product (GDP) would be considered acceptable.
The United States' GDP in 2010 was $14.5 trillion; 0.25 percent of $14.5 trillion is $36 billion. If all 80,000 colonists paid $500,000 per seat for their Mars trip, $40 billion would be raised.
"Some money has to be spent on establishing a base on Mars. It’s about getting the basic fundamentals in place," Musk said. "That was true of the English colonies [in the Americas]; it took a significant expense to get things started. But once there are regular Mars flights, you can get the cost down to half a million dollars for someone to move to Mars. Then I think there are enough people who would buy that to have it be a reasonable business case."
Again, while SpaceX is comprised of brilliant engineers and has an insane(ly dedicated) leader, it isn't made entirely out of money, nor does it have infinite time. Making and improving on the MCT will take up a lot of their money and time (in fact I'm pretty sure all of it) which they won't be able to spend on reusable BFR upper stage design.
I can't imagine a client wanting to use a semi-expendable rocket when a cheaper secondary/tertiary payload slot is available on a fully reusable HLV. People seem hung up over the size, all that matters is the price.
(Those semis will often carry many small packages because the per-item cost is lower than carrying them individually in a smaller vehicle.)
People seem to be trying to compare BFR cost to launch satellites with other existing or near term launchers (Ariane 5, SLS, F9 etc etc). But that is not what my argument is about, it is about which VISION FOR BFR is cheaper to launch satellites with. Musk's goal is not to slightly undercut existing launchers, it is to create massive paradigm-shifting reductions in $ to LEO, even if BFR launch blows every other launch vehicle out of the water it still needs to compete with variants of itself.
I'm arguing that BFR with mostly normal reusable 2nd stage is better at launching satellites then BFR with the giant MCT combo 2nd stage. These latter will cost LESS because a 2nd stage even reusable is a much simpler and lower mass vehicle then the whole MCT which has MUCH more demanding requirements on lifespan, reentry heat, lifespan etc etc.
The first stage is identical and presumably all other logical and launch related costs are too, so the only difference is in the 2nd stages, one which is conventional with a voluminous payload fairing which is light and designed for optimal mass delivery to orbit, the other is a huge heavy vehicle totally over engineered for this job and has a small cargo hold.
Their is no contest the normal 2nd stage will out perform the giant vehicle to any orbit, just as an EELV booster outperformed the shuttle at launching satellites. And the normal 2nd stage is going to be vastly cheaper to develop as well, the only argument that anyone has left is that because Elon absolutely MUST have his Mars oriented vehicle he will choose to shoehorn it into every possible usage even for things it is not optimized for so as to amortize the cost over as many flights as possible. I don't recall that strategy working well for Shuttle.
Musk is blessed with inordinate patience, he could have blown his money on a stunt ages ago but has always focused on building a viable BUSINESS first and foremost. In pursuit of greater revenue he is now getting into the satellite business. He doesn't leave any potential revenue source on the table. Se Musk is not going to pass on designing the best conventional satellite launcher simply because he also wants to use the vehicle for Mars adventures, he knows that it must be a viable vehicle in it's own right.
And please stop repeating that big-tank are the ONLY way, I have shown you several times how the vehicle can designed with much smaller tanks, your not a fan of these options but it is dishonest to begin your argument with your preferred solution as the only option, it is simply begging the question.
First off you can go to LMO and then dock with a transit vehicle like Mars Semi-Direct, no one here can claim that they are unfamiliar with Semi-Direct. Second, inflatable tanks in the cargo-hold, even rigid tanks in the cargo hold if you think inflatables are to low TRL.
And back to design concepts, it is the most fun :)
What if BFR and MCT will be the essentially same thing?
Kind of universal module sized similar to S-IC, 9 Raptors at the bottom, with lending legs, around 1900mt.
Than all thing will be three core (I know, Elon said one-core, but look at F5), MCT in the center will have less fuel load replaced by cargo bay and improved thermal protection, and probably have less than 9 Raptors(3?).
Just a thought to standardize and reduce cost...
Try crushing a beer can containing several bars of pressure against your forehead.
And back to design concepts, it is the most fun :)
What if BFR and MCT will be the essentially same thing?
Kind of universal module sized similar to S-IC, 9 Raptors at the bottom, with lending legs, around 1900mt.
Than all thing will be three core (I know, Elon said one-core, but look at F5), MCT in the center will have less fuel load replaced by cargo bay and improved thermal protection, and probably have less than 9 Raptors(3?).
Just a thought to standardize and reduce cost...
Setting aside Elon's actual words aside for a moment about it being single core, the 2-piece concept myself and a few others have been debating about would do what you are doing, but with just two pieces rather than 3. One big monolithic RTLS booster, and one combo upperstage/spacecraft that can get itself to LEO where it will be refueled prior to going to Mars.
And back to design concepts, it is the most fun :)
What if BFR and MCT will be the essentially same thing?
Kind of universal module sized similar to S-IC, 9 Raptors at the bottom, with lending legs, around 1900mt.
Than all thing will be three core (I know, Elon said one-core, but look at F5), MCT in the center will have less fuel load replaced by cargo bay and improved thermal protection, and probably have less than 9 Raptors(3?).
Just a thought to standardize and reduce cost...
Setting aside Elon's actual words aside for a moment about it being single core, the 2-piece concept myself and a few others have been debating about would do what you are doing, but with just two pieces rather than 3. One big monolithic RTLS booster, and one combo upperstage/spacecraft that can get itself to LEO where it will be refueled prior to going to Mars.
Actually this looks to me like the first innovative new idea for a while.
Elon Musk said single core. But the idea behind that was to my understanding, not a 3 core heavy configuration because the central core would go too fast for easy RTLS and would incur heavy payload loos for reuse. This concept avoids that problem.
This concept would be like a first stage in two parts, something completely different. The "central core" would be the MCT. The vac engine problem might be solvable with a retractable engine bell extension. The mechanism shown in that Falcon Heavy animation seems to allow fast efficient reconnection so should not be a major problem for simple operation.
Two side cores with 9 engines each plus a central core with 5? engines would give a total number of engines 23 for lift off. Most of them would be switched off as soon as the T/W ratio allows it to retain fuel for reaching orbit.
A single MCT won't be billions of dollars of investment or the $500k per passenger figure will be impossible. An order of magnitude less like $100-400 million. Also, electronics can easily be made reliable enough. We have lots an lots of experience running spacecraft for years at a time without maintenance. 6-9 months won't be a challenge for a company that will launch thousands of satellites into LEO.
A single MCT won't be billions of dollars of investment or the $500k per passenger figure will be impossible. An order of magnitude less like $100-400 million. Also, electronics can easily be made reliable enough. We have lots an lots of experience running spacecraft for years at a time without maintenance. 6-9 months won't be a challenge for a company that will launch thousands of satellites into LEO.
As these craft are supposed to effectively be the DC-3s of space, acting as both cargo and passenger carriers, Billions would be possible. After all, airliners cost tens of millions of dollars and most airline tickets are less than $400.
No, billions would not be possible. 15 reuses for passenger MCTs, 100 passengers max, $500k per passenger yields a maximum cost per MCT of $750m. But you also have the BFR (first stage) and refueling flights plus operations and refurb cost, etc, plus cargo (although some of that will be funded in other ways perhaps), cost of capital over 30 years, plus the desire to reduce costs to below $500k, and you really, really need to get costs to around $200-400m per MCT. Which also means lightweighting the heck out of it, using as small of volume as you can get away with, etc.A single MCT won't be billions of dollars of investment or the $500k per passenger figure will be impossible. An order of magnitude less like $100-400 million. Also, electronics can easily be made reliable enough. We have lots an lots of experience running spacecraft for years at a time without maintenance. 6-9 months won't be a challenge for a company that will launch thousands of satellites into LEO.
As these craft are supposed to effectively be the DC-3s of space, acting as both cargo and passenger carriers, Billions would be possible. After all, airliners cost tens of millions of dollars and most airline tickets are less than $400.
Don't forget that your income calculation is only for colonists. There will also be a market for cargo on the MCT as well, which may well be able pull in more per KG (scientific payloads etc). I reckon you might be able to pull in $1B on each MCT trip.About once a decade, NASA might spend a billion dollars to put a fancy rover on Mars. Maybe $1-4 billion per year to transport stuff for a crewed research outpost, I don't imagine much more than that. But that's still only a tiny fraction of the $500k per person ticket price. At such high flight rates, a single MCT will not be able to command anywhere near $1B.
How much does it currently cost to put a 100kg payload on Mars?
You may have smaller tanks, but they'll still be large in relation to the overall vehicle.
It's not what I am or am not a fan of, but what must be at a minimum. MCT must be a single stage to Mars orbit vehicle at a minimum. And thus, it will still be a big gas can that must get through EDL, whether it does direct return or not...whether it's it's own 2nd stage on Earth ascent or not.
:-)
Again please link me to this quote.http://www.amazon.com/Elon-Musk-SpaceX-Fantastic-Future/dp/0062301233
Musk has said a lot of stuff that did not survive contact with realityAnd? You've done better? Musk and occasionally Shotwell are basically the ONLY source for anything about MCT. So like it or not, that's all we've got right now.
and we know they are looking at SEP right now.And one of the best ways of using SEP is to haul propellant around, from LEO to EML1/2 (or high Earth orbit).
And Musk doesn't think there's anything worthwhile exporting from Mars, so MCT has to be able to pay for itself.
I've extremely high doubts that once the BFR/MCT package is complete, SpaceX will just cancel its breadwinning Falcon9/Falcon Heavy lines and launch all commercial satellites on a Saturn V class rocket.
I've extremely high doubts that once the BFR/MCT package is complete, SpaceX will just cancel its breadwinning Falcon9/Falcon Heavy lines and launch all commercial satellites on a Saturn V class rocket.
.. and assuming that's the case, what's BFR for? Annual (at best) launches of MCTs?
You would be sending up cargo and propellant flights all year round, waiting in Earth orbit for the departure window, not waiting on the ground. (Or possibly more exotic ballistic trajectories.)I've extremely high doubts that once the BFR/MCT package is complete, SpaceX will just cancel its breadwinning Falcon9/Falcon Heavy lines and launch all commercial satellites on a Saturn V class rocket.
.. and assuming that's the case, what's BFR for? Annual (at best) launches of MCTs?
And? You've done better? Musk and occasionally Shotwell are basically the ONLY source for anything about MCT. So like it or not, that's all we've got right now.
In 2016 SpaceX has bookings for ~$1.3B of rockets (22 at $57m via Google and the US launch schedule thread). They receive about $650M annually from the CCtCap award ($2.6B through 2019) and about $200M from CRS, that's about $2.15B in revenues. A very cursory search suggests that Tesla and Google both have 25% profit margins so let's assume that SpaceX gets the same and makes a profit of ~$550M. Suppose half that profit goes into funding private Mars work, so that 1/8th of all revenue is being spent on internal Mars R&D and mission costs. That's a $275m annual budget for Mars efforts, on the order of half their yearly costs on the CCtCap award.
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Proportion of MCT module (same as side booster modules), as I mentioned can be similar to S-1C, which are ~10m diameter and ~40m long, but can be wider. It is not anywhere as skinny as F9 :)d.And back to design concepts, it is the most fun :)
What if BFR and MCT will be the essentially same thing?
Kind of universal module sized similar to S-IC, 9 Raptors at the bottom, with lending legs, around 1900mt.
Than all thing will be three core (I know, Elon said one-core, but look at F5), MCT in the center will have less fuel load replaced by cargo bay and improved thermal protection, and probably have less than 9 Raptors(3?).
Just a thought to standardize and reduce cost...
Setting aside Elon's actual words aside for a moment about it being single core, the 2-piece concept myself and a few others have been debating about would do what you are doing, but with just two pieces rather than 3. One big monolithic RTLS booster, and one combo upperstage/spacecraft that can get itself to LEO where it will be refueled prior to going to Mars.
Actually this looks to me like the first innovative new idea for a while.
Elon Musk said single core. But the idea behind that was to my understanding, not a 3 core heavy configuration because the central core would go too fast for easy RTLS and would incur heavy payload loos for reuse. This concept avoids that problem.
This concept would be like a first stage in two parts, something completely different. The "central core" would be the MCT. The vac engine problem might be solvable with a retractable engine bell extension. The mechanism shown in that Falcon Heavy animation seems to allow fast efficient reconnection so should not be a major problem for simple operation.
Two side cores with 9 engines each plus a central core with 5? engines would give a total number of engines 23 for lift off. Most of them would be switched off as soon as the T/W ratio allows it to retain fuel for reaching orbit.
It is innovative (good job fast). And it could work for the reason you say. It eliminates the central core that's staging too fast. There's also be a booster core for the intermediate size LV that's been discussed over on the SFR thread.
But, I just don't know that it has advantages over an in-line wider core? Do you see where this would be an advantage? Maybe if they weren't planning to build the cores near the launch facility, this concept would result in thinner cores that are more easily transported. But that's really not a problem unless SpaceX changes from what they've stated.
And it'd result in a tall and skinny MCT. That's probably the major issue with it vs. monolithic, that I see.
But, I just don't know that it has advantages over an in-line wider core? Do you see where this would be an advantage?
In 2016 SpaceX has bookings for ~$1.3B of rockets (22 at $57m via Google and the US launch schedule thread). They receive about $650M annually from the CCtCap award ($2.6B through 2019) and about $200M from CRS, that's about $2.15B in revenues. A very cursory search suggests that Tesla and Google both have 25% profit margins so let's assume that SpaceX gets the same and makes a profit of ~$550M. Suppose half that profit goes into funding private Mars work, so that 1/8th of all revenue is being spent on internal Mars R&D and mission costs. That's a $275m annual budget for Mars efforts, on the order of half their yearly costs on the CCtCap award.They have enough revenues to develop Raptor and perhaps the first few MCT/BFR, but Mars will not be paid for by Dragon or F9/FH, but according to Musk, Mars will be paid for by new growth from the SpaceX constellation. Did you miss that?
A doubling of commercial sales -- 50 launches a year -- adds about $170m to this estimate and puts the effort at $450M yearly. A great success of reuse might drastically affect profit margins, or it might not, since now you need to build fewer rockets for the same flight rate, but you also need to maintain reflown stages, and you might end up discounting your launch price anyways. Imagining a 50% profit margin and doubled sales, SpaceX can spend $900M on Mars work yearly.
What I am getting at is that the scale of SpaceX's commercial rocketry business is borderline in terms of building and operating individual exploration craft. Fleets of $500M MCTs are beyond SpaceX's independent means, even under optimistic expectations.
I am also skeptical of SpaceX trying to switch horses in midstream between kerolox Falcon and methalox BFR architectures with this cost structure. Before they had an order book and DoD certification, they had a lot of freedom to change rocket configurations. Now, they have an ongoing business of maintaining Falcon service and associated Falcon costs. If SpaceX tries to leap directly to methalox and BFR in one go, that adds on Raptor costs and BFR costs and MCT costs all at once. These are going to be Large Rocket Costs, much larger than those incurred for Falcon. I am skeptical that the existing business can bear them. SLS costs are $2.2b/year, and BFR is larger than SLS. I don't see how SpaceX can develop a rocket larger than SLS and a very advanced upper stage / spacecraft capable of Mars EDL/ascent for a small fraction of the price.
I think Musk's comments about a "single monster boost stage" can encompass a range of monstrosity. A single core equivalent of Falcon Heavy would, after all, be at least a smidge monstrous. A craft that delivers "100 metric tons of useful payload to the surface of Mars" can be assembled in Earth orbit and does not need to launch intact from Earth's surface.
Ultimately I think that SpaceX might well achieve a manned Mars landing, but I am doubtful of their ability to independently fund a major colonization architecture. It does make me wonder if SpaceX might try to get into the satellite business to scale up their revenues in order to better follow up on the Mars goal.
In 2016 SpaceX has bookings for ~$1.3B of rockets (22 at $57m via Google and the US launch schedule thread). They receive about $650M annually from the CCtCap award ($2.6B through 2019) and about $200M from CRS, that's about $2.15B in revenues. A very cursory search suggests that Tesla and Google both have 25% profit margins so let's assume that SpaceX gets the same and makes a profit of ~$550M. Suppose half that profit goes into funding private Mars work, so that 1/8th of all revenue is being spent on internal Mars R&D and mission costs. That's a $275m annual budget for Mars efforts, on the order of half their yearly costs on the CCtCap award.
...
Assembly in LEO is not feasible in my opinion, it has proved extremely expensive to do any assembly in space.
Thanks for the link, this seems to be the key quote.
“And then one of the key questions is can you get to the surface of Mars and back to Earth on a single stage. The answer is yes, if you reduce the return payload to approximately one-quarter of the outbound payload, which I thought made sense because you are going to want to transport a lot more to Mars than you’d want to transfer from Mars to Earth. For the spacecraft, the heat shield, the life support system, and the legs will have to be very, very light."
It's largely as I suspected, Musk is describing what COULD be done and the constraints he would face in doing it, but it is by no means a commitment that this is how he will proceed even with the initial design. I think the incredible lightness necessary to make it work will prove too risky of a development challenge, he could end up in Venture Star territory if just one of his lightening strategies fails to work the whole thing could collapse.
By setting low bars like Mars surface to Low Mars orbit and low entry velocity the whole design process becomes a much less risky and cutting edge. SpaceX has traditionally not done high risk designs so I think it is far more likely that in the end he chooses the safer design even if it dose require a second SEP vehicle to function and a good deal of rendezvous in space for refueling.And? You've done better? Musk and occasionally Shotwell are basically the ONLY source for anything about MCT. So like it or not, that's all we've got right now.
Actually their is a LOT of information on rocketry out their on the inter-webs and we can and should do our own research if we expect to speculate with any kind of informed way. If Musk & Shotwell quotes were the only permissible source material then this thread would be nothing more then a religious war between the SpaceX fan-boys and the SpaceX haters. If my analysis disagrees with anyone else's, even Musk's I have every right and indeed the responsibility to point that out and I will not heckled by you or anyone else simply because I don't own a rocket company.
In 2016 SpaceX has bookings for ~$1.3B of rockets (22 at $57m via Google and the US launch schedule thread). They receive about $650M annually from the CCtCap award ($2.6B through 2019) and about $200M from CRS, that's about $2.15B in revenues. A very cursory search suggests that Tesla and Google both have 25% profit margins so let's assume that SpaceX gets the same and makes a profit of ~$550M. Suppose half that profit goes into funding private Mars work, so that 1/8th of all revenue is being spent on internal Mars R&D and mission costs. That's a $275m annual budget for Mars efforts, on the order of half their yearly costs on the CCtCap award.Note that there are 3 CRS flights in 2016 and that is $400M
What I am getting at is that the scale of SpaceX's commercial rocketry business is borderline in terms of building and operating individual exploration craft. Fleets of $500M MCTs are beyond SpaceX's independent means, even under optimistic expectations.
I am also skeptical of SpaceX trying to switch horses in midstream between kerolox Falcon and methalox BFR architectures with this cost structure. Before they had an order book and DoD certification, they had a lot of freedom to change rocket configurations. Now, they have an ongoing business of maintaining Falcon service and associated Falcon costs. If SpaceX tries to leap directly to methalox and BFR in one go, that adds on Raptor costs and BFR costs and MCT costs all at once. These are going to be Large Rocket Costs, much larger than those incurred for Falcon. I am skeptical that the existing business can bear them. SLS costs are $2.2b/year, and BFR is larger than SLS. I don't see how SpaceX can develop a rocket larger than SLS and a very advanced upper stage / spacecraft capable of Mars EDL/ascent for a small fraction of the price.
Ultimately I think that SpaceX might well achieve a manned Mars landing, but I am doubtful of their ability to independently fund a major colonization architecture. It does make me wonder if SpaceX might try to get into the satellite business to scale up their revenues in order to better follow up on the Mars goal.
Actually, you probably want to dump the water shielding before Mars entry. Mars has lots of water, and it will be absolutely essential to tap that water for any of this to work.So I guess it might be a good idea to jettison the water before the final injection burn to Mars orbit, that might reduce the fuel required for that manoeuver. A linked question is would we want to jettison the waste, or is the compost value higher than the value of fuel saving? Is the injection burn a large part of the overall fuel use for Mars transfer? My understanding is the faster we go, the more fuel is required to stop at Mars, but for a 6 month mission, is the Mars burn a large portion of the deltaV requirement?
Additionally, clever arrangement of propellant tanks (and surface tension devices) could use your propellant as shielding. Methane is actually significantly more efficient than water for radiation shielding (Water has an average atomic mass of 6, while methane has an average of ~3). Only hydrogen is superior.
That would probably shift the expected fuel:oxidizer ratio to be more fuel rich than it otherwise would be (if you could adjust that ratio on the fly, you may depart EML1/2 with a more stoich ratio but do final burn above Mars with significantly more fuel rich).
So I guess it might be a good idea to jettison the water before the final injection burn to Mars orbit, that might reduce the fuel required for that manoeuver. A linked question is would we want to jettison the waste, or is the compost value higher than the value of fuel saving? Is the injection burn a large part of the overall fuel use for Mars transfer? My understanding is the faster we go, the more fuel is required to stop at Mars, but for a 6 month mission, is the Mars burn a large portion of the deltaV requirement?
From Statements made by SpaceX representatives:Thanks, this is a great summary. I guess this means not much water based radiation shielding? And if as Guckyfan proposes there is no final injection burn, not much fuel at the end for radiation protection either?
- 100mt payload delivery to Mars
- 1/4 payload SSTO return to Earth from mars surface
- prop density 1m^3 for 1mt (LOX and CH4)
- 15m diameter vehicle (this was hited at not actually specified by SpaceX
- Raptor engines 380-385 vacuum ISP 500klbf
A vehicle like this results:
- Vehicle structure+engines+ shield =40mt
- Max propellant load 900mt
- propulsion section (engines and tanks) cylindircal or nearly cylindrical section at base 15m diameter and 6m tall
- bi-conal payload section (first section 15m to 10m diameter 10m tall) (second section 10m to 0m 10m tall) ~1800m^3 volume
-MCT can be its own 2nd stage on the BFR (BFR is basically just the 1st stage) would have ~7.5km/s delta v capability with a 100mt payload+40mt vehicle dry weight +900mt propellant load
-An MCT tanker variant would be a Cargo MCT without any cargo which could deliver ~150mt of propellant to LEO would have 6km/s delta v capability
In order to get to Mars 6-9 tankers docking in LEO-MEO are required
Edit Added: BTW An MCT cargo used as the 2nd stage going just to LEO would be capable of delivering 180mt of payload. Note the 1st stage needs to be capable of ~3km/s delta v with a fully loaded MCT + 180mt of payload on top ~1120mt MCT+payload GLOW
From Statements made by SpaceX representatives:Thanks, this is a great summary. I guess this means not much water based radiation shielding? And if as Guckyfan proposes there is no final injection burn, not much fuel at the end for radiation protection either?
- 100mt payload delivery to Mars
- 1/4 payload SSTO return to Earth from mars surface
- prop density 1m^3 for 1mt (LOX and CH4)
- 15m diameter vehicle (this was hited at not actually specified by SpaceX
- Raptor engines 380-385 vacuum ISP 500klbf
A vehicle like this results:
- Vehicle structure+engines+ shield =40mt
- Max propellant load 900mt
- propulsion section (engines and tanks) cylindircal or nearly cylindrical section at base 15m diameter and 6m tall
- bi-conal payload section (first section 15m to 10m diameter 10m tall) (second section 10m to 0m 10m tall) ~1800m^3 volume
-MCT can be its own 2nd stage on the BFR (BFR is basically just the 1st stage) would have ~7.5km/s delta v capability with a 100mt payload+40mt vehicle dry weight +900mt propellant load
-An MCT tanker variant would be a Cargo MCT without any cargo which could deliver ~150mt of propellant to LEO would have 6km/s delta v capability
In order to get to Mars 6-9 tankers docking in LEO-MEO are required
Edit Added: BTW An MCT cargo used as the 2nd stage going just to LEO would be capable of delivering 180mt of payload. Note the 1st stage needs to be capable of ~3km/s delta v with a fully loaded MCT + 180mt of payload on top ~1120mt MCT+payload GLOW
From Statements made by SpaceX representatives:15m diameter and other such details have not been mentioned. Please cite your sources and put the source quote in the MCT source thread so we know exactly what was said: http://forum.nasaspaceflight.com/index.php?topic=37839.0
- 100mt payload delivery to Mars
- 1/4 payload SSTO return to Earth from mars surface
- prop density 1m^3 for 1mt (LOX and CH4)
- 15m diameter vehicle (this was hited at not actually specified by SpaceX
- Raptor engines 380-385 vacuum ISP 500klbf
A vehicle like this results:
- Vehicle structure+engines+ shield =40mt
- Max propellant load 900mt
- propulsion section (engines and tanks) cylindircal or nearly cylindrical section at base 15m diameter and 6m tall
- bi-conal payload section (first section 15m to 10m diameter 10m tall) (second section 10m to 0m 10m tall) ~1800m^3 volume
-MCT can be its own 2nd stage on the BFR (BFR is basically just the 1st stage) would have ~7.5km/s delta v capability with a 100mt payload+40mt vehicle dry weight +900mt propellant load
-An MCT tanker variant would be a Cargo MCT without any cargo which could deliver ~150mt of propellant to LEO would have 6km/s delta v capability
In order to get to Mars 6-9 tankers docking in LEO-MEO are required
Edit Added: BTW An MCT cargo used as the 2nd stage going just to LEO would be capable of delivering 180mt of payload. Note the 1st stage needs to be capable of ~3km/s delta v with a fully loaded MCT + 180mt of payload on top ~1120mt MCT+payload GLOW
- Vehicle structure+engines+ shield =40mt
The MCT is not a small vehicle. It could conceivably reach Earth orbit as an SSTO with a little payload about 20mt.
I've extremely high doubts that once the BFR/MCT package is complete, SpaceX will just cancel its breadwinning Falcon9/Falcon Heavy lines and launch all commercial satellites on a Saturn V class rocket.
.. and assuming that's the case, what's BFR for? Annual (at best) launches of MCTs?
I like the quote page, thanks! the 15m had a qualifying text with it :-) so I'm keeping with my own ideas for a 10m core for the moment. Although the 15m core is a tempting design basis.From Statements made by SpaceX representatives:15m diameter and other such details have not been mentioned. Please cite your sources and put the source quote in the MCT source thread so we know exactly what was said: http://forum.nasaspaceflight.com/index.php?topic=37839.0
- 100mt payload delivery to Mars
- 1/4 payload SSTO return to Earth from mars surface
- prop density 1m^3 for 1mt (LOX and CH4)
- 15m diameter vehicle (this was hited at not actually specified by SpaceX
- Raptor engines 380-385 vacuum ISP 500klbf
A vehicle like this results:
- Vehicle structure+engines+ shield =40mt
- Max propellant load 900mt
- propulsion section (engines and tanks) cylindircal or nearly cylindrical section at base 15m diameter and 6m tall
- bi-conal payload section (first section 15m to 10m diameter 10m tall) (second section 10m to 0m 10m tall) ~1800m^3 volume
-MCT can be its own 2nd stage on the BFR (BFR is basically just the 1st stage) would have ~7.5km/s delta v capability with a 100mt payload+40mt vehicle dry weight +900mt propellant load
-An MCT tanker variant would be a Cargo MCT without any cargo which could deliver ~150mt of propellant to LEO would have 6km/s delta v capability
In order to get to Mars 6-9 tankers docking in LEO-MEO are required
Edit Added: BTW An MCT cargo used as the 2nd stage going just to LEO would be capable of delivering 180mt of payload. Note the 1st stage needs to be capable of ~3km/s delta v with a fully loaded MCT + 180mt of payload on top ~1120mt MCT+payload GLOW
Except the Vac-optimized Raptors may not have enough thrust to get it off the ground nor the Isp to get to orbit. Just because it might conceivably get 9km/s in free space doesn't mean it is a SSTO.
Except the Vac-optimized Raptors may not have enough thrust to get it off the ground nor the Isp to get to orbit. Just because it might conceivably get 9km/s in free space doesn't mean it is a SSTO.
The MCT is not a small vehicle. It could conceivably reach Earth orbit as an SSTO with a little payload about 20mt.
20 tonnes to LEO is not a little payload.
If the MCT could really achieve a dry mass of 40 tonnes it would make a very useful reusable SSTO.
Except the Vac-optimized Raptors may not have enough thrust to get it off the ground nor the Isp to get to orbit. Just because it might conceivably get 9km/s in free space doesn't mean it is a SSTO.
15m diameter vehicle (this was hited at not actually specified by SpaceX
Now this is an interesting bit, if accurate. For those looking at commonality with MCT to replace Falcon, here you have it. Not a new Raptor powered SFR. The MCT spacecraft acting as a SSTO LV.
Of course, most unmanned paylaods are not just going to LEO, so a kick stage or something would need to be used for BLEO trajectories. And that kick stage would need to be significantly cheaper than F9US to make any economic advantage over F9R or FHR.
But interesting, nontheless.
From Statements made by SpaceX representatives:Thanks, this is a great summary. I guess this means not much water based radiation shielding? And if as Guckyfan proposes there is no final injection burn, not much fuel at the end for radiation protection either?
- 100mt payload delivery to Mars
- 1/4 payload SSTO return to Earth from mars surface
- prop density 1m^3 for 1mt (LOX and CH4)
- 15m diameter vehicle (this was hited at not actually specified by SpaceX
- Raptor engines 380-385 vacuum ISP 500klbf
A vehicle like this results:
- Vehicle structure+engines+ shield =40mt
- Max propellant load 900mt
- propulsion section (engines and tanks) cylindircal or nearly cylindrical section at base 15m diameter and 6m tall
- bi-conal payload section (first section 15m to 10m diameter 10m tall) (second section 10m to 0m 10m tall) ~1800m^3 volume
-MCT can be its own 2nd stage on the BFR (BFR is basically just the 1st stage) would have ~7.5km/s delta v capability with a 100mt payload+40mt vehicle dry weight +900mt propellant load
-An MCT tanker variant would be a Cargo MCT without any cargo which could deliver ~150mt of propellant to LEO would have 6km/s delta v capability
In order to get to Mars 6-9 tankers docking in LEO-MEO are required
Edit Added: BTW An MCT cargo used as the 2nd stage going just to LEO would be capable of delivering 180mt of payload. Note the 1st stage needs to be capable of ~3km/s delta v with a fully loaded MCT + 180mt of payload on top ~1120mt MCT+payload GLOW
One item I forgot to mention was the number of Raptors on MCT would be 5 to give a 3g liftoff at Mars with immediate 1 engine out continue mission capability. Later in Mars launch even 2 engines out would still enable continue mission. A very low risk value for mission success results from this.
Edit added: GLOW at Mars liftoff would be 965mt or 2.135Mlb more than 2x the GLOW of the F9v1.0. (65mt dry weight [40mt vehicle 25mt payload]). The MCT is not a small vehicle. It could conceivably reach Earth orbit as an SSTO witha little payload about 20mt.
I have just noticed this year's number for the raptor engine, at 230 tonnes rather than 820 tonnes (Raptor engine Wikipedia article). As this is a direct quote from Musk, it seems cannon. That means something like 28+ engines for the fully fuelled ship on Earth. That probably doesn't fit on a 10m core. So the single core design would have to be larger, leading to a very short rocket, 35 to 40 m high for a 15m core, for example. Is that the present consensus, more or less?
From Statements made by SpaceX representatives:That gives you and MCT with a ΔV of 7.475 km/s (380 isp) I am under the impression that ΔV budget of 4.5km/s worst case for LEO-Mars TMI and that variably we have been discussing needs in the order of 1kms for EDL presuming Aerocapture/braking. That leave us with a need for about 5.5km/s, since making LEO as a 2nd stage for F9 needs about 5.9, I am happy with making 6.2km/s the round number leaving a healthy margin of fuel for Earth EDL. At 6.2km/s and 380isp I see the amount of propellant needed for the dry weight and payload you mentioned being 600t at launch, and 472t leaving LEO.
- 100mt payload delivery to Mars
- 1/4 payload SSTO return to Earth from mars surface
- prop density 1m^3 for 1mt (LOX and CH4)
- 15m diameter vehicle (this was hited at not actually specified by SpaceX
- Raptor engines 380-385 vacuum ISP 500klbf
A vehicle like this results:
- Vehicle structure+engines+ shield =40mt
- Max propellant load 900mt
- propulsion section (engines and tanks) cylindircal or nearly cylindrical section at base 15m diameter and 6m tall
- bi-conal payload section (first section 15m to 10m diameter 10m tall) (second section 10m to 0m 10m tall) ~1800m^3 volume
-MCT can be its own 2nd stage on the BFR (BFR is basically just the 1st stage) would have ~7.5km/s delta v capability with a 100mt payload+40mt vehicle dry weight +900mt propellant load
-An MCT tanker variant would be a Cargo MCT without any cargo which could deliver ~150mt of propellant to LEO would have 6km/s delta v capability
In order to get to Mars 6-9 tankers docking in LEO-MEO are required
Edit Added: BTW An MCT cargo used as the 2nd stage going just to LEO would be capable of delivering 180mt of payload. Note the 1st stage needs to be capable of ~3km/s delta v with a fully loaded MCT + 180mt of payload on top ~1120mt MCT+payload GLOW
From Statements made by SpaceX representatives:That gives you and MCT with a ΔV of 7.475 km/s (380 isp) I am under the impression that ΔV budget of 4.5km/s worst case for LEO-Mars TMI and that variably we have been discussing needs in the order of 1kms for EDL presuming Aerocapture/braking. That leave us with a need for about 5.5km/s, since making LEO as a 2nd stage for F9 needs about 5.9, I am happy with making 6.2km/s the round number leaving a healthy margin of fuel for Earth EDL. At 6.2km/s and 380isp I see the amount of propellant needed for the dry weight and payload you mentioned being 600t at launch, and 472t leaving LEO.
- 100mt payload delivery to Mars
- 1/4 payload SSTO return to Earth from mars surface
- prop density 1m^3 for 1mt (LOX and CH4)
- 15m diameter vehicle (this was hited at not actually specified by SpaceX
- Raptor engines 380-385 vacuum ISP 500klbf
A vehicle like this results:
- Vehicle structure+engines+ shield =40mt
- Max propellant load 900mt
- propulsion section (engines and tanks) cylindircal or nearly cylindrical section at base 15m diameter and 6m tall
- bi-conal payload section (first section 15m to 10m diameter 10m tall) (second section 10m to 0m 10m tall) ~1800m^3 volume
-MCT can be its own 2nd stage on the BFR (BFR is basically just the 1st stage) would have ~7.5km/s delta v capability with a 100mt payload+40mt vehicle dry weight +900mt propellant load
-An MCT tanker variant would be a Cargo MCT without any cargo which could deliver ~150mt of propellant to LEO would have 6km/s delta v capability
In order to get to Mars 6-9 tankers docking in LEO-MEO are required
Edit Added: BTW An MCT cargo used as the 2nd stage going just to LEO would be capable of delivering 180mt of payload. Note the 1st stage needs to be capable of ~3km/s delta v with a fully loaded MCT + 180mt of payload on top ~1120mt MCT+payload GLOW
However, personally, even with the smaller propellant load, I would presume MCT's dry weight (which includes potentially some active cooling for propellant and/or sun shades/reflectors, solar power, a higher proportion of RCS than your average US, and several other systems not seen on upper stages) is 60t.
I see it launching with less than full cargo though and having a full tank for the TMI burn.
Can you tell me where your 7.475km/s came from (or where I might have made an assumption about your vehicle concept that reduces its ΔV)
Except the Vac-optimized Raptors may not have enough thrust to get it off the ground nor the Isp to get to orbit. Just because it might conceivably get 9km/s in free space doesn't mean it is a SSTO.
Btw how would MCT land on Earth if it had vac-Raptors?
Mars surface to earth is much less than 9km/s. More like 7-7.5km/s.
The TPS is attached to the tanks and intertank structures through the TPSS. Depending on the tank structural concept and stiffening arrangement, the TPSS may be attached to external stiffeners, such as ring frames and longerons, or to the outer skin of a sandwich tank structure.
As a note, it occured to me where Rocketplane Kistler must have gotten then idea for the shape of the K-1.
;-)
Mars surface to earth is much less than 9km/s. More like 7-7.5km/s.
But still the largest delta v requirement. At some point for growing vehicle dry weight other delta v requirements become larger. A 90mt vehicle weight may even be beyond that point. But leaving Mars surface it would be prudent to have significantly more capability than just the minimum. The item here is that use as 2nd stage to LEO, use as EDS and the use as direct return all have close to the same propellant requirements. The vehicle dry weight minus the payload weight drives which one is the driver for the amount the MCT must hold.
My values gives the information that it is doable and the reasoning behind some of the speculation on what the MCT would do. For Earth departure you just keep filling the tanks until you get enough prop. Note here is that you need also prop for the landing phase on Mars of ~1km/s or more based on the vehicle shape and weight.
Edit: Difference between 2nd stage or EDS and Mars liftoff is payload size 100mt vs 25mt.
MCT will only be reused 15-30 times (once each for Mars entry and Earth entry), so for the crew version, PICA-X may be fine, perhaps with TPS replacement every decade or so. Metallic TPS probably doesn't make sense for MCT, except for perhaps a tanker or cargo version that would just travel to LEO but could be reused hundreds or even thousands of times and which isn't as mass-sensitive.
PICA-X should be just fine for MCT, though, if it works as good as SpaceX says it does. metallic would be suboptimal except for tanker/cargo duty.
From Statements made by SpaceX representatives:15m diameter and other such details have not been mentioned. Please cite your sources and put the source quote in the MCT source thread so we know exactly what was said: http://forum.nasaspaceflight.com/index.php?topic=37839.0
- 100mt payload delivery to Mars
- 1/4 payload SSTO return to Earth from mars surface
- prop density 1m^3 for 1mt (LOX and CH4)
- 15m diameter vehicle (this was hited at not actually specified by SpaceX
- Raptor engines 380-385 vacuum ISP 500klbf
A vehicle like this results:
- Vehicle structure+engines+ shield =40mt
- Max propellant load 900mt
- propulsion section (engines and tanks) cylindircal or nearly cylindrical section at base 15m diameter and 6m tall
- bi-conal payload section (first section 15m to 10m diameter 10m tall) (second section 10m to 0m 10m tall) ~1800m^3 volume
-MCT can be its own 2nd stage on the BFR (BFR is basically just the 1st stage) would have ~7.5km/s delta v capability with a 100mt payload+40mt vehicle dry weight +900mt propellant load
-An MCT tanker variant would be a Cargo MCT without any cargo which could deliver ~150mt of propellant to LEO would have 6km/s delta v capability
In order to get to Mars 6-9 tankers docking in LEO-MEO are required
Edit Added: BTW An MCT cargo used as the 2nd stage going just to LEO would be capable of delivering 180mt of payload. Note the 1st stage needs to be capable of ~3km/s delta v with a fully loaded MCT + 180mt of payload on top ~1120mt MCT+payload GLOW
Basically, yes, this is my thought process.MCT will only be reused 15-30 times (once each for Mars entry and Earth entry), so for the crew version, PICA-X may be fine, perhaps with TPS replacement every decade or so. Metallic TPS probably doesn't make sense for MCT, except for perhaps a tanker or cargo version that would just travel to LEO but could be reused hundreds or even thousands of times and which isn't as mass-sensitive.
PICA-X should be just fine for MCT, though, if it works as good as SpaceX says it does. metallic would be suboptimal except for tanker/cargo duty.
So how would the PICA-X be arranged? Just on the nose like F9USR and K-1? Or on the nose and along one half of the cylinder like most biconic concepts?
And how long would/could the PICA-X be made to last? Assuming two EDL's per round trip, and no more than 1 trip every 2 years, would mean 5 trips per decade max per MCT, and 10 EDL's? I think Dragon's PICA-X is supposedly good for 10 EDL's, but that's from LEO reentry speeds.
The idea is to not have to replace the whole TPS system too often. If it only needed replaced once a decade, that wouldn't seem to be too cumbersome.
A tanker MCT could do several LEO EDL's per year though. Then again, from LEO, metallic TPS should be adequate without the overheating issues of interplanetary return?
But do they want two different TPS configurations? I would think they might rather go with a common platform. Would replacing a PICA-X TPS on a cargo/tanker MCT be too cumbersome to do every 10 missions? (which could be within one year theoretically) Can PICA-X tiles be made to last more than 10 LEO reentries? Or would that be overly heavy?
Maybe a cargo/tanker MCT could have the PICA-X just on the nose like F9US-R and K-1, as it would be coming back only from LEO speeds like them, and would be uncrewed, so fairly high g-loading for that reentry profile wouldn't be a problem for humans. That's be a pretty ballistic reentry without much lift vs. a belly entry. But SpaceX and Rocketplace Kistler seemed to think it would work without crushing their beer cans.
I really don't understand why SpaceX would not limit the BFR to 12 million lbs thrust to take advantage of the Kennedy, then so what if they can only get 75 tons to Mars on MCT instead of 100. Once this BFR/MCT gets going, it will put SLS out of business since it will cost less and the entire Kennedy facility with 4 high bay doors and room for 6 BFR/MCT rockets at a time. That would save SpaceX a ton of money, it seems to me, building the infrastructure for such a large rocket. What would Kennedy be limited to 14m in diameter and 12 million lbs thrust?They may have some clever ideas about how're get beyond usual limits, maybe how to reduce acoustic loads, as we already heard from what's his name's LC39A pad tour. Additionally, they could operate at lower thrust initially.
A small forward facing heat-shield won't present enough surface area to slow down on Mars, you would probably impact the surface at high speed. A smaller ballistic coefficient works on Earth because we have a much deeper atmosphere to slow down in.
- propulsion section (engines and tanks) cylindircal or nearly cylindrical section at base 15m diameter and 6m tall
- bi-conal payload section (first section 15m to 10m diameter 10m tall) (second section 10m to 0m 10m tall) ~1800m^3 volume
Peak heat is also peak deceleration, small forward face delays deceleration in time and in depth into the atmosphere so the time remaining to do any other deceleration is reduced even if your able to increese your cross-sectional area and create more drag. It's the same principle as opening a parachute on Mars, if your too late you impact the surface so I favor presenting maximum surface area as soon as possible.
And even under my lower velocity entry assumption something like ceramic tiles may be needed on the nose and control flaps, stagnation point temperatures can still be quite high.
I don't favor any kind of ablative because I believe we can eventually do more then 2 EDL per synod by sending cargo containers outside of a vehicle and loading them into the lander in Low Mars orbit. Then landing unloading refueling and returning to orbit with enough propellent to land again. At a generous 1 week turn around one landing can bring down ~100 cargo loads per synod, so an indefinite lifespan TPS would be needed for that. While only applicable to cargo this would be a huge improvement in efficiency over the baseline of 1 cargo in 1 lander in 1 synod.
Unfortunately, NASA's funding is not purely for exploration, science, tech, etc, it is also a regional development program for Dixie and a couple other places. So you can have the coolest spaceship ever imagined for super cheap and amazing capability, and it still won't get but a small fraction of funding from NASA. Additionally, NASA is so much more than human spaceflight and launch services. $4 billion annually is the max possible I can imagine anything like BFR/MCT ever getting from NASA (in the next half century), with maybe $1-2 billion being far more likely.This is a good point. But it's also of note that ~$2B/year flowing into the coffers of SpaceX would really help with expeneses. If they made $50M profit per Falcon launch, $2B would be the equivalent of about 40 Falcon launches per year worth of profit.
Then there's the 1st stage with expen$ive 10-15 meter tooling
- propulsion section (engines and tanks) cylindircal or nearly cylindrical section at base 15m diameter and 6m tall
- bi-conal payload section (first section 15m to 10m diameter 10m tall) (second section 10m to 0m 10m tall) ~1800m^3 volume
Your volume estimate is badly off, I calculate your segments are from bottom to top 1060 m ^3 for the cylinder, 1244 m ^3 for the first frustum, 261 m^3 for the top cone (which would in reality need to be blunted to some degree).
Total 2565 m^3, a 60% increase over the vehicle size I'm proposing.
Keep in mind that a F9 first stage has a volume of 480 m^3 and has dry masses of 25 mT and holds just shy of 400 mT of propellents, this should really show how absurd a vehicle with 5 times the volume holding more then twice the propellent and with all the heat shields necessary to do high speed EDL could possibly have a mass only 15 mT more then a F9 first stage.
It's a pressure vessel (and internal equipment needs internal supports) so weight scales with volume, not just area.
Perhaps a mass estimate based on adding up the parts of the MCT would be more reasonable.I like this method for estimation better than just scaling up Dragon (or just random guessing ;) .)
You'll need,
# of Raptors and mass of Raptor (lets assume 100:1 thrust:weight which would make them 2.3 mT each.
Thrust structure mass, probably proportional to thrust, few good examples to base a comparison on
Tank volume and tankage fraction, F9 tanks are reasonable basis for comparison
Surface Area and mass per unit area of Thermal protection
Structural mass, probably proportional to internal volume and peak g-forces.
Landing legs, I've read that these are generally 10% of touch down mass.
Auxiliary systems, solar panels, radiators, batteries, avionics etc etc, again hard to estimate.
Why would MCT weigh that much dry, particularly in a cargo config (because you mentioned 100mt payload, and Musk keeps talking about cargo flights as separate from passenger flights) and without yet counting the heat shield?
I would guess more like 30-35 tons.
L2 MCT Rending Effort (ongoing, large collection):
http://forum.nasaspaceflight.com/index.php?topic=35307.0
Musk has said 80,000 people per year (and ten times as many cargo shipments), which is 1000 Passenger MCTs at once, plus 10,000 cargo MCTs (or actually, there ways around this, but it remains to be seen if they're worth it). So yeah, at any one time, there would need to be thousands of MCTs.
No, he has suggested a mature colony developed over decades would reach a population of 80k, not that eighty thousand settlers would be sent every year....
Incorrect.
https://twitter.com/elonmusk/status/273483420468932608
Did this figure initially take you aback??
I like this method for estimation better than just scaling up Dragon (or just random guessing ;) .)
Was that 10% of touch down mass or touchdown weight? On Mars, it will have 100 tons of cargo but at a fraction of the gravity. Earth landing should be less 0-20 tons of cargo.
The structural mass depends a lot on design details. For example if the crew/cargo section is below the propellant tanks, the walls have to be beefed up to support the full propellant load (~1000 tons) through max Q. Also, having the TPS somewhere other than the bottom (top or side), requires it to be beefed up further to handle load in multiple directions.
I think that there will only be a pressurized version (as opposed to an un-pressurised cargo version) for reasons of the cargo/crew pressure vessel doing double duty as the support structure, just like the propellant tanks.
Did this figure initially take you aback??
Honestly it still does.
But see it that way: combine 80000 people a year, 10 cargo flights for each colonist flight and 50 Million $ per flight the total cost of 440 Billion $
So .. uh .. the cargo is free?
So .. uh .. the cargo is free?
You are right, I miscalculated by one order of magnitude, sorry. Yes it is much higher than the US defense budget. Corrected my post.
Your original flight cost was calculated correctly (800 passenger flights, 8000 cargo flight, 8800 total at $50M a pop yields the $440B.
The cost of the cargo is presently unknown. My gut feeling is that the answer to question "what do you have to pack in order to live on Mars" is quite lengthy, complex and thus expensive.
I would think once the flight rate of 80, 000 colonists per year is reached, the cargo requirements per colonist fight will be much less than 10:1.See this earlier post
By that time, the mars industrial base should be able to produce anything they need short of integrated circuits.
I would think once the flight rate of 80, 000 colonists per year is reached, the cargo requirements per colonist fight will be much less than 10:1.
By that time, the mars industrial base should be able to produce anything they need short of integrated circuits.
I would think once the flight rate of 80, 000 colonists per year is reached, the cargo requirements per colonist fight will be much less than 10:1.
By that time, the mars industrial base should be able to produce anything they need short of integrated circuits.
You may well be right. A colony needs to be well advanced to be able to absorb that many people per year. I really don't see that mass exodus happen, ever. It is just that it is not completely impossible with available ressources.
Why would MCT weigh that much dry, particularly in a cargo config (because you mentioned 100mt payload, and Musk keeps talking about cargo flights as separate from passenger flights) and without yet counting the heat shield?
I would guess more like 30-35 tons.
Just the propellant tank portion, 5 Raptor engines and possibly landing legs too would weight ~40mt. Now add the reentry shield and the cargo bay structure. Of course the cargo variant will not have as high a dry weight as the crew variant but where is the tradeoff in crew payload size and crew vehicle dry weight increase. If you could get the cargo variant to have a dry weight as low as 60mt then reduce the payload size of the crew variant (crew + supplies) to only 60mt on a crew variant that dry weight 100mt things will work out better in that the overall system becomes smaller. You shrink the size and maybe some savings on the propellant tank dry weight due to smaller tanks.
My only problem with the estimates is that the more detail we go the heavier the MCT gets.
Yes, the depot could be a specially equipped MCT for 0 boil-off since it would have large enough tanks to refuel 1+ MCT's for Earth departure. This would make it easy to orbit the depots since they are just another cargo specialized version of the MCT which are then manufactured in the 10's to 100's.Why would MCT weigh that much dry, particularly in a cargo config (because you mentioned 100mt payload, and Musk keeps talking about cargo flights as separate from passenger flights) and without yet counting the heat shield?
I would guess more like 30-35 tons.
Just the propellant tank portion, 5 Raptor engines and possibly landing legs too would weight ~40mt. Now add the reentry shield and the cargo bay structure. Of course the cargo variant will not have as high a dry weight as the crew variant but where is the tradeoff in crew payload size and crew vehicle dry weight increase. If you could get the cargo variant to have a dry weight as low as 60mt then reduce the payload size of the crew variant (crew + supplies) to only 60mt on a crew variant that dry weight 100mt things will work out better in that the overall system becomes smaller. You shrink the size and maybe some savings on the propellant tank dry weight due to smaller tanks.
My only problem with the estimates is that the more detail we go the heavier the MCT gets.
My visualization for the MCT version of the BFR upper stage is 4 raptors, but the hardware to cant them for Mars landing/take off. I think 60t works for the dry weight of a cargo only version, and I am not committed one way or the other yet as to whether the passenger ECLSS and quarters are just cargo 'modules' that fit on an otherwise standard MCT or a seperately designed and built MCT. What I do expect is that a passenger MCT is less loaded with payload than the cargo only one so that it has more ΔV partly for slighlty shorter transit time, partly for more safety margin.
My visualization for a reusable Earth orbit tanker Upper Stage for the BFR is a slightly smaller volume all fuel vehicle that adds little to the launch weight of the BFR and has about 10% lower dry weight than the MCT cargo. Running some numbers this morning I am only seeing 120mt of propellant left over (after margins) to transfer to a depot. Note the idea is that there doesn't need to be active cooling on the tanker since it goes immediately to its offload rendezvous. I want this vehicle to be specially designed because operationally it flies the most. With 4.1 flights per MCT going to Mars, plus whatever flights get made to satisfy other BLEO business the depot, MCTs and other LEO and Earth based Mars infrastructure gets used for. So a small fleet of these makes sense. Cargo MCTs may well work for just launching cargo bound anywhere in LEO or beyond and I suspect that there will be far less volume of this than keeping the depot topped up so no need to specifically develop a LEO cargo MCT.
Yes, the depot could be a specially equipped MCT for 0 boil-off since it would have large enough tanks to refuel 1+ MCT's for Earth departure. This would make it easy to orbit the depots since they are just another cargo specialized version of the MCT which are then manufactured in the 10's to 100's.Why would MCT weigh that much dry, particularly in a cargo config (because you mentioned 100mt payload, and Musk keeps talking about cargo flights as separate from passenger flights) and without yet counting the heat shield?
I would guess more like 30-35 tons.
Just the propellant tank portion, 5 Raptor engines and possibly landing legs too would weight ~40mt. Now add the reentry shield and the cargo bay structure. Of course the cargo variant will not have as high a dry weight as the crew variant but where is the tradeoff in crew payload size and crew vehicle dry weight increase. If you could get the cargo variant to have a dry weight as low as 60mt then reduce the payload size of the crew variant (crew + supplies) to only 60mt on a crew variant that dry weight 100mt things will work out better in that the overall system becomes smaller. You shrink the size and maybe some savings on the propellant tank dry weight due to smaller tanks.
My only problem with the estimates is that the more detail we go the heavier the MCT gets.
My visualization for the MCT version of the BFR upper stage is 4 raptors, but the hardware to cant them for Mars landing/take off. I think 60t works for the dry weight of a cargo only version, and I am not committed one way or the other yet as to whether the passenger ECLSS and quarters are just cargo 'modules' that fit on an otherwise standard MCT or a seperately designed and built MCT. What I do expect is that a passenger MCT is less loaded with payload than the cargo only one so that it has more ΔV partly for slighlty shorter transit time, partly for more safety margin.
My visualization for a reusable Earth orbit tanker Upper Stage for the BFR is a slightly smaller volume all fuel vehicle that adds little to the launch weight of the BFR and has about 10% lower dry weight than the MCT cargo. Running some numbers this morning I am only seeing 120mt of propellant left over (after margins) to transfer to a depot. Note the idea is that there doesn't need to be active cooling on the tanker since it goes immediately to its offload rendezvous. I want this vehicle to be specially designed because operationally it flies the most. With 4.1 flights per MCT going to Mars, plus whatever flights get made to satisfy other BLEO business the depot, MCTs and other LEO and Earth based Mars infrastructure gets used for. So a small fleet of these makes sense. Cargo MCTs may well work for just launching cargo bound anywhere in LEO or beyond and I suspect that there will be far less volume of this than keeping the depot topped up so no need to specifically develop a LEO cargo MCT.
My visualization for the MCT version of the BFR upper stage is 4 raptors, but the hardware to cant them for Mars landing/take off. I think 60t works for the dry weight of a cargo only version, and I am not committed one way or the other yet as to whether the passenger ECLSS and quarters are just cargo 'modules' that fit on an otherwise standard MCT or a seperately designed and built MCT. What I do expect is that a passenger MCT is less loaded with payload than the cargo only one so that it has more ΔV partly for slighlty shorter transit time, partly for more safety margin.
My visualization for the MCT version of the BFR upper stage is 4 raptors, but the hardware to cant them for Mars landing/take off. I think 60t works for the dry weight of a cargo only version, and I am not committed one way or the other yet as to whether the passenger ECLSS and quarters are just cargo 'modules' that fit on an otherwise standard MCT or a seperately designed and built MCT. What I do expect is that a passenger MCT is less loaded with payload than the cargo only one so that it has more ΔV partly for slighlty shorter transit time, partly for more safety margin.
I don't believe integrated habitat and direct Earth-reutrn are compatible.
My favorite,Quote from: Musk
I mean, if you do a densified liquid methalox rocket with on-orbit refueling, so like you load the spacecraft into orbit and then you send a whole bunch of refueling missions to fill up the tanks and you have the Mars colonial fleet - essentially - that gets built up during the time between Earth-Mars synchronizations, which occur every 26 months, then the fleet all departs at the optimal transfer point.
Elon Musk at MIT
http://shitelonsays.com/transcript/elon-musk-at-mits-aeroastro-centennial-part-2-of-6-2014-10-24
My visualization for the MCT version of the BFR upper stage is 4 raptors, but the hardware to cant them for Mars landing/take off. I think 60t works for the dry weight of a cargo only version, and I am not committed one way or the other yet as to whether the passenger ECLSS and quarters are just cargo 'modules' that fit on an otherwise standard MCT or a seperately designed and built MCT. What I do expect is that a passenger MCT is less loaded with payload than the cargo only one so that it has more ΔV partly for slighlty shorter transit time, partly for more safety margin.
I don't believe integrated habitat and direct Earth-reutrn are compatible.
An integrated hab for 6-10 people could probably mass less than 25 tonnes, so it seems possible for initial missions. Later missions with more would need a larger hab and that could not be integrated.
So I believe you have made an important point, integrated habs have no long term future on the MCT, so will probably not be designed in the first place.
You are correct about mass vs. weight, but I would think the Martian landing will be at lower speed due to the lower gravity (wider tolerance for v=0 altitude=0 point). Good point about Mars GLOW though I think we would both agree that that should not be the peak force on the legs (Earth landing will).
The landing gear mass is almost certainly driven by the force of impact with the surface NOT the static weight of the vehicle, in other words objects still have inertia irregardless of gravity. And even if static weight weight were the concern you would need to size the legs based on the gross take off weight which we all agree will be greater then landing weight.
F9 first stage has 8% of dry mass in the leg system, and this is designed for flat artificial surfaces and is not carry precious human cargo. The LEM had around 3% of touch down mass in legs, but that was a soft-touchdown with a deeply throttling engine, not the SpaceX 'hover-slam'.
I think Gross take off Weight will be ~450 mT total, not these monstrous 1000 ton figures. And their would not be any kind of integral habitat in a 'crew' version. Their will just be a single version with an unpressurized cargo bay into which a habitat module would be placed.
Max Q is aerodynamic pressure peak, in the Martian atmosphere it is an almost irrelevant force compared to the force experienced during launch from Earth, it is not the same as max g-forces which is what would be relevant for not crushing the vehicle.
You are correct about mass vs. weight, but I would think the Martian landing will be at lower speed due to the lower gravity (wider tolerance for v=0 altitude=0 point). Good point about Mars GLOW though I think we would both agree that that should not be the peak force on the legs (Earth landing will).
Good comparison on the lunar lander vs F9 S1 legs. The F9 legs had the additional constraints of having to be deployable, aerodynamic when folded, support a higher COV vehicle, and as you mentioned a higher speed impact (through the nominal should be close to 0). Probably all the same constraints the MCT legs will have.
The reason I think they will all be pressurized is that I think having an aero shell and a separate pressure vessel is a waste of mass. The exception to this would be if you are leaving the habitat behind on mars. With a top or side TPS I think a pressurized volume is all but required. Are you assuming a capsule design with the TPS on the bottom?
Regarding max Q and max Gs. I agree with you in the general sense, but I think we may be talking past each other. Different phases of flight have different masses, and a larger mass at a given G load requires more structure. There are also different "sources" of force acting on the MCT (inner stage, engine thrust structure, TPS, nose). The highest G phase will be the peak pressure on most structural members, but not all.
At max Q (earth assent) the MCT structure has to support the aerodynamic pressure plus the payload mass (times Gs) plus the full propellant mass (times Gs) (yes I'm assuming combined S2, you might not be). My point was that certain structural members will see a higher pressure at max Q than at higher G phases of the flight (MECO, SECO, mars assent, and mars or earth entry)
Edit: spelling
A 5-person crew's luxurious tourist hab is a 25-person crew's adequate expedition hab is a 100-person crew's short-term transfer hab. Design once, and use it for multiple campaigns.
Agree that static weight is not an issue, it is speed of impact with the surface. But I disagree that gravity will be the main determinant of that, on both Mars and Earth the vehicle will be under propulsive decent and gravity is simply a part of that 'dance' the Raptor engine is TOO MUCH thrust to hover on at either Earth or Mars (230 mT hover on Earth, 605 mT on Mars), no one belives MCT would have that much mass at touchdown. I favor a set of vernier engines specifically to make a softer touchdown (and avoid cratering the surface)Agreed on the terminal landing thrusters. On my design back in the first thread I actually used pressure fed metholox thrusters for the entire EDL (Earth and Mars) without using raptor at all. I am re-thinking that, especially now that raptor is smaller. Raptor would have an ISP advantage in the super-sonic retro-propulsion. Terminal landing is where raptor is far from ideal.
I have never heard of an space vessel in which the outer aero shell IS the pressure vessel, I would speculate that it presents for too much of a thermal pathway into the vessel and would literally COOK the passengers, note that reentry capsules get quite warm inside during re-entry and this is with considerable insulation between the TPS and pressure vessel. So I do not believe what your describing is possible.
Leaving habitats on surface is exactly what I proposed. The overall shape I'm going with (originally Lobo's configuration) is that of a biconic with TPS on the top/sides, engines and legs on the bottom and an unpressurized cargo-bay door on the side.
Direct-return vehicles are essentially a dead-end configuration because they have a low maximum flight rate, so why go down that path.
Sometimes I use a first approximation and say the cargo is as much as transport cost. That would average 500000 $ for one t. Averaged between simple tools and Intel CPUs.
Direct-return vehicles are essentially a dead-end configuration because they have a low maximum flight rate, so why go down that path.
But SEP stages and transit habs also have a low flight rate. So replacing a direct-return MCT with SEP, transit hab and a smaller MCT lander does not seem to be a win.
The main advantage of a direct-return MCT is that it can be maintained, refurbished and upgraded on Earth.
Any architecture which involves space-only or Mars-only stages has to explain how they will be maintained in space or on Mars. This is not an easy problem to solve, especially if we assume new and upgraded versions of SEP, transit hab and MCT lander are produced every few years. Earth has so many advantages, presence of jigs and tooling, unlimited supplies of water and other working fluids, a full local supply chain, local presence of the design engineers, clean rooms, large hangers under pressure, etc.
I think that the overheads of maintaining equipment off-earth removes any advantage your architecture might have in terms of lower initial mass in LEO, at least until we reach colony sizes of 100,000.
Impaler and CyclerPilot, this idea should be compatible with the COOPS you are developing:Possible yes. Here is a LINK (http://forum.nasaspaceflight.com/index.php?topic=35424.0) to my old design with one big raptor. Very similar to what you proposed. My hab was detachable but for purposes of a LAS. I never thought of leaving it behind.
Consider the passenger version of MCT, the trans-Hab and the surface-Hab to be all the same unit. By this I mean that the passenger section and the propulsion section of MCT would be separate and functionally independent units, i.e. propulsion avionics entirely in the propulsion unit and ECLSS components entirely within the passenger unit.. The passenger unit would sit on top of the propulsion unit, which sports a wide heat shield on its underbelly for EDL. The bottom rim of the passenger unit would be joined to the top rim of the propulsion unit only by a ring of bolts through both rims.
Passengers would ride in this vehicle from Earth and land on Mars' surface. A mobile robotic arm that was previously deployed onto Mars' surface would remove the bolts holding the units together. Then a pre-deployed crane would raise the passenger unit and place it on the ground in a desirable location. The passengers have now landed on Mars along with a permanent habitat unit complete with ECLSS. The propulsion unit, now rather lightweight, could then be launched back to Earth and reused. Note that the heat shield on the underbelly also returns.
This system could also pre-deploy habitats on Mars prior to the first human landing.
As more colonists arrive and build ISRU-based habitats, these original habs would continue to be employed as backup in case of emergency or simply additional housing to give colonists more living space. Also note that not all habs would be permanently located on Mars; some would be launched back to Earth with persons wishing to return.
A cargo version of MCT could also perform double-duty. Once landed, the cargo unit would likewise be removed and set on the ground. Unloading cargo would proceed from ground level to ground level. After unloading, the cargo hatch door(s) would be closed and permanently welded shut, both the interior pressure vessel and the exterior shell. Now we have a sizable tank for storing propellants or other liquids produced on Mars.
Do you think this is feasible?
Direct-return vehicles are essentially a dead-end configuration because they have a low maximum flight rate, so why go down that path.
But SEP stages and transit habs also have a low flight rate. So replacing a direct-return MCT with SEP, transit hab and a smaller MCT lander does not seem to be a win.
Any architecture which involves space-only or Mars-only stages has to explain how they will be maintained in space or on Mars.
No, not remotely.I value your critique.
First off, all sides of an object doing reentry need Thermal protection systems because hot air swirls around the back of a capsule shaped vehicle, so the habitat your placing on the top would need extensive TPS which then gets left on Mars. Likewise their would need to be yet more on the now exposed top of the propulsion stage to allow it to land on Earth. Lastly their is no way to send anyone or anything back to Earth which is required.
The crane necessary to remove this habitat would be monstrous, and it would need to be mobile both before and AFTER picking up the habitat for it to do anything other then put it on the ground right next to the propulsion section which needs to blast off again, a very bad place to be. The crane would have a higher mass then what it is lifting and would be extremely dangerous.
I'm proposing a habitat that is INSIDE the lander and deployed by WHEELS down a ramp, I can't see anything being simpler then that, and am perplexed why anyone feels this needs improving.
...Your proposal is excellent and should be employed. I would add only this:
...
I'm proposing a habitat that is INSIDE the lander and deployed by WHEELS down a ramp, I can't see anything being simpler then that, and am perplexed why anyone feels this needs improving.
...Your proposal is excellent and should be employed. I would add only this:
...
I'm proposing a habitat that is INSIDE the lander and deployed by WHEELS down a ramp, I can't see anything being simpler then that, and am perplexed why anyone feels this needs improving.
After you have unloaded your habitat, now go back and unload the whole top section of the MCT as another habitat.
Maximum cargo delivered in just one trip of the MCT. Win-win for SpaceX!
###
[Edit: A bonus -- we will be returning to Earth the absolute minimum mass that is physically possible.]
First off, all sides of an object doing reentry need Thermal protection systems because hot air swirls around the back of a capsule shaped vehicle, so the habitat your placing on the top would need extensive TPS which then gets left on Mars.
...I agree. And I bow to your earlier posting (reference?)
...
...
I have suggested before, that the whole cabin or cargo compartment may be removable and reused as habitat space on Mars. But that as an initial method. Not in a later stage when large numbers of colonists are transfered, that means the 100 people per flight are actually transported. At that stage the colony needs to be able to provide habitats and work places for all the arriving colonists.
I agree. And I bow to your earlier posting (reference?)
No, not remotely.I value your critique.
First off, all sides of an object doing reentry need Thermal protection systems because hot air swirls around the back of a capsule shaped vehicle, so the habitat your placing on the top would need extensive TPS which then gets left on Mars. Likewise their would need to be yet more on the now exposed top of the propulsion stage to allow it to land on Earth. Lastly their is no way to send anyone or anything back to Earth which is required.
The crane necessary to remove this habitat would be monstrous, and it would need to be mobile both before and AFTER picking up the habitat for it to do anything other then put it on the ground right next to the propulsion section which needs to blast off again, a very bad place to be. The crane would have a higher mass then what it is lifting and would be extremely dangerous.
I'm proposing a habitat that is INSIDE the lander and deployed by WHEELS down a ramp, I can't see anything being simpler then that, and am perplexed why anyone feels this needs improving.
You mentioned needing TPS on all sides of the MCT; I am only familiar with TPS on one side, where it serves as a leading edge during aerobraking on Mars,i.e. Design Reference Architecture 5A. Is the all-around TPS now a requirement for all Mars landers? Did your design include this?
I have seen a suggested design for MCT that is an enlarged version of a Dragon V.2, which sports a 15 m shield on the bottom and heat-resistant metal or composite for the rest of the "capsule'". Is this approach now obsolete?
The system I suggested could have TPS on all surfaces, but that would probably be expensive. If so, it would be less desirable to leave a whole section on Mars permanently. The large crane I suggested could be replaced by a different, low-mass system. But the first issue is TPS.
Your proposal is excellent and should be employed. I would add only this:
After you have unloaded your habitat, now go back and unload the whole top section of the MCT as another habitat.
Maximum cargo delivered in just one trip of the MCT. Win-win for SpaceX!
###
[Edit: A bonus -- we will be returning to Earth the absolute minimum mass that is physically possible.]
Except that the proposals are in headon conflict with two basic principles Elon Musk has stated over and over again.
One is mass fraction. Best possible mass fraction is required for full reusability. Having an outer shell capable of withstanding atmospheric reentry at interplanetary speeds and the resultant heating plus an inner habitat capable of holding pressure for crew is extremely mass inefficient.
The second is full reusability. I think it might be possible that some low value but heavy equipment not needed for a smaller return crew might be removed if necessary. Especially if they can be reused on Mars. But except early on as a special startup arrangement the complete MCT will go back to reach the cost goals.
I have suggested before, that the whole cabin or cargo compartment may be removable and reused as habitat space on Mars. But that as an initial method. Not in a later stage when large numbers of colonists are transfered, that means the 100 people per flight are actually transported. At that stage the colony needs to be able to provide habitats and work places for all the arriving colonists.
Why would MCT weigh that much dry, particularly in a cargo config (because you mentioned 100mt payload, and Musk keeps talking about cargo flights as separate from passenger flights) and without yet counting the heat shield?
I would guess more like 30-35 tons.
Just the propellant tank portion, 5 Raptor engines and possibly landing legs too would weight ~40mt. Now add the reentry shield and the cargo bay structure. Of course the cargo variant will not have as high a dry weight as the crew variant but where is the tradeoff in crew payload size and crew vehicle dry weight increase. If you could get the cargo variant to have a dry weight as low as 60mt then reduce the payload size of the crew variant (crew + supplies) to only 60mt on a crew variant that dry weight 100mt things will work out better in that the overall system becomes smaller. You shrink the size and maybe some savings on the propellant tank dry weight due to smaller tanks.
My only problem with the estimates is that the more detail we go the heavier the MCT gets.
F9 first stage has 8% of dry mass in the leg system, and this is designed for flat artificial surfaces and is not carry precious human cargo. The LEM had around 3% of touch down mass in legs, but that was a soft-touchdown with a deeply throttling engine, not the SpaceX 'hover-slam'.
And their would not be any kind of integral habitat in a 'crew' version. Their will just be a single version with an unpressurized cargo bay into which a habitat module would be placed.
Yes, the depot could be a specially equipped MCT for 0 boil-off since it would have large enough tanks to refuel 1+ MCT's for Earth departure. This would make it easy to orbit the depots since they are just another cargo specialized version of the MCT which are then manufactured in the 10's to 100's.
So it would take one MCT depot to refuel one MCT going to Mars. The Depot if emptied each time it goes up could come back to be refueled and checked out then. No need to worry about boil off. Launch the Fuel depot, launch the MCT to dock, refuel and head to Mars. Fuel depot returns, refuels, and relaunches with the next MCT.
I too like the idea of a cylinder MCT. It could be stretched for a fuel depot without much expense.
My visualization for the MCT version of the BFR upper stage is 4 raptors, but the hardware to cant them for Mars landing/take off. I think 60t works for the dry weight of a cargo only version, and I am not committed one way or the other yet as to whether the passenger ECLSS and quarters are just cargo 'modules' that fit on an otherwise standard MCT or a seperately designed and built MCT. What I do expect is that a passenger MCT is less loaded with payload than the cargo only one so that it has more ΔV partly for slighlty shorter transit time, partly for more safety margin.
I don't believe integrated habitat and direct Earth-reutrn are compatible.
An integrated hab for 6-10 people could probably mass less than 25 tonnes, so it seems possible for initial missions. Later missions with more would need a larger hab and that could not be integrated.
So I believe you have made an important point, integrated habs have no long term future on the MCT, so will probably not be designed in the first place.
So it would take one MCT depot to refuel one MCT going to Mars. The Depot if emptied each time it goes up could come back to be refueled and checked out then. No need to worry about boil off. Launch the Fuel depot, launch the MCT to dock, refuel and head to Mars. Fuel depot returns, refuels, and relaunches with the next MCT.
I too like the idea of a cylinder MCT. It could be stretched for a fuel depot without much expense.
I see the depot as much larger than that. I would prefer to see passenger carrying MCT's launch in pairs as close to simultaneously as possible. Also because of the intensity of the black body radiation of the earth and its daytime reflection of heat, I see the depot needing far more active cooling than the MCT which will only need to keep its propellant from boiling off near Mars and between Mars and Earth but will not need to keep it cool for long in the 10 radii range of the Earth. I also see the depot with a hab as transit station, and a place where PicaX can be recoated on MCT's along with engine swaps (engines taken off BFR tanker stages)
If you have to have 10 flights of cargo with 1 passenger flight of 100 with say 4 crew, why not have a crew of 4 with 10 passengers on each flight. Have the rest as cargo. That way every MCT would be identical, and so it only carries about 80-90 tons of cargo. However, the cargo could be loaded in modules, that could be unloaded and when emptied can be used for habitat. No need for separate cargo and human flights. The 10 colonists could stay on Mars to work, build habitats, landing pads, solar power stations, and maintain ISRU equipment. 10 MCT's would get you 100 colonists.Lots of reasons this is less efficient.
So it would take one MCT depot to refuel one MCT going to Mars. The Depot if emptied each time it goes up could come back to be refueled and checked out then. No need to worry about boil off. Launch the Fuel depot, launch the MCT to dock, refuel and head to Mars. Fuel depot returns, refuels, and relaunches with the next MCT.Not clear from your post that you need multiple refuel flights. A single launch does not have enough propellant for a MCT TMI. Even the most optimistic mission designs require at least 3 refuel flights. I have seen some estimates as high as 8.
I too like the idea of a cylinder MCT. It could be stretched for a fuel depot without much expense.
If your saying that the Thermal protection can be integrated with the pressure hull then that's a non-starter for the simple reason that it provides a direct heat path into the interior and will COOK people. The pressure hull has to have a stand-off gap between it and the TPS, likewise for propellent tanks.If there is an issue with thermal flux to the structure / pressure vessel, a layer of insulation could be added between them over most of the area. If it gets as hot as you say, your design will probably need this insulation layer too because your load bearing layer will lose strength at elevated temperatures.
Thus a removable habitat inside of an unpressurized bay is hardly any less efficient then an integrated one, the only extra mass is a little sheet metal wall for the payload bay and the system to secure the payload during launch and landing. The benefits are getting BIG habitats deployed on the surface for early missions and reducing the mass at launch from Mars surface to a minimum.
What is inefficient is doing high speed entry from interplanetary speeds, I'm advocating for much much lower speeds which will tax the vehicle design far less, which should more then make up for mass costs of a removable hab. And it will also allow a single type of lander to do both crew and cargo flights, a key factor in making it cheaper to develop, manufacture and operate.
If there is an issue with thermal flux to the structure / pressure vessel, a layer of insulation could be added between them over most of the area. If it gets as hot as you say, your design will probably need this insulation layer too because your load bearing layer will lose strength at elevated temperatures.
In your design isn't the "little sheet metal wall for the payload bay" a critical and heavy load bearing member? It has to support the load of the propellant tanks, propellant, TPS, and aerodynamic drag during launches. If it is a combined S2 on earth launch (full prop tanks at max Q and MECO) I would say the added mass requirement is a non-starter. If your MCT launches as an empty S3 then the load is much less but still significant. Hard to say if Earth or Mars launch would be a higher peak load without doing the math.
For nose first reentry that wall has to support the load of the engines and cargo during max G. You said you were using a low speed entry, so probably not a major constraint. Are you accomplishing this propulsively with an oberth burn at Mars? SEP deceleration? Slower transits? Aerobraking? EM drive?
I think your design would be much better and lighter if you made your crew pressure vessel (and cargo containers) load bearing. They could still be modular and removable.
So it would take one MCT depot to refuel one MCT going to Mars. The Depot if emptied each time it goes up could come back to be refueled and checked out then. No need to worry about boil off. Launch the Fuel depot, launch the MCT to dock, refuel and head to Mars. Fuel depot returns, refuels, and relaunches with the next MCT.Not clear from your post that you need multiple refuel flights. A single launch does not have enough propellant for a MCT TMI. Even the most optimistic mission designs require at least 3 refuel flights. I have seen some estimates as high as 8.
I too like the idea of a cylinder MCT. It could be stretched for a fuel depot without much expense.
Having the depot semi-permanently in orbit with good insulation / shading and active cooling allows the operational freedom to launch propellant all throughout the synod. As opposed to a high flight rate sprint followed by 20 months of inactivity.
Musk and his team have already designed a methane-based rocket for the job, and the idea would be it for it to refuel once outside of the Earth’s orbit at a type of fueling station and then make a high-speed journey to Mars in three months.
No, not remotely.
First off, all sides of an object doing reentry need Thermal protection systems because hot air swirls around the back of a capsule shaped vehicle, so the habitat your placing on the top would need extensive TPS which then gets left on Mars.
Question:
Why do we cite EM-L1/L2 which require active stationkeeping propulsion instead of the more stable EM-L4/5 points for orbital depots?
Question:
Why do we cite EM-L1/L2 which require active stationkeeping propulsion instead of the more stable EM-L4/5 points for orbital depots?
Because the goal is to get MCT to Mars. Less delta-v from L1 and especially L2.
No, not remotely.
First off, all sides of an object doing reentry need Thermal protection systems because hot air swirls around the back of a capsule shaped vehicle, so the habitat your placing on the top would need extensive TPS which then gets left on Mars.
hot air => plasma
swirls around => not at hypersonic/supersonic speeds (hot air/plasma basically limited to sonic velocities)
It is not pedantic, as was pointed out earlier in the thread by the quote from Max Faget TPS is unnecessary because there in fact is no such flow of matter dense enough to provide any convective heating.No, not remotely.
First off, all sides of an object doing reentry need Thermal protection systems because hot air swirls around the back of a capsule shaped vehicle, so the habitat your placing on the top would need extensive TPS which then gets left on Mars.
hot air => plasma
swirls around => not at hypersonic/supersonic speeds (hot air/plasma basically limited to sonic velocities)
Don't be pedantic.
Use a modified first stage as a giant depot. Launch it partially filled (to reach orbit dry, so acting as its own upper stage) on top of another first stage.
It is not pedantic, as was pointed out earlier in the thread by the quote from Max Faget TPS is unnecessary because there in fact is no such flow of matter dense enough to provide any convective heating.No, not remotely.
First off, all sides of an object doing reentry need Thermal protection systems because hot air swirls around the back of a capsule shaped vehicle, so the habitat your placing on the top would need extensive TPS which then gets left on Mars.
hot air => plasma
swirls around => not at hypersonic/supersonic speeds (hot air/plasma basically limited to sonic velocities)
Don't be pedantic.
But the space shuttles traveled at 17,000 miles per hour, while Orion will be coming in at 20,000 miles per hour on this first flight test. The faster a spacecraft travels through Earth’s atmosphere, the more heat it generates. So even though the hottest the space shuttle tiles got was about 2,300 degrees Fahrenheit, the Orion back shell could get up to 3,150 degrees, despite being in a cooler area of the vehicle. - See more at: http://www.parabolicarc.com/2014/09/01/heat-shield-installed-orion-spacecraft/#sthash.eNwEh6dP.dpuf
When did he mention an elliptical orbit? I had considered that as well, but have never seen Musk mention a kind of orbit. Please put it in the MCT source quotes thread.Use a modified first stage as a giant depot. Launch it partially filled (to reach orbit dry, so acting as its own upper stage) on top of another first stage.
The modifications to do that would be significant. Rockets aren't legos, after all. Not saying it isn't possible, but I think it's unlikely.
Elon was at one point considering refueling in a high elliptical orbit, which has delta-v advantages, but launch opportunity is more limited. Who knows if he has switched to LEO or L2. It's anybody's guess.
It is not pedantic, as was pointed out earlier in the thread by the quote from Max Faget TPS is unnecessary because there in fact is no such flow of matter dense enough to provide any convective heating.No, not remotely.
First off, all sides of an object doing reentry need Thermal protection systems because hot air swirls around the back of a capsule shaped vehicle, so the habitat your placing on the top would need extensive TPS which then gets left on Mars.
hot air => plasma
swirls around => not at hypersonic/supersonic speeds (hot air/plasma basically limited to sonic velocities)
Don't be pedantic.
It is not pedantic, as was pointed out earlier in the thread by the quote from Max Faget TPS is unnecessary because there in fact is no such flow of matter dense enough to provide any convective heating.No, not remotely.
First off, all sides of an object doing reentry need Thermal protection systems because hot air swirls around the back of a capsule shaped vehicle, so the habitat your placing on the top would need extensive TPS which then gets left on Mars.
hot air => plasma
swirls around => not at hypersonic/supersonic speeds (hot air/plasma basically limited to sonic velocities)
Don't be pedantic.
Yes it is pedantic, he is not disputing the content, he is nitpicking my use of terms, hot air vs plasma. He fails to considered that I might have been using simplified terms because I'm responding to someone who doesn't have all the basics on re-entry and I might not be try to intimidate people with technical terms like describing detached shock-layers and radiative heating which DOSE heat the back side of the vehicle.
Some decade old sour-grapes quotes from Max Faget dose not constitute a counter argument to the fact that EVERY entry vehicle has had a back-shell with thermal protection.
http://www.parabolicarc.com/2014/09/01/heat-shield-installed-orion-spacecraft/QuoteBut the space shuttles traveled at 17,000 miles per hour, while Orion will be coming in at 20,000 miles per hour on this first flight test. The faster a spacecraft travels through Earth’s atmosphere, the more heat it generates. So even though the hottest the space shuttle tiles got was about 2,300 degrees Fahrenheit, the Orion back shell could get up to 3,150 degrees, despite being in a cooler area of the vehicle. - See more at: http://www.parabolicarc.com/2014/09/01/heat-shield-installed-orion-spacecraft/#sthash.eNwEh6dP.dpuf
Some decade old sour-grapes quotes from Max Faget dose not constitute a counter argument to the fact that EVERY entry vehicle has had a back-shell with thermal protection.
Not sure if this has been posted elsewhere or of the veracity of the quote (bold mine):QuoteMusk and his team have already designed a methane-based rocket for the job, and the idea would be it for it to refuel once outside of the Earth’s orbit at a type of fueling station and then make a high-speed journey to Mars in three months.
http://www.bloomberg.com/news/articles/2015-06-24/elon-musk-first-martian-a-serious-conversation-about-the-future-in-space
If the quote is true, which I cannot determine independently, then it shows that there will be a DEPOT SYSTEM... deliver fuel to LEO, say with FH-R in 50mT increments, and then transfer it to outside of the Earth's orbit as staging for the MCT. This is exactly the concept published by ULA for their ACES depot system. (Ref below)
http://www.ulalaunch.com/uploads/docs/Published_Papers/Exploration/DepotBasedTransportationArchitecture2010.pdf
EML-2 is the optimum location delta-v-wise for depots, exploration outpost(s), and Mars departures. EML-1 is almost as good. I have proposed EML-1 as the refit location and EML-2 as fuel topping and departure staging point for the fleet. Some of the fuel could be loaded in LEO, and then the vehicle immediately departs for EML-1/2 for fit-out as a hybrid approach.
When did he mention an elliptical orbit? I had considered that as well, but have never seen Musk mention a kind of orbit. Please put it in the MCT source quotes thread.Use a modified first stage as a giant depot. Launch it partially filled (to reach orbit dry, so acting as its own upper stage) on top of another first stage.
The modifications to do that would be significant. Rockets aren't legos, after all. Not saying it isn't possible, but I think it's unlikely.
Elon was at one point considering refueling in a high elliptical orbit, which has delta-v advantages, but launch opportunity is more limited. Who knows if he has switched to LEO or L2. It's anybody's guess.
Some decade old sour-grapes quotes from Max Faget dose not constitute a counter argument to the fact that EVERY entry vehicle has had a back-shell with thermal protection.
Apollo imagery of people fishing out the capsules with leeward side in near pristine condition proves that Faget was right.
India's SRE-1 (https://en.wikipedia.org/wiki/Space_Capsule_Recovery_Experiment) proves that you are wrong.
Well, I kinda was being pedantic. Sorry Impaler.
Sometimes a simplified explanation becomes technically incorrect, but still can be useful. (Full disclosure: I do this all the time with non-technical listeners to get the big picture point across.)
Impaler, let me push back a bit against the premise of what you are pushing. If I understand you correctly, you advocate lower speed reentry, such as from an orbit (both on Mars and Earth) than direct reentry for the MCT, primarily because then one could use metallic TPS rather than ablative TPS (i.e. PICA-X). And that this is practically necessary to ensure high reusability and high flight rates. Do I have that right?
But is avoiding ablative TPS really that important in the grand scheme? Doesn't SpaceX intend to rapidly and frequently reuse the Dragon 2, which will surely have PICA-X. What do we suppose is the answer here? Is PICA-X something that can be de-ablated (reblated?) back on to the bottom of a capsule without too much hassle. I'm imagining that there is an inch or so of the material, and half an inch ablates off during reentry (a little more some places, a little less others) and then then additional PICA-X is applied and added to what remains before the next flight - sort of like retreading a tire. Or conversely, the entire backshell of PICA-X is designed and installed as a bolt on module, and a new one will be bolted onto the capsule for each flight (i.e. changing the tires). I ask because I don't know.
But in any case, this seems to solve the problem fully, without a great deal of bother. Surely a simpler solution than radically modifying the flight profiles to include orbital insertion on each end, which requires a great deal of additional fuel, don't you think?
First off trying to use your EYE as a after-the-fact calorimeter on a surface that was heat-shielded is a profoundly flawed. You have no idea what surface temperature it reached nor do you have any idea if it is truly pristine, that would take chemical analysis after-the-fact or ideally a temperature probe during the entry itself.
...
The rear of this IRSO vehicle looks to be composed of small solar panels, solar panel are covered with glass which has a high melting point, we would not expect this to blacken or char, the forward TPS is non-ablative ceramic tile which would not deposit black char streaks. The shuttle didn't LOOK chared when it returned from orbit but the top of it had TPS of a thinner, lower temperature type, but TPS none the less. What is behind the panels? you have no idea.
Use a modified first stage as a giant depot. Launch it partially filled (to reach orbit dry, so acting as its own upper stage) on top of another first stage.
Use a modified first stage as a giant depot. Launch it partially filled (to reach orbit dry, so acting as its own upper stage) on top of another first stage.
Given the speculated mass fractions, isn't the actual first stage just about capable of SSTO if it doesn't have a second stage or any payload attached?
It may need a pre-launched second stage to dock and act as a shepherd, raising it to its working location - but that should be do-able?
Very good point! And a much better idea than mine. Especially a huge first stage with no legs. Because it's big, low drag loss. Because no upper stage, low gravity losses. f9 v1.2-like mass fractions and propellant densification. Raptor's high Isp. All those engines allow you to have ability to throttle way down to prevent crushing your stage due to over acceleration of such a light stage. Just put an aerodynamic fairing on top, and yeah, it should have no problem reaching orbit.Use a modified first stage as a giant depot. Launch it partially filled (to reach orbit dry, so acting as its own upper stage) on top of another first stage.
Given the speculated mass fractions, isn't the actual first stage just about capable of SSTO if it doesn't have a second stage or any payload attached?
It may need a pre-launched second stage to dock and act as a shepherd, raising it to its working location - but that should be do-able?
Use a modified first stage as a giant depot. Launch it partially filled (to reach orbit dry, so acting as its own upper stage) on top of another first stage.
Given the speculated mass fractions, isn't the actual first stage just about capable of SSTO if it doesn't have a second stage or any payload attached?
It may need a pre-launched second stage to dock and act as a shepherd, raising it to its working location - but that should be do-able?
Yes but it would mean sending a lot of expensive Raptors on a one off mission. If inflatable LOX/methane tanks are possible it seems like the more cost efficient solution to me.
Use a modified first stage as a giant depot. Launch it partially filled (to reach orbit dry, so acting as its own upper stage) on top of another first stage.
Given the speculated mass fractions, isn't the actual first stage just about capable of SSTO if it doesn't have a second stage or any payload attached?
It may need a pre-launched second stage to dock and act as a shepherd, raising it to its working location - but that should be do-able?
Yes but it would mean sending a lot of expensive Raptors on a one off mission. If inflatable LOX/methane tanks are possible it seems like the more cost efficient solution to me.
That's a good point. Of course if this depot had appropriate fuel levels it could later de-orbit and land for servicing - say every two years following the mars departure window. It is no longer a one off mission, just another part of the reusable infrastructure?
Very good point! And a much better idea than mine. Especially a huge first stage with no legs. Because it's big, low drag loss. Because no upper stage, low gravity losses. f9 v1.2-like mass fractions and propellant densification. Raptor's high Isp. All those engines allow you to have ability to throttle way down to prevent crushing your stage due to over acceleration of such a light stage. Just put an aerodynamic fairing on top, and yeah, it should have no problem reaching orbit.Use a modified first stage as a giant depot. Launch it partially filled (to reach orbit dry, so acting as its own upper stage) on top of another first stage.
Given the speculated mass fractions, isn't the actual first stage just about capable of SSTO if it doesn't have a second stage or any payload attached?
It may need a pre-launched second stage to dock and act as a shepherd, raising it to its working location - but that should be do-able?
But if it's going to be a depot, it might need better insulation than a first stage usually has. That might be a major reason why you wouldn't do this, since you might need to fly it inside a fairing.
Another possibility though would be to retrieve the engines only or keep them as a stock of on orbit spares for replacements (or even ship them to Mars as a pool of spares there).
Use a modified first stage as a giant depot. Launch it partially filled (to reach orbit dry, so acting as its own upper stage) on top of another first stage.
Given the speculated mass fractions, isn't the actual first stage just about capable of SSTO if it doesn't have a second stage or any payload attached?
It may need a pre-launched second stage to dock and act as a shepherd, raising it to its working location - but that should be do-able?
Yes but it would mean sending a lot of expensive Raptors on a one off mission. If inflatable LOX/methane tanks are possible it seems like the more cost efficient solution to me.
That's a good point. Of course if this depot had appropriate fuel levels it could later de-orbit and land for servicing - say every two years following the mars departure window. It is no longer a one off mission, just another part of the reusable infrastructure?
But it would need significant modifications to handle re-entry as it was designed for something like 2.5 km/s re-entry. Another possibility though would be to retrieve the engines only or keep them as a stock of on orbit spares for replacements (or even ship them to Mars as a pool of spares there).
Also because of the intensity of the black body radiation of the earth and its daytime reflection of heat, I see the depot needing far more active cooling than the MCT which will only need to keep its propellant from boiling off near Mars and between Mars and Earth but will not need to keep it cool for long in the 10 radii range of the Earth.Not sure this is a problem. Other depot studies have noted this as an issue for hydrogen but methane is quite a mild cryogen in comparison. With solar power a methane prop depot should be able to be zero boiloff anywhere.
Your original flight cost was calculated correctly (800 passenger flights, 8000 cargo flight, 8800 total at $50M a pop yields the $440B.I think Musk is assuming most things will be manufactured locally before flights reach that level.
The cost of the cargo is presently unknown. My gut feeling is that the answer to question "what do you have to pack in order to live on Mars" is quite lengthy, complex and thus expensive.
Yes boiling points are:You forgot negative signs. Also: Please just use kelvins. It's much easier.
252.8 degrees C for hydrogen
182.9 degrees C for oxygen
162 degrees C for methane.
Hydrogen is much harder to keep from boiling off.
I worked for a natural gas company and we liquefied natural gas in the summer for winter peeks. Boil off was not that big of a problem on the ground, and space is colder. Tanks on the ground were doubled like a thermos bottle, with a vacuum pulled between the inner storage tank and outer shell. There was about 3' of space between them (1m), so keeping cold wasn't hard, and that is in the deep south.
Also because of the intensity of the black body radiation of the earth and its daytime reflection of heat, I see the depot needing far more active cooling than the MCT which will only need to keep its propellant from boiling off near Mars and between Mars and Earth but will not need to keep it cool for long in the 10 radii range of the Earth.Not sure this is a problem. Other depot studies have noted this as an issue for hydrogen but methane is quite a mild cryogen in comparison. With solar power a methane prop depot should be able to be zero boiloff anywhere.
And remember a LEO depot will spend roughly half its time above a sunlit Earth that is radiating significantly more than its black body night time amount and that it will cover a significant fraction of the visible area around the depot.I could say the same of LNG depots on the ground.
True, it is much more difficult in LEO, but an actuated passive system is very flexible, and even a static passive system can be done. Mount a cone-shaped reflective thermal shroud around the tanks, and point it normal to the orbital plane, and so long as your choice of orbital plane isn't very far from the ecliptic, you can be mostly in radiative thermal contact with deep space rather than the Earth or the Sun.Also because of the intensity of the black body radiation of the earth and its daytime reflection of heat, I see the depot needing far more active cooling than the MCT which will only need to keep its propellant from boiling off near Mars and between Mars and Earth but will not need to keep it cool for long in the 10 radii range of the Earth.Not sure this is a problem. Other depot studies have noted this as an issue for hydrogen but methane is quite a mild cryogen in comparison. With solar power a methane prop depot should be able to be zero boiloff anywhere.
There was a paper I read in the last month (and I know it is linked to here on NSF and I will look for it later) that suggested LOX and Methane would be fine more than 10 radii from Earth at Earth's distance from the sun with simply passive cooling, but that near Earth and potentially Mars more cooling would be required. And remember a LEO depot will spend roughly half its time above a sunlit Earth that is radiating significantly more than its black body night time amount and that it will cover a significant fraction of the visible area around the depot.
I worked for a natural gas company and we liquefied natural gas in the summer for winter peeks. Boil off was not that big of a problem on the ground, and space is colder. Tanks on the ground were doubled like a thermos bottle, with a vacuum pulled between the inner storage tank and outer shell. There was about 3' of space between them (1m), so keeping cold wasn't hard, and that is in the deep south.
Also because of the intensity of the black body radiation of the earth and its daytime reflection of heat, I see the depot needing far more active cooling than the MCT which will only need to keep its propellant from boiling off near Mars and between Mars and Earth but will not need to keep it cool for long in the 10 radii range of the Earth.Not sure this is a problem. Other depot studies have noted this as an issue for hydrogen but methane is quite a mild cryogen in comparison. With solar power a methane prop depot should be able to be zero boiloff anywhere.
There was a paper I read in the last month (and I know it is linked to here on NSF and I will look for it later) that suggested LOX and Methane would be fine more than 10 radii from Earth at Earth's distance from the sun with simply passive cooling, but that near Earth and potentially Mars more cooling would be required. And remember a LEO depot will spend roughly half its time above a sunlit Earth that is radiating significantly more than its black body night time amount and that it will cover a significant fraction of the visible area around the depot.
True, it is much more difficult in LEO, but an actuated passive system is very flexible, and even a static passive system can be done. Mount a cone-shaped reflective thermal shroud around the tanks, and point it normal to the orbital plane, and so long as your choice of orbital plane isn't very far from the ecliptic, you can be mostly in radiative thermal contact with deep space rather than the Earth or the Sun.
True, it is much more difficult in LEO, but an actuated passive system is very flexible, and even a static passive system can be done. Mount a cone-shaped reflective thermal shroud around the tanks, and point it normal to the orbital plane, and so long as your choice of orbital plane isn't very far from the ecliptic, you can be mostly in radiative thermal contact with deep space rather than the Earth or the Sun.
So, basically use Webb telescope shielding technology to build a "crater" in orbit, well insulated from Earth. Place propellant tanks in the bottom of that crater. Add a sun shield to make it permanently shadowed and you're set. Illustrated simple setting with articulating boom. Some other geometry might not even need that. Having openings to "vent" the thermal radiation into 2.7K space improves shield efficiency a lot. In regular MLI it just bounces between layers without escape.
Sorry about the negative signs. I'm not a rocket scientist and have never used kelvin. When I looked up the temps, they were not listed in kelvin but C and F. I know kelvin is from absolute zero, but a lot of people here are not rocket scientists but all should know degrees C or F.
Sorry about the negative signs. I'm not a rocket scientist and have never used kelvin. When I looked up the temps, they were not listed in kelvin but C and F. I know kelvin is from absolute zero, but a lot of people here are not rocket scientists but all should know degrees C or F.
Too funny Just asked 6th grade daughter and she knows the Kelvin scale 😀
On another question, could one of these large fuel depots be towed to L1 or L2 for say fueling some MCT's going to and from Mars without them landing every time? Seems like a lot of cargo, in cargo containers that could fit in an MCT could be brought up in Falcon Heavies, say two 50 ton containers. Then towed with SEP tugs to L1 or L2 to be loaded into an MCT to be sent back to Mars. Fuel and LOX in 50 ton units could also be towed to the fuel depot for refilling. If Vulcan comes on line, it too, could send up shipments of cargo/fuel to be loaded and sent to Mars. SpaceX wouldn't have to provide everything. Everyone might eventually get involved in Mars colonization, ESA, Russia, China, India, Japan, NASA, and other American companies. SpaceX just seems to be leading the way.
Sorry about the negative signs. I'm not a rocket scientist and have never used kelvin. When I looked up the temps, they were not listed in kelvin but C and F. I know kelvin is from absolute zero, but a lot of people here are not rocket scientists but all should know degrees C or F.
Too funny Just asked 6th grade daughter and she knows the Kelvin scale 😀
Yes, you don't have to be a rocket scientist to use Kelvin, but most here do have some kind of STEM background and fully understand it. Kelvin is taught in the younger grades and it is the metric used in high school sciences. In high school chemistry, basic calorimetry is measured in Kelvin. Thermal calculations in high school physics are done in Kelvin. Celsius and Kelvin both have a 100 degree difference between the state change temperatures of pure H2O @ STP (Standard Temperature and Pressure), i.e. solid/liquid and liquid/gas. Thus, Celsius and Kelvin scale on a 1:1 ratio. Since absolute zero is 273.15C below the first state change temperature of pure H2O @ STP, given y = temp K and x = temp C, y = x +273.15. Kids do learn how to do this in middle school.
Lets just correct the signs if they are wrong and not mock people for their chosen temperature scales. An american using celsius has already shown generosity towards rest of the world. Not long ago I read scientific paper about gas turbines which had temperatures in rankines.
With the size of a large depot the boiloff problem in LEO may go away. The square cube law helps. Plus constantly arriving sub cooled propellant. What's left of boiloff may justbe accepted for the sake of simplicity of operations.
True but your tank size is going to be constrained by the size of the launch vehicle payload, unless you weld it together in orbit (also a possibility). A 12m diameter tank should be fine though - and hold plenty of propellant.
Plus you have to make ullage burns to get the stuff flowing, or else use something clever like low-temperature bladders. Rotate for (very weak) artificial gravity? That'll probably cause more problems than it solves.
Plus you have to make ullage burns to get the stuff flowing, or else use something clever like low-temperature bladders. Rotate for (very weak) artificial gravity? That'll probably cause more problems than it solves.
You are right, I forgot ullage. If the tank becomes really large that becomes an issue.
With two sets of tanks, it should be possible to arrange for flow of the liquid in the tanks using low velocity mixers. If the mixers were contra rotating, the angular acceleration would tend to cancel out. It would be important to keep velocities low, and to be able to vary the velocity of the mixers for operation with mixed phase fluids. Rotating the whole tank seems complex, and might require rotating joints for fluid transfer, which are weak points.With the size of a large depot the boiloff problem in LEO may go away. The square cube law helps. Plus constantly arriving sub cooled propellant. What's left of boiloff may justbe accepted for the sake of simplicity of operations.
True but your tank size is going to be constrained by the size of the launch vehicle payload, unless you weld it together in orbit (also a possibility). A 12m diameter tank should be fine though - and hold plenty of propellant.
Plus you have to make ullage burns to get the stuff flowing, or else use something clever like low-temperature bladders. Rotate for (very weak) artificial gravity? That'll probably cause more problems than it solves.
With two sets of tanks, it should be possible to arrange for flow of the liquid in the tanks using low velocity mixers. If the mixers were contra rotating, the angular acceleration would tend to cancel out. It would be important to keep velocities low, and to be able to vary the velocity of the mixers for operation with mixed phase fluids. Rotating the whole tank seems complex, and might require rotating joints for fluid transfer, which are weak points.With the size of a large depot the boiloff problem in LEO may go away. The square cube law helps. Plus constantly arriving sub cooled propellant. What's left of boiloff may justbe accepted for the sake of simplicity of operations.
True but your tank size is going to be constrained by the size of the launch vehicle payload, unless you weld it together in orbit (also a possibility). A 12m diameter tank should be fine though - and hold plenty of propellant.
Plus you have to make ullage burns to get the stuff flowing, or else use something clever like low-temperature bladders. Rotate for (very weak) artificial gravity? That'll probably cause more problems than it solves.
As a negative point the mixers would add energy and possible increase boil off because there would be constant friction, but this could be offset by using active mechanical cooling of the fuel. Careful operation should create laminar flow; that would create the lowest friction.
Rather than guessing this or that, how about some math?
The solar constant is about 1300 W/m2. Since the depot is in the Earth's shadow 50% of the time this can be reduced to 650 W/m2 average. A good reflector foil wrap with a low emissivity of e=0.1 will reduce this to 65 W/m2. Depending on insulation effectiveness, this energy will either be absorbed into the fuel, or radiated back out into space. If the insulation was 100% effective, the exterior hull temperature would be determined by Q=Be(Ts^4-ta^4), where B is 5.7e-8, e is the emissivity of 0.1 and ta is the average ambient temperature in earth orbit, that I believe NASA usually sets at about 200°K. Solving this gives a surface temperature of 337°K, or 65°C.
Rather than guessing this or that, how about some math?
The solar constant is about 1300 W/m2. Since the depot is in the Earth's shadow 50% of the time this can be reduced to 650 W/m2 average. A good reflector foil wrap with a low emissivity of e=0.1 will reduce this to 65 W/m2. Depending on insulation effectiveness, this energy will either be absorbed into the fuel, or radiated back out into space. If the insulation was 100% effective, the exterior hull temperature would be determined by Q=Be(Ts^4-ta^4), where B is 5.7e-8, e is the emissivity of 0.1 and ta is the average ambient temperature in earth orbit, that I believe NASA usually sets at about 200°K. Solving this gives a surface temperature of 337°K, or 65°C.
I don't quite understand this. The effective temperature (https://en.wikipedia.org/wiki/Effective_temperature) of a blackbody at 1AU should be 254K (https://en.wikipedia.org/wiki/Black-body_radiation#Temperature_of_Earth). You can set the emissivity fairly high (for rocky Earth, emissivity approaches 1.0), but albedo is the variable you can tweak heavily, and with higher albedo should come lower hull temperatures.
In Low Earth Orbit, you have the added factor of a hemisphere radiating at ~288K (may be subject to some corrections). But nothing, as far as I understand it, should raise the hull temperature up to 337K.
The fuel depot has low emissivity and likely fairly high albedo, these are usually more or less inversely proportional. So it reflects a lot of radiation (high albedo), and therefore absorbs very little , but for the radiation it doesn't reflect, it had a lot of difficulty getting rid of, so it has to heat up quite a bit.
The fuel depot has low emissivity and likely fairly high albedo, these are usually more or less inversely proportional. So it reflects a lot of radiation (high albedo), and therefore absorbs very little , but for the radiation it doesn't reflect, it had a lot of difficulty getting rid of, so it has to heat up quite a bit.
In another discussion it was mentioned that there is something as simple as a paint that combines both properties. It is high albedo - white - in the visible spectrum where the sun emits most of its energy and at the same time low albedo in infrared where a depot needs to get rid of excess energy.
The clearest proof that landing is the plan was the statement that for the first crews MCT would be the habitat on Mars.
The clearest proof that landing is the plan was the statement that for the first crews MCT would be the habitat on Mars.
And their is nothing inconsistent with that statement and having a SEP transit stage. Musk calls it the "Mars Colonial transport SYSTEM" which clearly implies multiple parts such as the BFR first stages and what ever LEO propellent depots are need, neither of which will go to Mars.
...Given the depot system discussed here and given that Elon thinks an all propulsive system could work, consider the following:
...
I recall Elon Musk saying something like it can be done with chemical propulsion, no advanced propulsion systems are necessary. As far as I know he never repeated that statement and it does not preclude SEP. I would not be too surprised if it is added at some point in time to improve efficiency. But I am very sure it will not be part of the initial system at the time when a first base is set up because it adds complexity.
Four fully-fueled Tanker-MCT (TMCT) are clustered around and attached to a Mars-Bound-MCT (MBMCT). The engines of the four TMCTs are lit and push the cluster to HEO using about 1/2 of their fuel. The MBMCT is released and begins its TMI burn while the remaining four return to LEO and then individually RTLS. Alternatively, just to a fuel depot in LEO.
Feasible?
Edit: Starting point is LEO.
...You are right, it's not the most fuel efficient. But it might be a method for reaching Mars faster than the least-energy transfer orbit to reach Mars, perhaps in in 3-4 months rather than 6. Also a method that would represent a non-SEP architecture if Elon is serious about it. I am sure this has already been addressed somewhere and lies on someone's spreadsheet.
...
I don't think it is a very efficient architecture. It means several MCT with all their mass would need to be accelerated a significant part of TMI. Also you mention using half of their fuel. It would not be necessary to reserve half of the fuel for return. Injecting into a highly ellicptic orbit would give the Mars bound MCT much of the needed delta-v and brings the booster MCT back to earth basically free.
Why do you propose to get them back to LEO? More efficient to land them for a new launch with payload.
I think the most efficient way is giving MCT tanks large enough to do TMI burn and Mars EDL by themselves. Use tanker MCT to refuel in LEO either directly fuelling up an MCT or filling depots. They need that tankage and the delta-v to get back to earth from the Mars surface.
Plus you have to make ullage burns to get the stuff flowing, or else use something clever like low-temperature bladders. Rotate for (very weak) artificial gravity? That'll probably cause more problems than it solves.
Good ideas. I would think the fan could work if it covered most of the diameter. It could be an interesting ISS or Dragon lab experiment.Plus you have to make ullage burns to get the stuff flowing, or else use something clever like low-temperature bladders. Rotate for (very weak) artificial gravity? That'll probably cause more problems than it solves.
Maybe a totally crazy idea. But could a fan be used to herd the propellant to the pumps? That would save the need to accelerate a depot with thousands of tons of propellant so it can settle.
There have been several MCT designs that had large enough tanks to do the 4 month transfer direct from LEO. (I think the jump from 4 to 3 months is pretty insane most synods)....You are right, it's not the most fuel efficient. But it might be a method for reaching Mars faster than the least-energy transfer orbit to reach Mars, perhaps in in 3-4 months rather than 6. Also a method that would represent a non-SEP architecture if Elon is serious about it. I am sure this has already been addressed somewhere and lies on someone's spreadsheet.
...
I don't think it is a very efficient architecture. It means several MCT with all their mass would need to be accelerated a significant part of TMI. Also you mention using half of their fuel. It would not be necessary to reserve half of the fuel for return. Injecting into a highly ellicptic orbit would give the Mars bound MCT much of the needed delta-v and brings the booster MCT back to earth basically free.
Why do you propose to get them back to LEO? More efficient to land them for a new launch with payload.
I think the most efficient way is giving MCT tanks large enough to do TMI burn and Mars EDL by themselves. Use tanker MCT to refuel in LEO either directly fuelling up an MCT or filling depots. They need that tankage and the delta-v to get back to earth from the Mars surface.
...Excellent. I have seen a lot of discussion with the SEP option, but not the propellant only option. Do you know where I could find those designs -- on this forum?
...
There have been several MCT designs that had large enough tanks to do the 4 month transfer direct from LEO. (I think the jump from 4 to 3 months is pretty insane most synods).
If you think those tanks are too large, and want to transfer to a higher energy orbit before TMI, you should just use one extra MCT instead of 4. That will save hundreds of tons of propellant because you are moving less dry mass. A reusable SEP tug is another popular option.
One way to speculate about the BFR launcher for the MCT is to look at mass fraction to LEO efficiency. The F9 has a mass fraction of ~2.7% to LEO. Despite full re-usability most of us amazing peoples/girls here expect that SX will somehow improve on that with the fully re-useable BFR. The tables below 1st assume 180mT to LEO with the dry MCT massing 80mT and the 2nd table assumes a dry MCT at 100mT. Various optimistic mass fractions yield different BFR takeoff weights. A T/W ratio of 1.2 is assumed yielding 1st stage thrust and dividing by 500KLBs/Raptor, the # of Raptors needed.
MCT + Payload = 180mT to LEO
MASS FRACTION BFR BFR TAKEOFF
TO LEO WEIGHT mT WEIGHT M LBS THRUST M LBS # RAPTORS @ 500KLB
5.0% 3500 7.7 9.2 19
4.5% 3889 8.6 10.3 21
4.0% 4375 9.6 11.6 23
3.5% 5000 11.0 13.2 26
3.0% 5833 12.8 15 31
MCT + Payload = 200mT to LEO
MASS FRACTION
TO LEO BFR BFR TAKEOFF
WEIGHT mT WEIGHT M LBS THRUST M LBS # RAPTORS
5.0% 4000 8.8 10.6 21
4.5% 4444 9.8 11.7 24
4.0% 5000 11.0 13.2 26
3.5% 5714 12.6 15.1 30
Assuming Raptor engine bells are ~1.6m wide, it's likely that 1st stage diameters of over 10m are preferred with 12.5m or even better 13.5m best to allow for max # of engines in case mass fraction drops. A smaller MCT dry weight really helps reduce BFR mass & # of engines as would be expected.
One way to speculate about the BFR launcher for the MCT is to look at mass fraction to LEO efficiency. The F9 has a mass fraction of ~2.7% to LEO. Despite full re-usability most of us amazing peoples/girls here expect that SX will somehow improve on that with the fully re-useable BFR. The tables below 1st assume 180mT to LEO with the dry MCT massing 80mT and the 2nd table assumes a dry MCT at 100mT. Various optimistic mass fractions yield different BFR takeoff weights. A T/W ratio of 1.2 is assumed yielding 1st stage thrust and dividing by 500KLBs/Raptor, the # of Raptors needed.
MCT + Payload = 180mT to LEO
MASS FRACTION BFR BFR TAKEOFF
TO LEO WEIGHT mT WEIGHT M LBS THRUST M LBS # RAPTORS @ 500KLB
5.0% 3500 7.7 9.2 19
4.5% 3889 8.6 10.3 21
4.0% 4375 9.6 11.6 23
3.5% 5000 11.0 13.2 26
3.0% 5833 12.8 15 31
MCT + Payload = 200mT to LEO
MASS FRACTION
TO LEO BFR BFR TAKEOFF
WEIGHT mT WEIGHT M LBS THRUST M LBS # RAPTORS
5.0% 4000 8.8 10.6 21
4.5% 4444 9.8 11.7 24
4.0% 5000 11.0 13.2 26
3.5% 5714 12.6 15.1 30
Assuming Raptor engine bells are ~1.6m wide, it's likely that 1st stage diameters of over 10m are preferred with 12.5m or even better 13.5m best to allow for max # of engines in case mass fraction drops. A smaller MCT dry weight really helps reduce BFR mass & # of engines as would be expected.
Are you counting the mass of MCT as payload, or as the stage itself? If it is it's own 2nd stage, then you should figure 180 or 200mt -gross- to LEO.
F9R 1.1 can put approx 13.1mT to LEO with recovery of the first stage. That might improve for 1.2 but we dont have numbers on that yet.
Second stage weights around 3.9mT while F9R weights about 505mT at launch. That makes a gross mass to orbit of about 3.36%.
But that is with KeroLOX all the way. With MethaLOX, alone that number would increase significantly. So yes, I guess it is safe to assume that the mass to orbit fraction will be around 4 to 5 % as you listed in your table. If SpaceX has enough unicorn hair and ferry dust left from their Dragon production, it might even go higher than 5%, simply because there is no payload adapter, fairings or what have you necessary.
Another interesting discussion is how much of the total delta V budget of the total LV is going into the booster and how much is going into the MCT (second stage)?
I think that there will be a big difference from EELV LV's where most of the fuel is in the booster with a small second stage.
1. You will not be able to land the booster on a barge, it has to come back to a well prepared and solid landing site. This means you cant get to far away (unless you have an island to land on).
2. You will have to have more thrust capacity on the MCT (second stage). But you will need this anyway if you are going to get of Mars.
3. The MCT (second stage) is going to need a high delta-V both for high energy transfer to Mars (3-4 month as stated by Elon) and to get from Mars (delta-v budget of 6-9 km/s). This means you can take advantage of this when staging to LEO.
Is there any flaws in my reasoning or mayor points I have missed?
Another interesting discussion is how much of the total delta V budget of the total LV is going into the booster and how much is going into the MCT (second stage)?
I think that there will be a big difference from EELV LV's where most of the fuel is in the booster with a small second stage.
1. You will not be able to land the booster on a barge, it has to come back to a well prepared and solid landing site. This means you cant get to far away (unless you have an island to land on).
2. You will have to have more thrust capacity on the MCT (second stage). But you will need this anyway if you are going to get of Mars.
3. The MCT (second stage) is going to need a high delta-V both for high energy transfer to Mars (3-4 month as stated by Elon) and to get from Mars (delta-v budget of 6-9 km/s). This means you can take advantage of this when staging to LEO.
Is there any flaws in my reasoning or mayor points I have missed?
First, I am not an aerospace engineer (I'm a EE) so I'm not sure I know what I'm doing.
I have reached the exact same conclusions you cite above.
Unlike the F9, the BFR/MCT will probably have a 1st stage that stages "low & slow" making boostback to launch site a given. And yes, the 2nd stage, a.k.a. MCT will after orbital re-fueling need sufficient delta V to escape LEO & transit to Mars in a few months and will also later need sufficient delta v to launch from Mars (having refueled again on the surface) and return to Earth or HEO. All these requirements dictate a high delta v capability and consequently a larger 2nd stage to 1st stage weight & propellant capacity design point than required for a simple LEO/GEO launcher. My latest models have a 1st stage with 14-15 million LBS thrust & 7 Raptors (slight overkill) powering the 2nd stage. The Vacuum Raptors are assumed to have ~610 thousand pounds thrust following the same 1.22 vac/sea level thrust ratios of the Falcon 9.
Think of the MCT as a near SSTO that fell short but gets enough boost from the stage one BFR such that it's good to go.
(I just hope that the 2nd stage doesn't have that explode just before staging feature thingy the F9R has)
(I just hope that the 2nd stage doesn't have that explode just before staging feature thingy the F9R has)
Assuming that event has something to do with the helium pressurization, it won't. Both methane and LOX tank will have self pressurization, no helium involved, I am sure.
...Excellent. I have seen a lot of discussion with the SEP option, but not the propellant only option. Do you know where I could find those designs -- on this forum?
...
There have been several MCT designs that had large enough tanks to do the 4 month transfer direct from LEO. (I think the jump from 4 to 3 months is pretty insane most synods).
If you think those tanks are too large, and want to transfer to a higher energy orbit before TMI, you should just use one extra MCT instead of 4. That will save hundreds of tons of propellant because you are moving less dry mass. A reusable SEP tug is another popular option.
I am not sure where to put this.
Yesterday there was an interview with Hans Koenigsmann in german TV ZDF. He repeated the argument that rockets need to be reusable like airplanes. He added that planes fly for decades and rockets will not fly that much but it should be 100 flights. He did not specify if this would be the Falcon Family or the goal for BFR/MCT.
Do you happen to know the show he said that in? It might be possible to still see it in the zdf mediathek from within Germany.
Do you happen to know the show he said that in? It might be possible to still see it in the zdf mediathek from within Germany.
It is available.
http://www.zdf.de/ZDFmediathek#/beitrag/video/2452860/ZDF-heute-journal-vom-21-Juli-2015
The part with Hans Koenigsmann is near the end. Skip through most of it.
Edit: It's at 21:40
For all non-German speakers:
[...]
(just to elaborate on what was mentioned above)
Thank you. I usually do enjoy reading transcripts done by other people ;-), but you are right, there wasn't really anything new.For all non-German speakers:
[...]
(just to elaborate on what was mentioned above)
Thx Marcon. I thought about writing a translated transcript but decided its not worth it since he doesnt mentions anything new. Your summary is way better than either nothing or a transcript. Thx.
We know now first hand that they are not aiming low. 2-5 reuses would only help to make their competetive situation better which is already very good.
Again, using existing solar panels on Mars, how big an area would 100kw of solar panels cover? Of course it will only work effectively 8-10 hours a day whereas a small nuke unit will go 24-7.
10kg per m^2 for just the arrays sounds really high given modern thin films.
Thin films just placed on the ground would have lower efficiency per square meter due to dust and lack of reflectors... but probably significantly better per kilogram.
The cells in IKAROS were 25 micrometers thick - if they were amorphous silicon, that's something like 58 grams per m^2!
It might need to be somewhat thicker on Mars due to wind, but even so, I think you could do way better than 10kg per m^2.
The weight will come more from good uv protection than the panel itself.
So it might be months before the first MCT can return.
The weight will come more from good uv protection than the panel itself.
Not necessarily. UV protective coating can be thin and invisible. I learned that when I built a roof for my terrace. The transparent polycarbonate panels come with an UV coating on one side. You don't even see that. The panels need a marking for the side that goes up. If you mount them wrong side up they don't last.
Not that I know much about this, but hard UV straight from the Sun is rather different from the bit of it that gets through our atmosphere.
It hinges on the mass of the ISRU equipment and power systems. If a complete automated deployment turn-key system capable of refueling the MCT-Lander in 1 synod will fit within one such Lander then I see no reason to ever abandon any mechanically sound lander.
Rather you go right into propellant production, return the first vehicle and leave the ISRU equipment in place. This achieves the two most important goals, 1) Have propellant in place before crew is risked, 2) Validate the entire round-trip flight of the vehicle before crew is risked.
The only reason to temporarily or permanently 'strand' an expensive vehicle on Mars is if the ISRU equipment is so massive that it needs to be broken-up over multiple landers, but all of my estimates show that it should easily fit within one landers 100 mT capacity (and finish in 1 synod), provided that the return propellant mass is not some absurd amount like 1000 mT.
This achieves the two most important goals, 1) Have propellant in place before crew is risked, 2) Validate the entire round-trip flight of the vehicle before crew is risked.
Shouldn't the fuel production operation be validated before sending crew, because without in situ production Mars return is extremely difficult, and a rescue mission would need huge amounts of fuel and organisation. On the other hand if the fuel is already there, rescue isn't a likely developement.This achieves the two most important goals, 1) Have propellant in place before crew is risked, 2) Validate the entire round-trip flight of the vehicle before crew is risked.
Is that actually necessary? It might be easier just to carry, say, 3 synodic periods' worth of life support supplies...
Shouldn't the fuel production operation be validated before sending crew, because without in situ production Mars return is extremely difficult, and a rescue mission would need huge amounts of fuel and organisation. On the other hand if the fuel is already there, rescue isn't a likely developement.
About 700 metric tons. So semi absurd ;-)
One of the MCT will make a good storage tank. It has the storage capacity required for the second MCT. If you don't keep an MCT in place where will you store the fuel? So the first MCT of all will stay in place, roll out large solar arrays to get power and produce and store fuel for the second ship's return. Or it could switch out with the second ship, as long as it leaves all the production equipment in place.
Personally, I expect the first 2, possible the first 3 MCT to be entirely remote controlled and to not return.
How much energy does it take to extract the fuel from the air and water, and how quickly do we want to do it? That is what fundamentaly sets the power required, isn't it? So many many solar arrays at first, because there will not be a nuclear reactor developped in the next few years, unless things change dramatically on the energy front.
Here is a possible MCT propellant tank arrangement, CH4 and O2.
This achieves the two most important goals, 1) Have propellant in place before crew is risked, 2) Validate the entire round-trip flight of the vehicle before crew is risked.
Is that actually necessary? It might be easier just to carry, say, 3 synodic periods' worth of life support supplies...
And where does the ERV get its propellant from, to get everything from Mars orbit back to Earth? Any propellant you bring along for a return trip essentially counts as payload mass. (a bit less than 1-1 for payload to Mars surface, since you don't have to take it down to mars and back up again, but you do have to carry it all the way from Earth surface to Mars orbit)
About 700 metric tons. So semi absurd ;-)
One of the MCT will make a good storage tank. It has the storage capacity required for the second MCT. If you don't keep an MCT in place where will you store the fuel? So the first MCT of all will stay in place, roll out large solar arrays to get power and produce and store fuel for the second ship's return. Or it could switch out with the second ship, as long as it leaves all the production equipment in place.
Personally, I expect the first 2, possible the first 3 MCT to be entirely remote controlled and to not return.
How much energy does it take to extract the fuel from the air and water, and how quickly do we want to do it? That is what fundamentaly sets the power required, isn't it? So many many solar arrays at first, because there will not be a nuclear reactor developped in the next few years, unless things change dramatically on the energy front.
Here is a possible MCT propellant tank arrangement, CH4 and O2.
I think your assuming direct Earth return, but I believe just return to orbit and docking with an ERV is the way to go and would put propellant needs at ~400 mT.
The use of the lander vehicle as the storage tank is good and something I've been assuming for initial missions, eventually a tank farm would be set up but that's likely to be after permanent habitation has begun. Likewise the 'swap' of returning in a different vehicle then the one landed in is a strategy I've advocated for.
I'm in favor of SEP for all in space propulsion, this reduces round trip propellant needs to a fraction of what they would be if chemical propulsion is used, all the studies on SEP show significant reduction in IMLEO for the same delivered payload that's why they are becoming the standard for mission planning.
And where does the ERV get its propellant from, to get everything from Mars orbit back to Earth?
I'm in favor of SEP for all in space propulsion, this reduces round trip propellant needs to a fraction of what they would be if chemical propulsion is used, all the studies on SEP show significant reduction in IMLEO for the same delivered payload that's why they are becoming the standard for mission planning.
SEP from LEO means slow spiralling out of LEO through the van Allen belts. BFR will make fuel in LEO really cheap. ISRU makes fuel on Mars cheap. I don't see SEP as competetive.
And where does the ERV get its propellant from, to get everything from Mars orbit back to Earth?
Impaler apparently meant SEP, but there is another alternative. You fuel the ERV in Mars orbit from the surface ISRU facility. (Similar to the fuel depot on the Earth side.) The SSTO cargo-landers ferry fuel up to the ERV. By leaving most of your long-duration, heavily shielded, interplanetary infrastructure in orbit, you can reduce the dry-mass of the landers to the bare minimum. The ERV is the "ship", the landers are "boats" used only for the initial and final legs. This also means that you can use fewer smaller landers doing multiple trips to ferry the cargo/passengers down to Mars, rather than one-big-lander-per-100t of payload. Again, that may let you reduce the size of the SSTOs to something more manageable.
Even better, if you can put enough prop in Mars orbit, your incoming cargo-landers can do a deorbit burn that is a significant proportion of the entry velocity, several km/s, greatly simplifying the design of the landers.
["Ah", you say, "but when ferrying fuel, the landers will still have to reenter at full orbital velocity!" Yes, but they will be empty. As, most likely, will any MCTs used as shuttles/ferries to LEO on the Earth side; launch full, reenter empty. But on the Mars side, the cargo MCTs need to carry the full 100t payload down to the surface.]
Obviously, I don't think this is the model that Musk is going for, judging by the clues that have been dropped. But this concept may still allow an increase in scale after the basic (self-contained) MCTs have established the core infrastructure on Mars. Let those giant, 100 tonne MCTs become the mere surface ferries for an even larger main interplanetary transport. That's how you go from 50-100 hundred colonists per synod, to "a million in my lifetime".
Also 3 synod periods is 6.5 YEARS, this is absolutely beyond the limits of our ability to keep food from spoiling, not to mention the MASS, at 5 kg a day of consumables a 4 person crew would need nearly 50 mT of supplies. And 5 kg a day is conservative when you realize it includes all your spare for fixing EVERYTHING.
Lastly we can not expect a crew to be in remotely sane or healthy after that time period, regardless of what a certain Sci-Fi movie remake of Robinson Crusoe might have lead people to believe any person stranded on Mars is as good as dead. We are going to be stretching all our technology and physiological means to the maximum just to do a mission of 1 synod.
and a rescue mission would need huge amounts of fuel and organisation. On the other hand if the fuel is already there, rescue isn't a likely developement.
This makes no sense, if we have not validated the vehicles ability to return to Earth (as in if it will SURVIVE re-entry at Earth) then the risk to the crew is not mitigated by giving them more supplies.
Also 3 synod periods is 6.5 YEARS, this is absolutely beyond the limits of our ability to keep food from spoiling,
not to mention the MASS, at 5 kg a day of consumables a 4 person crew would need nearly 50 mT of supplies.
Lastly we can not expect a crew to be in remotely sane or healthy after that time period, regardless of what a certain Sci-Fi movie remake of Robinson Crusoe might have lead people to believe any person stranded on Mars is as good as dead. We are going to be stretching all our technology and physiological means to the maximum just to do a mission of 1 synod.
Additionally, no automated unfurling system is needed. Just set the spool on the ground and roll it out- over the rocks and everything, then plug it in to the junction box (on board power will be DC). Hammer in some ground stakes every 2m.
A UV coating will be needed. They make that stuff in "space application" strength, so there's no new technology there.
Well, I was assuming there would be one or more missions coming the next synod anyway. MCT isn't for one-off missions, it's meant to be a colonization infrastructure.
In a one-off mission scenario, yes it needs to be self-sufficient. But that's not MCT.Quote
MCT is not part of some MarsOne no-return suicide pact, people are going to do exploration missions first and they will come home promptly. Even once a base is established people will rotate in and out for a long time before anyone even thinks about settling permanently. A return option MUST exist at the time the first person sets foot on Mars.Also 3 synod periods is 6.5 YEARS, this is absolutely beyond the limits of our ability to keep food from spoiling,
No, it isn't.
You have no idea what your talking about. If you have some notion of canned or frozen food then you've completely blown the mass budget on the food alone as that would be ~75% water, the only practical way to send food is dry and that reduces it's self-life, even a conjunction mission to Mars spanning 1 synod pushes our tech to the limits.The 5 kg number assumes zero recycling, which pretty much can't be the case for a MCT that will be capable of carrying 50-100 people.
You know nothing about ECLSS is you think 5 kg a day is zero recycling, it is ISS current tech level based on nearly full closure of water and would be conservative once you factor in all the redundancy in parts and equipment needed. The MarsOne nonsense got shot down by MIT for exactly the same failure to consider spares.
http://sites.nationalacademies.org/cs/groups/depssite/documents/webpage/deps_063596.pdf
You have no idea what your talking about. If you have some notion of canned or frozen food then you've completely blown the mass budget on the food alone as that would be ~75% water, the only practical way to send food is dry and that reduces it's self-life, even a conjunction mission to Mars spanning 1 synod pushes our tech to the limits.
You know nothing about ECLSS is you think 5 kg a day is zero recycling, it is ISS current tech level based on nearly full closure of water and would be conservative once you factor in all the redundancy in parts and equipment needed
No one’s exactly sure how the transportation will work, but it’ll likely be something like this: the Mars Colonial Transporter will consist of two pieces—the giant, powerful first stage, and the second stage, which will also be the spacecraft. The first stage will launch a spacecraft into orbit, then come back down (landing propulsively), refuel, undergo a bit of maintenance, and head back up with another spacecraft. This will go on for a while in the weeks leading up to the point where Earth and Mars are next to each other in orbit. Then SpaceX will send up a tanker of some kind to refuel the orbiting spacecraft (which also functions as the second stage rocket, so it’ll have spent a lot of its fuel getting itself into orbit).
By the time the planets are in place, there will be a group of MCT spacecraft—what Musk calls the “colonial fleet”—orbiting the Earth, fueled up and ready to go, and at just the right moment, the fleet will take off for Mars.
Three-to-six months later, the spacecraft will get to Mars, descend through the atmosphere, and land propulsively. The people will get out, probably to a fun welcome celebration put on by the existing residents, and unload everything over the next few weeks.
About two years later, when the planets are again aligned, right around the time Earth is launching the next colonial fleet, the group of spacecraft that came to Mars two years earlier will head back to Earth, carrying anyone on Mars who’s over it.
Three-to-six months later, the spacecraft will arrive back on Earth, land propulsively, and head in for maintenance so they’ll be ready to head back to Mars in two more years.
A UV coating will be needed. They make that stuff in "space application" strength, so there's no new technology there.
Good to hear. I assume that you don't mean adding a glass panel on the front but some surface coating with not too much weight.
Minimum consumables requirements are something like:
–Water: 2 kg/person/day drinking + 0.2 kg/person/day for minimal washing. Probably more on long trips for better hygiene
–Oxygen: 0.8 kg/person/day for metabolic consumption (assumes exercise) + leaks + repressurization
–Nitrogen: mostly driven by leak rates, repressurization (e.g. for airlocks)
–Food: 1.8 kg/person/day (includes meal-level packaging) at ~380 kg/m3 density
–Be sure to account for both mass and volume
Minimum consumables requirements are something like:
–Water: 2 kg/person/day drinking + 0.2 kg/person/day for minimal washing. Probably more on long trips for better hygiene
–Oxygen: 0.8 kg/person/day for metabolic consumption (assumes exercise) + leaks + repressurization
–Nitrogen: mostly driven by leak rates, repressurization (e.g. for airlocks)
–Food: 1.8 kg/person/day (includes meal-level packaging) at ~380 kg/m3 density
–Be sure to account for both mass and volume
That's great. Sounds like my estimate of 2kg/person/day was very generous. Getting water and nitrogen from ISRU on Mars is a safe assumption. Also meal-level packaging would add a lot of unnecessary weight. For a long stay under gravity food can be stored in bulk. Maybe I should adjust my estimate from 2 kg to 1 kg for the duration of the surface stay.
Minimum consumables requirements are something like:
–Water: 2 kg/person/day drinking + 0.2 kg/person/day for minimal washing. Probably more on long trips for better hygiene
–Oxygen: 0.8 kg/person/day for metabolic consumption (assumes exercise) + leaks + repressurization
–Nitrogen: mostly driven by leak rates, repressurization (e.g. for airlocks)
–Food: 1.8 kg/person/day (includes meal-level packaging) at ~380 kg/m3 density
–Be sure to account for both mass and volume
That's great. Sounds like my estimate of 2kg/person/day was very generous. Getting water and nitrogen from ISRU on Mars is a safe assumption. Also meal-level packaging would add a lot of unnecessary weight. For a long stay under gravity food can be stored in bulk. Maybe I should adjust my estimate from 2 kg to 1 kg for the duration of the surface stay.
If you have ever had a LRP or freeze-dried meals for hiking, you know they make an MRE look like a gourmet meal. However, it beats starving. But you have to a reliable water supply or it's like eating salty sand and gravel.
Hopefully, the crew wouldn't have to resort to the emergency rations.
Why do you dismiss 3D printing as a primary source of spare parts?
Why do you dismiss 3D printing as a primary source of spare parts?
To be honest I would dismiss it for the moment too. However the food assumed by Impaler does not take into account bulk packing as we did. Also I don't see it as a valid assumption that an ECLSS completely new designed for MCT with all knowledge and experience available will need the same amount of spare parts as the present ISS systems do. It will be designed to be more robust and needing less spares. ECLSS on Mars is another item again. It is not part of what we discuss as consumables. It will be designed for Mars with completely different methods, mostly biological. More volume, more initial weight but more efficient to run for a long time.
Why do you dismiss 3D printing as a primary source of spare parts?
Why do you dismiss 3D printing as a primary source of spare parts?
Because I come from a family that has been in the tool-and-die, plastic injection molding and quality control industries for 3 generations, we know that these additive processes complement but do not replace traditional manufacturing processes.
Furthermore the reactions happening in most ECLSS equipment are high temperature and energy chemistry, that means metal and ceramic vessels, valves and catalysts, they can not be replaced with plastic widgets.
While their is some limited potential for packaging materials to be a source of plastic feed-stocks as soon as you start talking about metallic part printing your looking at bring large amounts of metallic feed-stock, and large amounts of secondary equipment for finishing, measuring and testing the parts created.
In summary 3D printing is not a Star Trek Replicator.
Agricultural systems will need to be redundant and as they are power intense there is the most potential for downtime from a disaster, needing reboot after a plant disease related flushing
MCT is not part of some MarsOne no-return suicide pact,
Even once a base is established people will rotate in and out for a long time before anyone even thinks about settling permanently. A return option MUST exist at the time the first person sets foot on Mars.
You have no idea what your talking about. If you have some notion of canned or frozen food then you've completely blown the mass budget on the food alone as that would be ~75% water, the only practical way to send food is dry and that reduces it's self-life,
You know nothing about ECLSS is you think 5 kg a day is zero recycling,
The MarsOne nonsense got shot down by MIT for exactly the same failure to consider spares.
I don't think food production via agriculture need be either complex/difficult or especially fragile.Agricultural systems will need to be redundant and as they are power intense there is the most potential for downtime from a disaster, needing reboot after a plant disease related flushing
Why power intense? Mars gets sufficient sunlight for plant growth -- probably more than, say, the Pacific Northwest temperate rainforests or England (due to clouds) which are both rather lush.
It should be rather simple to eliminate all plant diseases by not bringing them from Earth. Even if this fails, plant disease doesn't usually mean complete collapse of the system, as you'd have multiple species.
Your saying an entire mining industry needs to be established to support manufacturing of spare parts, to keep the ECLSS running to keep the FIRST LANDING of 4-6 astronauts alive for multiple synods so we can postpone having to figure out how to do a return trip???I am presuming 10 - 20 on the first expedition. I expect human exploration to branch out from the first landing site in the first synod by rover, maybe the 2nd, or 3rd synod will bring secondary settlements, Also by the 3rd synod the equipment to travel to other locations via high inclination orbit to refuel at a depot then land at any point on Mars and take off again to land at one of the settlements with propellant ISRU. At that point real exploration starts to happen, and by then the population is around 50 with maybe 20 or so of the people who had travelled to Mars having returned.
Did you not pay attention to the original premise of this tangent? I have been talking this ENTIRE time about initial landings of small numbers for exploratory purposes, you seem to be talking about Blue Mars level end-state total self sufficiency a century from now. Your scale and time range are so out of step with what I'm talking about it's like talking about how the James town settlers will will generate electricity.And I have been pointing out for almost as long that this is a thread about MCT and that is not about a few boots and flags - it is about settling Mars from the get go.
Any logical sequence would consist of first exploratory scouting missions, followed by outposts with personnel cycling in and out and then finally a permanent settlement once lots of infrastructure is built up and optimum sites are found.
You seem to be throwing your lot in with Vultur and the MarsOne nonsense of immediate colonization from the first footprint. This is not going to happen, it's like saying that Neil Armstrong should have colonized the moon rather then coming back.
Any logical sequence would consist of first exploratory scouting missions, followed by outposts with personnel cycling in and out and then finally a permanent settlement once lots of infrastructure is built up and optimum sites are found.
The one alternative I can see is if we do ISRU on Phobos or Deimos and supply many manned sorties to the surface of Mars before building the first settlements. This would still involve a permanent presence and the one reason I don't really elaborate on it here is the idea that it doesn't work all that well with the MCT model driving Mars settlement.
Yeah, these are seemingly absurd percentage improvements, however not impossible. The critical elements of the solution are rocket reusability and low cost propellant (CH4 and O2 at an O/F ratio of ~3.8 ). And, of course, making the return propellant on Mars, which has a handy CO2 atmosphere and lots of H2O frozen in the soil.
The design goal is technically 100+ metric tons of useful cargo per flight, so maybe more than 100 people can be taken. Depends on how much support mass is needed per person and the luggage allowable.
Avionics, sensors, communications, aspects of vehicle structure, landing pads and a few other things get better with scale, plus it is more fun to be on a cruise ship than a bus, so I suspect that the 100 people per flight number grows a lot over time, maybe to several hundred. Also, we could subsidize the equivalent of economy by charging a lot more for first class.
Factor in all of the above and getting below $100k/ton or person eventually is conceivable, as the trip cost is then dominated by propellant, which is mostly liquid oxygen at a mere $40/ton (although a lot of it is needed per useful ton of cargo). That would be really awesome!
The one alternative I can see is if we do ISRU on Phobos or Deimos and supply many manned sorties to the surface of Mars before building the first settlements. This would still involve a permanent presence and the one reason I don't really elaborate on it here is the idea that it doesn't work all that well with the MCT model driving Mars settlement.
This may or may not warrant an extended discussion. I do like that idea as I was always thinking of fuel ISRU at Phobos or Deimos. Do we know if MCT would be able to land with enough fuel to lift off again? It probably should not be much heavier than on a normal landing with 100t supplies. So with minimal life support for a small crew and very little cargo to maximise fuel. Still seems not enough with less than 100t to lift off and reach Phobos. Maybe entry from orbit at lower speed allows for some more payload.
Known 'facts' about MCT you can fit on the back of a stamp. A small one, that already has writing on the back.
Even things Musk has already said (which is practically nothing anyway) are liable to significant change. Just like most other things SpaceX have done. They are happy to change as they go along, as they discover more stuff.
New info from Musk about the Raptor and MCT(and sorry about the crossposting):Quote from: Elon MuskYeah, these are seemingly absurd percentage improvements, however not impossible. The critical elements of the solution are rocket reusability and low cost propellant (CH4 and O2 at an O/F ratio of ~3.8 ). And, of course, making the return propellant on Mars, which has a handy CO2 atmosphere and lots of H2O frozen in the soil.
The design goal is technically 100+ metric tons of useful cargo per flight, so maybe more than 100 people can be taken. Depends on how much support mass is needed per person and the luggage allowable.
Avionics, sensors, communications, aspects of vehicle structure, landing pads and a few other things get better with scale, plus it is more fun to be on a cruise ship than a bus, so I suspect that the 100 people per flight number grows a lot over time, maybe to several hundred. Also, we could subsidize the equivalent of economy by charging a lot more for first class.
Factor in all of the above and getting below $100k/ton or person eventually is conceivable, as the trip cost is then dominated by propellant, which is mostly liquid oxygen at a mere $40/ton (although a lot of it is needed per useful ton of cargo). That would be really awesome!
Looks like the Raptor will run oxidizer rich. That puts its niche even closer to the BE-4.
IMHO, a reality check on MCT has been long overdue as this thread wanders all over the place because there are no facts to pin us down.
SpaceX can't even build Dragon 2 without public support, and it's been delayed for years because of a lack of internal funding to make up for shortfalls in public funding measured in millions of dollars. Yet somehow SpaceX is going to commit to actually building and flying BFR and MCT with billions of their own dollars and no contracted return on the investment? I don't think so. Not a snowball's chance in Hades.
Actually that would be very inconsistent with everything that Elon Musk has done to date. From the start hyperloop was an Elon concept looking for a non Elon home. He introduced it that way and described his process for arriving at the idea publicly. He has been as open about his intent to colonize Mars. His first plan was to inspire people to want it by sending a greenhouse experiment to Mars. He even tried to make that work, and I suspect somewhere early on, he will pay homage to that original idea in some early launch of an unmanned craft to Mars, but if he does it will be as significant as the wheel of cheese on the first Dragon flight. He publicly explained each of the step changes in his rational towards getting people on Mars, and has all along since starting to build rockets said that he is building towards creating the infrastructure needed to settle Mars. If you consider his actions with Tesla and Solar City in the way he announces business development, everything he has done suggests that what he is doing right now with SpaceX includes several money making adjuncts that also carry it towards its ultimate goal. SpaceX also exists to help Tesla and Solar city if you haven't noticed, and if needed they will help SpaceX.
More likely IMHO is that we'll see a fleshed-out paper concept similar to Hyperloop. Then they will go fishing for governments to pony up the development costs in order to be occupants on the actual spacecraft.
That tactic may work, but SpaceX doesn't have that many supporters in Congress as a consequence of making everything under one roof.
Even if money does come in from a government source, which won't be possible to even have allocated until 2017 at the extreme earliest, we won't see parts manufactured for the spacecraft until sometime in the mid to late 2020's.
Now, could I be entirely wrong? Sure I could, I'd love to be wrong on this! But the Dragon2 development experience is telling me I'm not.
On the gripping hand, a large well-funded religious group could put up a few billion to have a planet all their own for their more fervent adherents with no problem at all. It worked for the Pilgrims, why not Scientologists!?!I hope it doesn't come to that. I personally would like to see only humanist motivations organizing off Earth communities and their rules and customs (and on Earth communities for that matter). Let us please avoid bringing prejudice and hate into space (eliminating nationalism as well as religion).
I rush to point out how much SpaceX has funded from their own pocket so far, and shown that they can continue to raise equity to fund their portion of developments (they just raised $1B).Like Tesla, SpaceX's 'pocket' is rapidly running out. The $1B was allocated to general funds, not just the satellite venture where it is much needed, which speaks to the shallow depth of the general funds. It is clear that the larger SpaceX plan depends on satellite income, much like Tesla's larger plan depends on model X, and then the model after that. And like Tesla, raising funds by diluting ownership is not sustainable at the level needed to compensate for foundering profits. I do think both companies can stay afloat, but this dangerous game was Musk's plan all along, and we should recognize how dangerous it is.
I personally would like to see only humanist motivations organizing off Earth communities and their rules and customs (and on Earth communities for that matter). Let us please avoid bringing prejudice and hate into space (eliminating nationalism as well as religion).Aldrin took communion on the moon, read the words of Jesus, and planted the American flag. This did not preclude the trip from being for all mankind. We should let our rules and customs be what they are, wherever we are, in the spirit of individual freedom.
I rush to point out how much SpaceX has funded from their own pocket so far, and shown that they can continue to raise equity to fund their portion of developments (they just raised $1B).Like Tesla, SpaceX's 'pocket' is rapidly running out. The $1B was allocated to general funds, not just the satellite venture where it is much needed, which speaks to the shallow depth of the general funds. It is clear that the larger SpaceX plan depends on satellite income, much like Tesla's larger plan depends on model X, and then the model after that. And like Tesla, raising funds by diluting ownership is not sustainable at the level needed to compensate for foundering profits. I do think both companies can stay afloat, but this dangerous game was Musk's plan all along, and we should recognize how dangerous it is.
I personally would like to see only humanist motivations organizing off Earth communities and their rules and customs (and on Earth communities for that matter). Let us please avoid bringing prejudice and hate into space (eliminating nationalism as well as religion).Aldrin took communion on the moon, read the words of Jesus, and planted the American flag. This did not preclude the trip from being for all mankind. We should let our rules and customs be what they are, wherever we are, in the spirit of individual freedom.
I rush to point out how much SpaceX has funded from their own pocket so far, and shown that they can continue to raise equity to fund their portion of developments (they just raised $1B).Like Tesla, SpaceX's 'pocket' is rapidly running out. The $1B was allocated to general funds, not just the satellite venture where it is much needed, which speaks to the shallow depth of the general funds. It is clear that the larger SpaceX plan depends on satellite income, much like Tesla's larger plan depends on model X, and then the model after that. And like Tesla, raising funds by diluting ownership is not sustainable at the level needed to compensate for foundering profits. I do think both companies can stay afloat, but this dangerous game was Musk's plan all along, and we should recognize how dangerous it is.
I'm confused again. Do folks here believe that Musk has said the MCT is 200mT to LEO (then refueled and onto Mars' surface) with 100mT useful payload? Or is it 100mT useful payload with X mT more being the Dry Mass of what it takes to carry that payload...airframe, empty fuel tanks, re-entry shield, engine tonnage?
I'm confused again. Do folks here believe that Musk has said the MCT is 200mT to LEO (then refueled and onto Mars' surface) with 100mT useful payload? Or is it 100mT useful payload with X mT more being the Dry Mass of what it takes to carry that payload...airframe, empty fuel tanks, re-entry shield, engine tonnage?
IMHO, a reality check on MCT has been long overdue as this thread wanders all over the place because there are no facts to pin us down.
SpaceX can't even build Dragon 2 without public support, and it's been delayed for years because of a lack of internal funding to make up for shortfalls in public funding measured in millions of dollars. Yet somehow SpaceX is going to commit to actually building and flying BFR and MCT with billions of their own dollars and no contracted return on the investment? I don't think so. Not a snowball's chance in Hades.
I'm confused again. Do folks here believe that Musk has said the MCT is 200mT to LEO (then refueled and onto Mars' surface) with 100mT useful payload? Or is it 100mT useful payload with X mT more being the Dry Mass of what it takes to carry that payload...airframe, empty fuel tanks, re-entry shield, engine tonnage?
I think that Musk has said that the MCT will deliver 100t to Mars via a trip to Earth orbit for refueling. How much it can put into LEO was not a matter of record from Elon and speculation here has been anything from 70t to LEO to 200t to LEO.
My personal expectation is that MCT will have an dry weight around 50t, carry 100t of payload, 670t of propellant and that a tanker version that is reusable to LEO has a dry weight of 30t, carries virtually no payload but 800t of propellant at launch and can nominally deliver 130t of that to a depot. Thus 5 to 6 tanker flights per MCT launch to Mars (note that I am presuming a ΔV budget of 6km/s).
You seem to be throwing your lot in with Vultur and the MarsOne nonsense of immediate colonization from the first footprint. This is not going to happen, it's like saying that Neil Armstrong should have colonized the moon rather then coming back.
Your delta V budget seems a little low from LEO.
Figures I've seen for reaching Mars surface from LEO run higher...aerobraking away a Km/sec or two?
http://www.lr.tudelft.nl/index.php?id=29335&L=1
I'm confused again. Do folks here believe that Musk has said the MCT is 200mT to LEO (then refueled and onto Mars' surface) with 100mT useful payload? Or is it 100mT useful payload with X mT more being the Dry Mass of what it takes to carry that payload...airframe, empty fuel tanks, re-entry shield, engine tonnage?
The core of the capability falls on several questions
5) Does MCT need to *return* 100 tons of cargo to Earth orbit?
The core of the capability falls on several questions
5) Does MCT need to *return* 100 tons of cargo to Earth orbit?
I believe Musk said in the past year or so that MCT would be able to carry about 25 tons back to Earth, which adds up pretty nicely if you assume the 4 month Marsbound trajectory that he has talked about and a 75ish ton dry mass.
Although, it's quite possible I'm misremembering something.
Regardless, it really doesn't make sense to design for 100 tons both ways because the return flight is more demanding in delta-V.
Excellent post Mike, with a few reservations.
The one I want to tackle first is 1-synod operation.
By what mission plan can an MCT be used once per synod? Would this be an opposition-class mission that refuels from prelanded ISRU assets?
@Burninate
I vote for D) None of the above. 100 is the number of people going to Mars in colonization mode. It is not anticipated and planned for that so many people will ever go back to earth. That number may be closer to 10 max. That could be provided for with 25t return mass. Provided the ECLSS for 100 people does not have too much weight by itself which could reduce the max number of people going back further or part of the ECLSS would need to be removed and go back on empty cargo MCT to maximize passenger capacity.
The core of the capability falls on several questions
1) Is MCT's structural/rocket-stage mass counted within this 100 tons?
2) How many pieces on the board are there: Will non-landed transit habitats be used?
3) Is MCT's human cargo, life support, & food counted within this 100 tons?
3.5) Is MCT's habitat integral to the design?
4) Is MCT's ISRU gear counted within this 100 tons?
4.5) Is MCT's ISRU gear integral to the design?
5) Does MCT need to *return* 100 tons of cargo to Earth orbit?
5.5) Is MCT's ISRU gear returned?
The core of the capability falls on several questions
1) Is MCT's structural/rocket-stage mass counted within this 100 tons?
2) How many pieces on the board are there: Will non-landed transit habitats be used?
3) Is MCT's human cargo, life support, & food counted within this 100 tons?
3.5) Is MCT's habitat integral to the design?
4) Is MCT's ISRU gear counted within this 100 tons?
4.5) Is MCT's ISRU gear integral to the design?
5) Does MCT need to *return* 100 tons of cargo to Earth orbit?
5.5) Is MCT's ISRU gear returned?
My answers:
1) No
2) BFR booster, MCT (cargo+crew which are similar as crew hab mainly acts as cargo), tanker. No
3) Yes - MCT payload is either cargo or hab + crew + consumables
3.5) Partly - the hab is plumbed in for crew flights, it can be removed/replaced but only with considerable effort, cargo MCT would not have the hab, but would have some cargo containment system.
4) No - some of the initial flights have as cargo ISRU
4.5) No - no reason to carry ISRU back to Earth, many reasons not to.
5) No - only 25 tonnes. Aborts might return a full 100 tonnes.
5.5) No - ISRU remains on Mars were it forms part of a ISRU farm.
I think that with higher delta-v, fast transits and fast turn-around on Mars 1-synod are possible. If I have time I might try and work out if this is actually true, and how large the delta-v, transit time and entry velocities are.I understand the desire, in the same way that I understand the desire for quarter-synod operations.
Your delta V budget seems a little low from LEO.
Figures I've seen for reaching Mars surface from LEO run higher...aerobraking away a Km/sec or two?
http://www.lr.tudelft.nl/index.php?id=29335&L=1
Your link gives 5,7km/s for LEO to LMO. Landing should require less. That's assuming the LMO figure is with propulsive braking. Not going into LMO saves a lot and much of the braking for landing is done with aerobraking. If I remember correctly usually 1km/s propulsive braking was usually assumed for Mars landing.
I'm confused again. Do folks here believe that Musk has said the MCT is 200mT to LEO (then refueled and onto Mars' surface) with 100mT useful payload? Or is it 100mT useful payload with X mT more being the Dry Mass of what it takes to carry that payload...airframe, empty fuel tanks, re-entry shield, engine tonnage?
First: "200mT to LEO" is something that probably corresponds to BFR the launch vehicle rather than to MCT the upper stage & lander. Anything which lands 100mT of useful payload on the Martian surface, will be much higher than 200mT IMLEO, and this implies that one MCT mission will be the culmination of multiple propellant-carrying, and possibly multiple payload-carrying launches.
With that said, there are still big questions.
The core of the capability falls on several questions
1) Is MCT's structural/rocket-stage mass counted within this 100 tons?
2) How many pieces on the board are there: Will non-landed transit habitats be used?
3) Is MCT's human cargo, life support, & food counted within this 100 tons?
3.5) Is MCT's habitat integral to the design?
4) Is MCT's ISRU gear counted within this 100 tons?
4.5) Is MCT's ISRU gear integral to the design?
5) Does MCT need to *return* 100 tons of cargo to Earth orbit?
5.5) Is MCT's ISRU gear returned?
First let me add another few questions that I feel are vital:
A) Are there going to be a propellant depots, or are MCT's going to be refueled by a succession of tanker rendezvous?
B) does each MCT have regenerative ZBO or does it rely on passive cooling?
C) where would propellant depots be located?
D) where will BFR launch from?
MCT is its own second stage. This seems to be confirmed by WaitButWhy blog.
MCT is its own second stage. This seems to be confirmed by WaitButWhy blog.
I have the disagree with the very idea that this WaitButWhy blog can be considered a confirmation of ANYTHING. The writer doesn't claim any privileged access to information beyond what we have on these forums and I strongly think that everything they described came FROM this forums speculation or the speculation of similar forums. Even if the author independently arrived at similar conclusions that's nothing more then another 'vote' for one a particular configuration.
MCT is its own second stage. This seems to be confirmed by WaitButWhy blog.
I have the disagree with the very idea that this WaitButWhy blog can be considered a confirmation of ANYTHING. The writer doesn't claim any privileged access to information beyond what we have on these forums and I strongly think that everything they described came FROM this forums speculation or the speculation of similar forums. Even if the author independently arrived at similar conclusions that's nothing more then another 'vote' for one a particular configuration.
No he discussed this at length with Elon.
MCT is its own second stage. This seems to be confirmed by WaitButWhy blog.
I have the disagree with the very idea that this WaitButWhy blog can be considered a confirmation of ANYTHING. The writer doesn't claim any privileged access to information beyond what we have on these forums and I strongly think that everything they described came FROM this forums speculation or the speculation of similar forums. Even if the author independently arrived at similar conclusions that's nothing more then another 'vote' for one a particular configuration.
No he discussed this at length with Elon.
I don't think Elon let slip all that much privileged information. Hell, I don't think Elon has finalized all that much information, but that's another matter. He's organizing the community's conjectures, that's all. Plus a 3.8 F/O ratio. He quotes Elon where Elon provides info.
Note he says "No one’s exactly sure how the transportation will work, but it’ll likely be something like this: "
I find it unlikely that Elon would specifically request that some piece of speculation be cut from an article (who has the time to wade through that morass of an article), especially when it is many others have already been speculating the exact same thing and he is is still entertaining it himself. We know Elon has through about direct Earth return and might PREFER that, but it doesn't prove it will work and if it can't work then he can't use it. It is the idea that this configuration was in anyway CONFIRMED that is bogus.
The only bit of information that has any provenance back to Elon is the O/F ratio and that's an extremely minor detail. If their was a 'long discussion' it must have consisted of either Musk describing his dreams of colonization without going into detail, or the author asking every basic 3rd grader question that could have been answered by reading 'shit that Elon says'.
MCT is its own second stage. This seems to be confirmed by WaitButWhy blog.
I have the disagree with the very idea that this WaitButWhy blog can be considered a confirmation of ANYTHING. The writer doesn't claim any privileged access to information beyond what we have on these forums and I strongly think that everything they described came FROM this forums speculation or the speculation of similar forums. Even if the author independently arrived at similar conclusions that's nothing more then another 'vote' for one a particular configuration.
No he discussed this at length with Elon.
I don't think Elon let slip all that much privileged information. Hell, I don't think Elon has finalized all that much information, but that's another matter. He's organizing the community's conjectures, that's all. Plus a 3.8 F/O ratio. He quotes Elon where Elon provides info.
Note he says "No one’s exactly sure how the transportation will work, but it’ll likely be something like this: "
Elon had a veto on what went in the article and certainly provided plenty of background that counts as access to me. No this article did not say that a two stage design was set in stone, but neither is it something that he just got by reading the forum here or guessing.
Note that Elon has endorsed the latest WBW article by tweeting a link to it twice. He has not to my knowledge ever tweeted a link to speculation we have on this site saying "oh hey they are pretty close"
@Burninate
I vote for D) None of the above. 100 is the number of people going to Mars in colonization mode. It is not anticipated and planned for that so many people will ever go back to earth. That number may be closer to 10 max. That could be provided for with 25t return mass. Provided the ECLSS for 100 people does not have too much weight by itself which could reduce the max number of people going back further or part of the ECLSS would need to be removed and go back on empty cargo MCT to maximize passenger capacity.
Potential interpretations:
A)
the consumable budget is going to be considerably closer to refined food powder & oils than ISS' partially-dehydrated whole food panty...
@Burninate
I vote for D) None of the above. 100 is the number of people going to Mars in colonization mode. It is not anticipated and planned for that so many people will ever go back to earth. That number may be closer to 10 max. That could be provided for with 25t return mass. Provided the ECLSS for 100 people does not have too much weight by itself which could reduce the max number of people going back further or part of the ECLSS would need to be removed and go back on empty cargo MCT to maximize passenger capacity.
If the ECLSS on passenger MCT was made up of, say, 5 identical modules then after arrival on Mars, 3 modules could be removed for use in ground habitats leaving the other 2 for the return trip. Saves weight and also rotates new equipment on every subsequent flight. Similarly, gas and water storage tanks.
If the ECLSS on passenger MCT was made up of, say, 5 identical modules then after arrival on Mars, 3 modules could be removed for use in ground habitats leaving the other 2 for the return trip. Saves weight and also rotates new equipment on every subsequent flight. Similarly, gas and water storage tanks.
If the ECLSS on passenger MCT was made up of, say, 5 identical modules then after arrival on Mars, 3 modules could be removed for use in ground habitats leaving the other 2 for the return trip. Saves weight and also rotates new equipment on every subsequent flight. Similarly, gas and water storage tanks.
I agree that everything really useful on Mars would be removed. Quite possibly storage tanks and bunks and dividers for the passengers may be useful. The ECLSS units not so much. ECLSS on Mars would be based on plants. Greenhouses that produce food enough to eat will also produce enough oxygen. Technical recycling like on spacecraft is not needed. So even if some of them need to be removed to maximise passenger capacity they would go back on cargo MCT for reuse unless single components like pipes, valves or fans are valuable on Mars.
It also depends on how much the weight of this equipment is. If the 25t return mass allow return of everything to earth and 10 people and supplies they may remove very little, only items of really high value on Mars.
The 100 passenger phase is likely to be deep into the 21st century, 2070 or thereabouts. Rocketry has seen very little radical development since the 60s so we could be seeing the MCT still flying, but there could also be a whole host of unforeseen events. Bigelow building a cycler to serve the Earth-Mars routes.
The 100 passenger phase is likely to be deep into the 21st century, 2070 or thereabouts. Rocketry has seen very little radical development since the 60s so we could be seeing the MCT still flying, but there could also be a whole host of unforeseen events. Bigelow building a cycler to serve the Earth-Mars routes.
My impression was, this design exercise was for a vehicle & mission architecture which would work without thus-far-imaginary technologies like induced torpor, would be capable of equipping for 100-person missions and most of Musk's other requirements, and would have first mission in the 2030's and first launch in the 2020's.
That's consistent with the Wait-but-Why article, which says that Elon described the early process as sending up the MCT, returning the BFR and sending up tankers to refuel. With more MCTs it'll make sense to make a depot so that this can allow the MCT to be sent up last, getting a refuel quickly before departing. But at the beginning using small reusable tankers makes more sense.A) Are there going to be a propellant depots, or are MCT's going to be refueled by a succession of tanker rendezvous?A) Elon Musk mentioned depots. They will be needed when many flights go every launch window. However I believe that they won't need them early on. Two or three MCT, one of them passenger, can easily be fuelled directly.
380 vacuum ISP will definitely reduce both stage's mass. Rocket equation. i'll update.
Any other errata?
MCT Mass 180 mT
S2 Mass w/MCT 1025 mT
S2 Mars 25mT Cargo 8.5 Km/sec Rocket Equation
MCT Mass 180 mT
S2 Mass w/MCT 1025 mT
S2 Mars 25mT Cargo 8.5 Km/sec Rocket Equation
I don't think those three go together.
A stage with 180mt dry mass, 8.5km/s delta v and 25mt payload has a wet mass of 1980mt.
MCT Mass 180 mT
S2 Mass w/MCT 1025 mT
S2 Mars 25mT Cargo 8.5 Km/sec Rocket Equation
I don't think those three go together.
A stage with 180mt dry mass, 8.5km/s delta v and 25mt payload has a wet mass of 1980mt.
In the context of what he wrote it was 80t dry mass and 100t cargo outbound and 25t cargo inbound
MCT Mass 180 mT
S2 Mass w/MCT 1025 mT
S2 Mars 25mT Cargo 8.5 Km/sec Rocket Equation
I don't think those three go together.
A stage with 180mt dry mass, 8.5km/s delta v and 25mt payload has a wet mass of 1980mt.
In the context of what he wrote it was 80t dry mass and 100t cargo outbound and 25t cargo inbound
MCT Mass 180 mT
S2 Mass w/MCT 1025 mT
S2 Mars 25mT Cargo 8.5 Km/sec Rocket Equation
I don't think those three go together.
A stage with 180mt dry mass, 8.5km/s delta v and 25mt payload has a wet mass of 1980mt.
So we're saying that the oft stated 200mT MCT landing on Mars is impossible?
So we're saying that the oft stated 200mT MCT landing on Mars is impossible?
I think we need to consider entry velocity when looking at any lander, it's the #1 driver of TPS, structural mass and retro-propulsion needs. Most lander heritage involves direct entry from interplanetary transit and very high entry velocity. The Viking lander was the exception and is probably the best comparison because it doesn't drop nearly as many parts along the way. It had a landed mass of 62% of entry mass compared with 10% for the Phoinex lander.
The lander would launch from Earth atop a 2 stage launch vehicle and would be loaded with ~50 mt of propellants providing ~900 m/s DeltaV to be used for emergency separation and propulsive landing in the even of an abort. This would put the total launch mass at 225 mt. And the lander could then be fully topped off via another launch of and transfer of 250 mt or propellant from a stretched 2nd stage tanker. Transit to Mars is then done via a hybrid propulsion system, first a large SEP vehicle would slow push the lander to high Earth orbit where crew would board via a dragon capsule.
Isn't PICA supposed to be significantly better than previous TPS materials though? Would that reduce the required dry mass?
The human lander concepts I referred to are all designed for either entry from Mars orbit or aerocapture into orbit + entry from Mars orbit. Granted, gs, heat peak rate and heat load are significantly higher in the aerocapture phase than in the entry phase, although its a lot better than direct entry.
"Parts dropped along the way" are part of the structural mass at entry, which I was refering to. I.e. the entry mass minus the fuel and payload.
The best lander in terms of payload to structural entry mass I cound find is in the pdf attached on page 15, top left. A value of ~1.2. With SIAD, from orbit. It has a total entry mass of only 20t though. The higher the entry mass the worse usually.
With HIAD it may get better, haven't seen the mass break down of such a lander yet. HIAD could be interesting for MCT.
Wait, you want to use SEP for transfering a 475t payload from LEO to HEO (e.g. LDHEO)? That would take about 8 years with a 300kw SEP.
Isn't PICA supposed to be significantly better than previous TPS materials though? Would that reduce the required dry mass?
PICA is the best ablative, but I intend for the lander to do repeated rapid cargo flights between mars surface and low orbit maybe as many as 100 such round trips before returning to Earth, no ablative can go through that many cycles so I reject it as an option.
As the Mars colony is built up the small SEP transit vehicle would be replaced with a large more powerful cargo hauler with extensive habitats aka the mother-ship. The lander would be fueled in LEO and used to rapidly deliver passengers (100 at a time) to the waiting mother-ship in high orbit and would then travel attached too it as the mother-ships speed is expected to be competitive with the earlier direct flight. At mars the lander is used to repeatedly ferry between surface and low orbit with containerized cargo being reloaded in orbit. At Earth the launch of chemical propellant and landers is almost eliminated in favor of launching naked cargo containers and SEP propellant to be loaded onto the mother-ship.
PICA is the best ablative, but I intend for the lander to do repeated rapid cargo flights between mars surface and low orbit maybe as many as 100 such round trips before returning to Earth
PICA is the best ablative, but I intend for the lander to do repeated rapid cargo flights between mars surface and low orbit maybe as many as 100 such round trips before returning to Earth, no ablative can go through that many cycles so I reject it as an option.
As opposed, I believe 50t could be the dry mass of a 820t fully loaded and fueled MCT, note that at most there would be 1.3km/s of ΔV or so a re-entry mass around 230t, and a landing mass approaching 150t.
PICA is the best ablative, but I intend for the lander to do repeated rapid cargo flights between mars surface and low orbit maybe as many as 100 such round trips before returning to Earth
Ah, I was assuming one landing, one launch direct to Earth.
MCT is supposed to be "land the whole thing".
So only one Earth launch, one Mars entry, one Mars launch, and one Earth entry between servicings (on Earth).
I think PICA-X will be used for MCT.
It is supposed to get the mission done as efficiently as possible, I'm making the case that this hybrid architecture is 1) Actually achievable without using pixie-dust and 2) more efficient to boot. Compared to that any 'supposed to be' is moot.
So I propose a vehicle designed to act as a ferry between the surface and orbit, able to land 100 mt and return to orbit with 25 mt PLUS enough propellant to make another landing on Mars. At 75 mt dry mass for the vehicle the landing requires 40 mt of propellants. The launch mass is 400 mt of which 260 is assent (4.1 km/s) propellant, 25 cargo, 40 return propellant (.8 km/s) and 75 the vehicle dry mass. Total propellant load is just 300 mt meaning propellant production on Mars can be a fraction of that needed for Direct return. The structural mass fraction at launch is a very conservative 18.75% and payload is 46% of entry mass (including propellant) both very reasonable figures.
The lander would launch from Earth atop a 2 stage launch vehicle and would be loaded with ~50 mt of propellants providing ~900 m/s DeltaV to be used for emergency separation and propulsive landing in the even of an abort. This would put the total launch mass at 225 mt. And the lander could then be fully topped off via another launch of and transfer of 250 mt or propellant from a stretched 2nd stage tanker. Transit to Mars is then done via a hybrid propulsion system, first a large SEP vehicle would slow push the lander to high Earth orbit where crew would board via a dragon capsule. Then the SEP would separate and the lander would make a lunar-Earth slingshot burn of ~1 km/s for fast transit to Mars and use propulsive capture and perhaps some airobraking to reach low mars orbit leaving just enough propellant for landing. On the surface the cargo is unloaded and a 25 mt return cabin is loaded in it's place. The SEP system would make a slow transit to mars arriving well after the lander and wait in low orbit for rendezvous after which it would bring the lander back to high Earth orbit, crew would disembark again via Dragon capsule and then the unoccupied combined vehicle would do a down-spiral to low Earth orbit where it would be refueled again and use it's propulsive capacity to reduce entry velocity to the 3.6 km/s velocity it can tolerate followed by landing on Earth for reloading and reintegration, SEP remains in orbit ready to be reused.
Total BFR launches would be 6, one for the lander, 2 tankers of chemical propellant, 1 for SEP hardware, 2 estimated for SEP propellant. In addition 2 trans-lunar Dragon capsule launches likely on Falcon Heavy. Each subsequent launch would need 5 launches when SEP is reused. The cycle would be one round trip every 2 synods. If philw1776 can run the numbers of the size of the BFR needed to do a 225 mt launch that would be most informative to see how it compares.
It is supposed to get the mission done as efficiently as possible, I'm making the case that this hybrid architecture is 1) Actually achievable without using pixie-dust and 2) more efficient to boot. Compared to that any 'supposed to be' is moot.
To be clear on this. You are arguing that you know better than SpaceX and Elon Musk?
Perhaps you should write it up as a paper.
It is supposed to get the mission done as efficiently as possible, I'm making the case that this hybrid architecture is 1) Actually achievable without using pixie-dust and 2) more efficient to boot. Compared to that any 'supposed to be' is moot.
To be clear on this. You are arguing that you know better than SpaceX and Elon Musk?
To be clear on this. You are arguing that you know better than SpaceX and Elon Musk?
Since the amount of information known about MCT outside of SpaceX can fit on a postage stamp, written with a crayon, and that SpaceX plans change considerable over time, how do we know this isn't the plan? No point in saying "MUSK SAID THIS", because aforesaid plans change.
I'm confident they will reach the same conclusion I have after figuring out (possibly the hard way) that Direct Earth return is impossible. People are being way to fast to grasp at nebulous ideas and speculations even if they come from Musk as THE ONE AND ONLY way it will be done, Musk like any good programmer tries to think of the simplest possible system he thinks could possibly work, we saw that with F9 reuse plans, they are now WAY more complex then originally planned, MCT will be the same.
As opposed, I believe 50t could be the dry mass of a 820t fully loaded and fueled MCT, note that at most there would be 1.3km/s of ΔV or so a re-entry mass around 230t, and a landing mass approaching 150t.
Does it really need that much delta-v to go from terminal velocity to landing?
I am with Gucky on this one. I dont think that the mission plan you described is realistic at the beginning.
1. Aerobraking has never been used before, for good reason.
2. SEPs are for small payloads, not ones of many mT. The amount of solar arrays needed to get any meaningful thrust would be ginormous.
3. Using first SEP and then chemical propulsion is really inefficient.
4. Your mission plan is as complicated as it possibly can get. Simple is most often more important than efficient. Any complication in the flight plan means a rats hole of additional design and engineering work and launch mass to avoid failures. The more things you have the more things can break. And if anything breaks, your mission is toast.
5. The answer to "direct return is impossible" would not be to make the plan more complicated but to reduce the payload mass. The 100 persons per MCT figure is probably something for the far future. Not something for the first try.
The plan you described is maybe something for the far future, when Humanity has figured Mars transport out and is on its way to optimize the process. When failure modes are known and engineers learn to cut the right corners. I am very skeptical for the beginning though.
Ok, lets make a back-of-the-envelope calculation.
A recent all-electric satellite Eutelsat 115 WEST B (2,205 kg) uses about 18 kW (4x 4.5 kW thrusters) to reach GSO. It takes 8 Month, so that is a reasonable timeframe for the not yet crewed MCT. Each thruster weights below 16 kg, say 15. So the mass of the spacecraft without the thrusters is 2141 kg. The solar arrays provide 18 kW of power for the thrusters. Wikipedia says that one gets about 300W/kg and 300W/m^2. That makes 60 kg of solar panels and 60 m^2 area. So the spacecraft without the engines and without the panels is 2081 kg.
Now, scaling that to the mass of the MCT of about 475mT in your design. Then the SEP tug would need 475/2* 18 kW= 4.25MW of solar, or 472/2 * 60kg approx 14 mT of solar panels which have a surface area of about 14000 m^2. Say launch that thing with an BFR that has a payload bay of 20m hight. Then the solar arrays need to have a wing span of about 700m. Since the individual elements cant be larger than the diameter, say 10m, that are 70 segments. Say we have 2 equal sized 350m panels (one on either side) and you use ISS type solar arrays, that would fill two boxes of about 5m height each.
I admit, I thought that the solar panels alone would be far larger than that. I thought that this alone would invalidate your concept. It is at these considerations borderline possible. But I did not include everything that needs to be included. I did not include the Xenon mass, the mass of the SEP tug structure, any electronics or transformers. No heat control system or other things that are needed.
sources:
https://directory.eoportal.org/web/eoportal/satellite-missions/a/all-electric
https://en.wikipedia.org/wiki/Solar_panels_on_spacecraft
https://en.wikipedia.org/wiki/Integrated_Truss_Structure#Truss_subsystems
Something I haven't seen many discussions on: Abort Modes
Is this somewhere where SEP may be superior? If you use a 6 months fast Hohmann transfer, you will naturally return to Earth in 2 years if you miss Mars for what ever reason. Swing out to 2 AU then back in again to meet up with Earth (source: the Case for Mars/Mars Direct).
This is fine for initial exploration, as you will only have a small crew (~4 people) as you only need to carry 2 years worth of supplies for 4, and you would arguable have to do that anyway.
But trying to carry 2 years worth of supplies for 100 people, and things get out of hand. Roughly 1000kg per person of supplies and you're looking at over 100 tonnes of just food, water and oxygen. We could just assume that nothing will go wrong, but it's not going to be fun when something does.
That's why IMO it makes sense to launch several MCT's in the same launch window - a small fleet - that way they could support and assist each other if a problem should occur.
Your basing your calculations on a freight SEP vehicle on a com-sats performance? That's like using a water bottle rocket to estimate the performance of a launch vehicle. Try using one of a million SEP design concepts for lunar tugs, ARM or Boeing's recent paper rather then this nonsense.
Ok, lets make a back-of-the-envelope calculation.What you did is assume away several factors: required delta V & burn duration, which determine both required power and required propellant. The SEP mass fraction of the MCT being identical to the 702SP is not a safe assumption.
A recent all-electric satellite Eutelsat 115 WEST B (2,205 kg) uses about 18 kW (4x 4.5 kW thrusters) to reach GSO. It takes 8 Month, so that is a reasonable timeframe for the not yet crewed MCT. Each thruster weights below 16 kg, say 15. So the mass of the spacecraft without the thrusters is 2141 kg. The solar arrays provide 18 kW of power for the thrusters. Wikipedia says that one gets about 300W/kg and 300W/m^2. That makes 60 kg of solar panels and 60 m^2 area. So the spacecraft without the engines and without the panels is 2081 kg.
Now, scaling that to the mass of the MCT of about 475mT in your design. Then the SEP tug would need 475/2* 18 kW= 4.25MW of solar, or 472/2 * 60kg approx 14 mT of solar panels which have a surface area of about 14000 m^2. Say launch that thing with an BFR that has a payload bay of 20m hight. Then the solar arrays need to have a wing span of about 700m. Since the individual elements cant be larger than the diameter, say 10m, that are 70 segments. Say we have 2 equal sized 350m panels (one on either side) and you use ISS type solar arrays, that would fill two boxes of about 5m height each.
I admit, I thought that the solar panels alone would be far larger than that. I thought that this alone would invalidate your concept. It is at these considerations borderline possible. But I did not include everything that needs to be included. I did not include the Xenon mass, the mass of the SEP tug structure, any electronics or transformers. No heat control system or other things that are needed.
sources:
https://directory.eoportal.org/web/eoportal/satellite-missions/a/all-electric
https://en.wikipedia.org/wiki/Solar_panels_on_spacecraft
https://en.wikipedia.org/wiki/Integrated_Truss_Structure#Truss_subsystems
Ok, lets make a back-of-the-envelope calculation.What you did is assume away several factors: required delta V & burn duration, which determine both required power and required propellant. The SEP mass fraction of the MCT being identical to the 702SP is not a safe assumption.
A recent all-electric satellite Eutelsat 115 WEST B (2,205 kg) uses about 18 kW (4x 4.5 kW thrusters) to reach GSO. It takes 8 Month, so that is a reasonable timeframe for the not yet crewed MCT. Each thruster weights below 16 kg, say 15. So the mass of the spacecraft without the thrusters is 2141 kg. The solar arrays provide 18 kW of power for the thrusters. Wikipedia says that one gets about 300W/kg and 300W/m^2. That makes 60 kg of solar panels and 60 m^2 area. So the spacecraft without the engines and without the panels is 2081 kg.
Now, scaling that to the mass of the MCT of about 475mT in your design. Then the SEP tug would need 475/2* 18 kW= 4.25MW of solar, or 472/2 * 60kg approx 14 mT of solar panels which have a surface area of about 14000 m^2. Say launch that thing with an BFR that has a payload bay of 20m hight. Then the solar arrays need to have a wing span of about 700m. Since the individual elements cant be larger than the diameter, say 10m, that are 70 segments. Say we have 2 equal sized 350m panels (one on either side) and you use ISS type solar arrays, that would fill two boxes of about 5m height each.
I admit, I thought that the solar panels alone would be far larger than that. I thought that this alone would invalidate your concept. It is at these considerations borderline possible. But I did not include everything that needs to be included. I did not include the Xenon mass, the mass of the SEP tug structure, any electronics or transformers. No heat control system or other things that are needed.
sources:
https://directory.eoportal.org/web/eoportal/satellite-missions/a/all-electric
https://en.wikipedia.org/wiki/Solar_panels_on_spacecraft
https://en.wikipedia.org/wiki/Integrated_Truss_Structure#Truss_subsystems
Not sure this has been discussed here. Is there a chance BFR gets a grasshopper type development vehicle?It has been discussed here: Developing BFR reusability first (http://forum.nasaspaceflight.com/index.php?topic=34665.0)
I know, its a back of the envelope analysis. Its quick and dirty, thats what estimations are for. Its a calculation to see if the concept makes sense in the most simplistic way. I expected that an SEP tug for such a huge payload is physically impossible to launch and I tried to prove that. I did not achieve that goal.
But that is not a prove that the concept works. Its just a failure to prove that it does not work. I hope you understand the difference.
Also, my initial information on Impalers design did only state "from LEO to HEO" which is as fuzzy as it can get and I cant possibly deduce dV requirements from that. I just assumed that GSO is roughly equivalent to what he envisions as HEO.
I still dont like Impalers concept for the other reasons I stated, even if a SEP tug is physically possible.
Ok, so you admit that you've already decided you don't like my proposed mission architecture before you did any actual research (and regardless of what the research says), and are now trying to 'prove' it is not feasible via some sloppy back of a napkin calculations that seem to be uninformed by ANY actual studies on large SEP vehicles of which their are dozens. And you expect me to do your research for you by providing these links? I think not.
If the return delta-v is that high, the vehicle would only ever be fully fueled on Mars, as it doesn't need nearly as much delta-v to get to Mars.
If the return delta-v is that high, the vehicle would only ever be fully fueled on Mars, as it doesn't need nearly as much delta-v to get to Mars.
It would likely be fully fueled in both directions. However the payload back to earth would be much lower to achieve the higher delta-v with the same amount of fuel.
If the return delta-v is that high, the vehicle would only ever be fully fueled on Mars, as it doesn't need nearly as much delta-v to get to Mars.
It would likely be fully fuelled in both directions. However the payload back to earth would be much lower to achieve the higher delta-v with the same amount of fuel.
Also you could utilize the extra delta-v to make the booster a little smaller and maybe make the trip to Mars take less time.
It is the leg that requires most delta-v that steers the parameters of the rest.
If the return delta-v is that high, the vehicle would only ever be fully fueled on Mars, as it doesn't need nearly as much delta-v to get to Mars.
It would likely be fully fueled in both directions. However the payload back to earth would be much lower to achieve the higher delta-v with the same amount of fuel.
It would likely be fully fueled in both directions. However the payload back to earth would be much lower to achieve the higher delta-v with the same amount of fuel.
Maybe, but the return payload has to be large enough to accommodate people + life support according to Musk, and less fuel means fewer tanker launches needed to fuel up the vehicle going to Mars.
If the return delta-v is that high, the vehicle would only ever be fully fueled on Mars, as it doesn't need nearly as much delta-v to get to Mars.
It would likely be fully fuelled in both directions. However the payload back to earth would be much lower to achieve the higher delta-v with the same amount of fuel.
Also you could utilize the extra delta-v to make the booster a little smaller and maybe make the trip to Mars take less time.
It is the leg that requires most delta-v that steers the parameters of the rest.
Actually, the dry weight of the MCT (and total mass on a return flight) also is very important in this equation since the ΔV budget of a fully fuelled craft could be significantly different when it departs LEO from when it departs Mars surface.
This, the dry weight is too small and encloses too large of a volume (spreading it very thin) to be able to survive re-entry on Earth which is a minimum of 11 km/s. Even once we take into account for the 75% reduction in cargo on the Earth return leg.
This was proposed on one of these threads, but I have no data on its plausibility. Can you provide any references on the concept of superpressure as lighter-weight substitute for structural strength during reentry?This, the dry weight is too small and encloses too large of a volume (spreading it very thin) to be able to survive re-entry on Earth which is a minimum of 11 km/s. Even once we take into account for the 75% reduction in cargo on the Earth return leg.
The volume would be pressurized so is stable when reentering head on. Only when it has slowed down a lot it would pivot over for flying engines first for landing. I have speculated before that they may pressurize for more than 1000 millibar for reentry.
Going for good mass fraction prohibits massive walls like on capsules.
And one of my standard arguments. :) The designers at SpaceX are certainly aware of that problem and have at least tentatively a solution. A Falcon 9 first stage cannot do it because it is too long and slender. A second stage or a MCT has different proportions.
This was proposed on one of these threads, but I have no data on its plausibility. Can you provide any references on the concept of superpressure as lighter-weight substitute for structural strength during reentry?
Internal pressure doesn't prevent you from burning up which is my main concern.
>
How heavy is a Dragon heat shield?
Googling I find 848 kg for Apollo heat shield, and PICA-X is supposed to be better than previous materials.
>
Internal pressure doesn't prevent you from burning up which is my main concern.
That's what PicaX is for. At its size I think MCT would be lighter per volume and per surface than a Dragon capsule. Especially with a small cargo entering earth's atmosphere at high speed. The shield would be thick and strong on the tip but could be thin at the sides.
PICAX is nice stuff but no one even knows how to seal seems in it yet so it can only be used in monolithic pieces, assuming this limitation is over come their are other issues.
PICAX is nice stuff but no one even knows how to seal seems in it yet so it can only be used in monolithic pieces, assuming this limitation is over come their are other issues.
I guess these don't count as seams then? SpaceX figured out how to seal PICAX seams on their very first cargo dragon flight back in 2010.
PICAX is nice stuff but no one even knows how to seal seems in it yet so it can only be used in monolithic pieces, assuming this limitation is over come their are other issues.
I guess these don't count as seams then? SpaceX figured out how to seal PICAX seams on their very first cargo dragon flight back in 2010.
Yes, I'm not sure where this idea comes from. PICA-X does not have to be monolithic. Statements like this makes me question your other assertions, Impaler.
S2 Mars Return 25mT Cargo 8.5Km/sec Rocket Equation
There should be no problem fitting the 5 Rvac nozzles into the 12.5m MCT or even inside a 10m.
8.5 Km/sec is probably a half Km/sec too low
Now I'm confused. Just calculated Mars' escape velocity and it's only ~5 Km/sec without needed allowance for gravity losses.
But I still feel 7.5km/s ΔV is sufficient, 8.5 or 9 should actually allow a much faster return if it was available.
Now I'm confused. Just calculated Mars' escape velocity and it's only ~5 Km/sec without needed allowance for gravity losses.
Yes but being lower escape velocity you get a lower benefit from hyperbolic velocity (Oberth effect) so you need more ΔV over escape for trans Earth injection from low Mars orbit than you do for trans Mars injection from low Earth orbit.
I do not intend to interrupt your current dV conversation. Its very interesting, so please go on. But I want to make a point that bugs me quite some time already.
There is one thing, I do see not enough in L2 and here in the open forums BFR/MCT designs.
They do not seem that anyone takes an evolution of the BFR+MCT design into account. I would suspect that the first versions of the BFR+MCT are much less fleshed out, much less capable than the announced 100mT cargo / 100 passengers to Mars surface. In fact, I would expect them to have less than half of that.
Just an example for early missions that do not need the full, stated MCT capacity. Its just an example, it does not need to go down that way.
* The first MCT goes to Mars and stays there. Having a fuel production plant on board. But no people and no intention to get back. It would require a way to collect water, witch probably is the largest challenge.
* The second MCT would be dedicated to make it back. Using the fuel production of the first lander, store some food, a precursor for Humens. On its way back, it might bring some rocks as well. But really, it would be a demonstrator of getting back.
* The third might be a ship that brings furthe supplies as a precurser to human arrival. It could function as an MCT in case one of the next human rated MCTs can not make it back.
* The forth might have some humans on board. The mission would be: survive and come back. And the equipment would be triple and quadruple redundant to make that happen. No base as of yet. No habitat, MCT will have to do. That and the last MCT in case something breaks.
* The fifth MCT might get the first crew that stays longer. With the mission to create a base. It might be accompanied by 2-3 cargo MCTs. That all are intended to get back.
All this will likely happen with a less capable craft, as explained above. I would expect a major redesign of the magnitude F9 to F9 1.1 and Dragon to Dragon2. But still after this redesign, BFR and MCT might not have the capability that Musk is promising and that you are guys designing right now.
My arguments completely neglects the uses of BFR/MCT for LEO. At the current state, I would expect a precursor to BFR+MCT that is capable of payload to LEO of around 70 to 100 mT. It might be enough for the Mars missions stated above after one evolution step. It also might well be used to phase out FH and be used to deploy large satellite constellations that SpaceX and others are planning. I do not believe that SpaceX will come up with a BFR+MCT design that fulfils Elons statements in the next 20 years.
Why only 500 cubic meters storage? Seems way too small by over a factor of 2.Just a stretch goal. Total internal volume should be between 1000 to 1500, but some of it is in impractical shapes for storing vehicles, and we might want to keep some of the furnished areas for living space latter, or storing multiple communication antennas, or whatever. 5m3 per lander seemed ample.
Even 1000-1500m^3 is still too small for a volume needed to hold 100 people. It needs to be >2000m^3 or about 2500m^3. A 15m diameter MCT with a 30m tall payload section has about 2000m^3. This is what I expect this section of the MCT to be like. Plus larger diameters solves some other problems such as Mars entry terminal velocity values. In general it makes the height of the MCT a lot shorter and manageable. Larger diameters also increases prop tank volume for a specified height and decrease the ratio of tank weight to volume.Why only 500 cubic meters storage? Seems way too small by over a factor of 2.Just a stretch goal. Total internal volume should be between 1000 to 1500, but some of it is in impractical shapes for storing vehicles, and we might want to keep some of the furnished areas for living space latter, or storing multiple communication antennas, or whatever. 5m3 per lander seemed ample.
ML
The question is what is the point of developing an intermediate capacity rocket for SpaceX? The goal is a fully recoverable rocket. That's what drops the prices by orders of magnitude and makes Mars even thinkable. Once you stop destroying the rocket at each launch, then the cost become much closer to the fuel costs+ development costs. the development costs are a huge portion for rockets that aren't used much. So I would expect that Spacex to develop a single large core, and then use that for all possible variants. The BFR could be used with only partial loads and part of its fuel, for example. Or for smaller loads is could fly back to the launch pad, since this cuts down on payload by almost 50%. 125 tons with fly back, 250 tons with barge or a remote landing area. That's a wide range of payloads. As for the MCT itself, whatever extra capacity it has can be filled with spare parts, or even just stock materials; we're going to need a lot of spare parts for this program :-)
If a market develops, a modification of the second stage could carry up many different payloads using adapted fairings, without breaking the bank in development costs.
In a similar vein, the shuttle almost never flew fully loaded, and increased capacity by various optimisations with the external tank and the tiles, but it always remained outwardly identical.
Michel Lamontagne
Even 1000-1500m^3 is still too small for a volume needed to hold 100 people. It needs to be >2000m^3 or about 2500m^3. A 15m diameter MCT with a 30m tall payload section has about 2000m^3. This is what I expect this section of the MCT to be like. Plus larger diameters solves some other problems such as Mars entry terminal velocity values. In general it makes the height of the MCT a lot shorter and manageable. Larger diameters also increases prop tank volume for a specified height and decrease the ratio of tank weight to volume.Why only 500 cubic meters storage? Seems way too small by over a factor of 2.Just a stretch goal. Total internal volume should be between 1000 to 1500, but some of it is in impractical shapes for storing vehicles, and we might want to keep some of the furnished areas for living space latter, or storing multiple communication antennas, or whatever. 5m3 per lander seemed ample.
ML
Well, isn't Falcon Heavy, fully recoverable, the perfect precursor? Why another model in between? It can get some nice experimental payloads to Mars.Falcon Heavy will do somewhere in the vicinity of 35-40 tons to LEO semi-reusably. BFR + MCT will allegedly do somewhere in the vicinity of 200 tons to LEO reusably. FH is a good precursor for unmanned missions that test single subsystems, but some intermediate LV *would* be useful for less ambitious goals and larger subsystems, I suspect. 4.6m of payload fairing isn't a lot to play with.
I indeed hope there is a case! More ships=less cost. I just question the need for an intermediate step between a fully recoverable Heavy and the BFR.
About deltaV, can aerobraking reduce the deltaV required at Earth return or does it have too great a time penalty?
Well, many reasons why I believe there will be an incremental approach.
1. SpaceX always did an incremental approach. The didnt start with Falcon 9 1.1 when they developed their first rocket. Same holds true for the BFR. They need to learn to manufacture, launch, etc. such big rockets and thats easier if they dont jump into it with both feat.
2. There is a business case. They can probably much more cheaply launch large constellations of satellites. Like one or more of the internet constellations. Maybe even modules of a new space station. Maybe even a commercial space station, serviced by Dragon2 for space tourists.
3. All the key technologies for BFR+MCT can be tested on a smaller rocket with a smaller MCT precursor. That includes:Sure they can, but they can also be tested with an FH - at most making a new methalox upper stage.
* landing on MarsAvionics, precursor aerodynamics of shape if not total mass, autonomous operation can all be tested with FH and dragon derived spacecraft.
* large scale energy production on MarsNo this won't happen until enough people get there (>1MW) to layout and maintain solar farm, while it could be done with automated craft, it is no more difficult to get to this level from several 13,000kg craft than several 25,000kg craft)
* water production on Marsas above
* methane/oxygen production on Mars
* landing, retanking and MCT (precursor) return to EarthProbably not retanking, though it isn't impossible, just seems too cumbersome, but certainly a 13,000kg lander can have a sample return as part of it (or several can put samples in Mars orbit to be collected by a single return craft that was the SEP testbed).
* flight operations and trajectory executionAlready done with the above steps, I expect one of the first SpaceX payloads to Mars from a single FH launch will be the local network for Mars with low latency 150kg satellites in a constellation with one or two larger high latency Earth link satellites. This network needs to be in place to facilitate the automated precursor flights including the first few MCT's.
* communications
* high bandwidth, high latency internet access on Mars
* booster launch and recovery (using Raptor)plenty of time to do this to launch and fuel the first MCT. That is 8 flights right there. I am betting on 4 MCT's being launched in the two synods before the first manned flight
* second stage recovery (if any)As above, but if we made an alternate FH upper stage that was raptor powered maybe it would make sense to work out the major kinks. Note that going from an intermediate MCT re-entering to a full sized one is probably as big a step of development as going from Dragon 2 to the intermediate. Certainly it is not just scaling and will not require significantly less expense because you went through an intermediate step that cost you more than half as much as the last step. I am even afraid that MCT and tanker version upper stages might have enough differences to involve considerable unique costs.
* re-tanking in LEOThis could justify some work on a reusable FH upper stage right there. I think it needs to be done. However, it really can be done with just the planned FH.
* first human on Mars and survivability
* first habitat on Mars
* first plant growth on Mars using Mars soil
* first sustainable food growth on Mars
And that are just the biggies. All this needs to be ready for colonization. And all this is easier with a smaller version of BFR/MCT than the ultimate goal noted by Musk.
Once these technologies are in place and tested and ready, a full size MCT makes sense. I would not be surprised that after the logistical hassles are overcome with the first missions, a bigger version of MCT are produced, one that is closer or even capable of 100 humans to Mars. Even though this particular capability would not be used.
4. I dont think that SpaceX will have the funds to do it alone. They will need the help of governments, NASA might not even be enough. When in cooperation with space agencies, pork needs to be provided. That is far easier with a precursor MCTs that focus on technology development and science rather than direct colonization of a naked planet.
I've done a mass breakdown of my proposed MCT vehicle.
Thermal Protection Airoshell: 5 to 10 kg/m^2 metallic fully reusable, 650 m^2 5 Mt
Tanks and plumbing: 5% of 300 Mt propellant load 15 Mt
Landing Gear: 10% of landed mass based on F9-R ratio with F9 expendable 18 Mt
Raptor Engines: 1.5 Mt each based on 150:1 T/W ratio and 2300 KN force x4 6 Mt
Vernier Engines: Hover 175 Mt on Mars with 100:1 T/W ratio 1 Mt
Miscellaneous: Transit solar panels, thermal radiators, avionics, cargo handling 5 Mt
Structural Frame: 11% of Entry mass, carbon-fiber skeleton 25 Mt
Total Dry Mass 75 Mt
Cargo 100 Mt
Landing Propellant: Provide 800 m/s Terminal decent DeltaV 40 Mt
Total Entry mass 215 Mt
Here's a fully fueled 180mT MCT outbound from LEO (100mT cargo) compared with the launch back from Mars' surface with "only" 25 mT cargo
MCT Dry Wt & Cargo 180 mT
S2 Mass w/MCT 1025 mT LEO departure. Mars Return is 75 mT less
S2 Mass w/MCT 2.3 Million LBS
Stage 2 Km/sec 6.48 Km/sec Rocket Equation LEO
S2 Mars Return 25mT Cargo 8.5Km/sec Rocket Equation
Exponentials help when you reduce the mass. Just refuel with less propellant to reduce "excess" Km/sec.
The technical challenge is a lightweight MCT vehicle able to withstand Earth re-entry if that is the goal rather than return to some high Earth/moon orbit.
If anything the 225 Mt number for BFR launch mass I'm speculating is aggressive. I'm concerned about what the 2nd stage recovery cost will be and if this will kill the performance and lower the % of launch mass reaching orbit as philw1776 thought his own analysis looked high in this regard.I think you might have misunderstood me here. My suggestion is that ~200 tons launched from Earth to LEO seems like the limit, but that limit only has to cover spending on structure & permanently-attached habitat; One adds to that with supplemental launches and reaches ~600 tons at Mars entry, with the extra spent on food, people, ECLSS, ISRU gear, surface equipment, and descent propellent. Then you factor in propellant for Earth Departure and reach 2000-3000 tons IMLEO (4.5 to 6km/s).
If anything the 225 Mt number for BFR launch mass I'm speculating is aggressive. I'm concerned about what the 2nd stage recovery cost will be and if this will kill the performance and lower the % of launch mass reaching orbit as philw1776 thought his own analysis looked high in this regard.I think you might have misunderstood me here. My suggestion is that ~200 tons launched from Earth to LEO seems like the limit, but that limit only has to cover spending on structure & permanently-attached habitat; One adds to that with supplemental launches and reaches ~600 tons at Mars entry, with the extra spent on food, people, ECLSS, ISRU gear, surface equipment, and descent propellent. Then you factor in propellant for Earth Departure and reach 2000-3000 tons IMLEO (4.5 to 6km/s).
And that buys you a conjunction-class colony-in-a-can mission, for 25 people at least; And maybe 100 people with prelanded assets.
Humm, if I am launching from Mars surface and I want to enter an elliptical hohoman transfer around the sun I need to decrease my heliocentric velocity aka I need to be going slower then Mars itself after having left it's sphere of influence. Thus gravity loss incurred during escape may actually be beneficial IF it results in the loss in heliocentric velocity that one needs to reach Earth.
According to https://en.wikipedia.org/wiki/File%3aDelta-Vs_for_inner_Solar_System.svg you need 5.5 km/s DeltaV to reach escape
I just had a wild idea about the MCT design. What if it was inflatable by itself, without a separate habitat?Too complicated, this system is designed to break all current limitations. You want to keep it as simple as possible. If there is an inflatable easier to be a docked module until Mars, but I don't think is the case.
Bigelow habitats generally have a solid cylindrical core surrounded by an inflatable shell. But is there any particular reason why a conical shape wouldn't work just as well? Imagine a scaled-up SuperDragon enclosed in a layer of inflatable material which is puffed-up in vacuum.
The capsule would be launched deflated and only inflate once safely out of the atmosphere. When it's time to enter the atmosphere of Mars you can repack all the furniture back into the solid core and deflate the exterior. This pull brings the inflatable portion back behind the heat shield. On the surface of mars you can inflate it again and will end up with a sort of mushroom-shaped habitat. You can mount light-weight flooring inside the inflated region when on the surface..
The "inflatable" portion doesn't have to touch actually the heat shield. You will have side-mounted engines surrounding the heatshield, similar to the current Dragon.
Such a design requires that the skin material have some thermal resistance, similar to the sides of a normal capsule. It would likely need a fairing for flying upwards through the earth's atmosphere. The habitat also needs to support inflation/deflation while people are inside the core. Deflation seems particularly difficult. The simplest way would be to make core itself mostly airtight and pump out the air from the exterior portion.
This design only makes sense for shipping large numbers of people. You give them plenty of space during the transit period and on the surface but pack them closely together for takeoff and landing. It solves the "habitable volume" problem without increasing payload diameter too much.
I was thinking of something more akin to a gravitational slingshot in which the vehicle is ahead of mars in it's orbit and mars pulls it back, the vehicle loses velocity relative to the sun and mars gains it.
In a previous article, the same author, Richard Heidmann, seems to have fixed on the idea that BFR is an SSTO, that MCT has no main engine, just belly thrusters for landing. That's led to some weird conclusions, and hence the current article.
Nice concept. But quite frankly, I like the concept developed in L2 better. Hyperion et al. seem to have a better handle on the subject. Its very interesting to the development though and only good things can come from independent groups tackle the same problem.
Or how about this radical idea. Drop the abort capsule altogether. I've argued this before, but the idea seems to be very ingrained in people, almost like the idea of landing a shuttle with wings "horizontally" on Mars. Both ideas make just as much sense, IMO. (In other words, not much sense)
There is no place to abort to when landing on or taking off from Mars. If you find launch abort essential for Earth departure, shuttle people up while the MCT is in LEO being refueled.
That approach doesn't really work to maximize the rest of the vehicle.
EITHER:
The MCT launches full of fuel and has lots of delta V, which maximizes the payload mass fraction to orbit, but has a TWR too low to use for launch abort
OR:
[The MCT launches empty of fuel and has high TWR, but with low enough delta V that part or all of the BFR has to come close to, or reach, orbit. This lowers payload mass fraction by a large amount, or for a fixed payload makes the BFR required substantially larger.
OR:
The MCT launches with an order of magnitude more engine power than it will require for the rest of the mission, which is retained as waste mass throughout the rest of the mission, reducing space for other payload, or increasing launch requirements, by a large factor.
That approach doesn't really work to maximize the rest of the vehicle.
I ran across this MCT speculation/article, from the French chapter of the Mars Society: (my apologies if this has already been discussed)
ANALYSIS OF A CONCEPT IN MARCH COLONIZATION TRANSPORTATION (MCT) LAUNCHES TWO [google translation]
http://planete-mars.com/analyse-dun-concept-mars-colonization-transport-mct-a-deux-lancements/
If you have the chrome browser, it will automatically translate the site for you.
Here are some interesting images from the MCT architecture of this article:
Image 1: Two MCT's docking in LMO for propellant transfer
Image 2: Launch abort module interior
Image 3: Launch abort module from behind
Image 4: MCT landed horizontally on Mars
Image 5: MCT interior
Image 6: BFR base
Image 7: BFR/MCT stack
In is an interesting concept, but I'm not sure that horizontal landing is practical. And their MCT seems to be lacking any kind of engines for propulsion.
BTW, here is a link to the author's (Richard Heidmann) previous articles on the subject: http://planete-mars.com/author/heidmann/
He has really gone all in on the horizontal landing idea... Anything else doesn't seem to enter his mind as a possibility.
That approach doesn't really work to maximize the rest of the vehicle.
I don't see it that way. I already stated that it will not be able to speed away from an explosion without warning. But it can separate from a failing first stage with shut down engines. Being heavy it would have to burn a lot of fuel before it can land. There is that one point that it will not have enough time to go through a lengthy precooling period. That's why I asked if Raptor can be kept in a state ready for ingnition throughout launch.
Or how about this radical idea. Drop the abort capsule altogether. I've argued this before, but the idea seems to be very ingrained in people, almost like the idea of landing a shuttle with wings "horizontally" on Mars. Both ideas make just as much sense, IMO. (In other words, not much sense)
There is no place to abort to when landing on or taking off from Mars. If you find launch abort essential for Earth departure, shuttle people up while the MCT is in LEO being refueled.
At some point you've got to put "big boy pants" on if you want to go to Mars. Or roll a hard six. Or name your own analogy. :)
First crew sizes are likely to be quite small - under 15. So will fit into a couple of Dragon 2.
We've been told that payload back is ~ 25% of payload there, so ~ 25 tonnes. It is really hard to fit an abort system (capsule?) + hab for 100 (or even 15) into 25 tonnes - even if they return dry and empty.
Later flights might use an optimised crew MCT.
That approach doesn't really work to maximize the rest of the vehicle.
EITHER:
The MCT launches full of fuel and has lots of delta V, which maximizes the payload mass fraction to orbit, but has a TWR too low to use for launch abort
OR:
The MCT launches empty of fuel and has high TWR, but with low enough delta V that part or all of the BFR has to come close to, or reach, orbit. This lowers payload mass fraction by a large amount, or for a fixed payload makes the BFR required substantially larger.
OR:
The MCT launches with an order of magnitude more engine power than it will require for the rest of the mission, which is retained as waste mass throughout the rest of the mission, reducing space for other payload, or increasing launch requirements, by a large factor.
Nice concept. But quite frankly, I like the concept developed in L2 better. Hyperion et al. seem to have a better handle on the subject. Its very interesting to the development though and only good things can come from independent groups tackle the same problem.
A top mounted thruster system like this would be inherrently stable, unlike landing on a tail engine (F9 booster)....like a helicopter is stable with it's thrust up on top and it's mass hanging below it.
Nice concept. But quite frankly, I like the concept developed in L2 better.
GORDAP - I like it! If abort is only practical for Earth launch, it is an elegant solution.
The MCT will somehow need to be able to use Methalox for its thrusters - not sure how that will work. Would there be smaller LOX and Methane containers that are pressurized at high pressure (continuously refilled from the main tanks) that feed the maneuvering thrusters? One could also imagine such a system could be scaled up for abort/landing thrusters.
Agreed, a pushing fast-start 1 second abort motor located in the interstate is a very clever solution.
I think the solid booster solution is viable (though SpaceX seems to dislike all pyrotechnics). Also it doesn't need to really be the whole of the inter-stage, all you really need is a naked thrust-structure that connects the abort motor to the bottom of the MCT which can generate and transmit the ~9 million N of thrust. As that is a lot of force I would recommend just sticking cones into the nozzles of the Raptors and pushing on them directly (a bit like the new center-pusher stage separator) so your reusing the thrust structure of the vehicle itself as if the engines were running without having to create new hard-points that bypass the engines. As the Raptors start up they will simply push out the abort motors as they are burning out.
Based on the total impulse of the Shuttle boosters and their total mass I would estimate that to deliver 9 million N-s of total impulse would require 15 mT of solid rocket motors, quite heavy. But total take off mass might be lessened by the fact that your putting less abort propellents in the MCT itself. Still it is a hit to take on returning the stage it is on. If that is a first stage it should be do able without much problem, but if it's a 2nd stage as I've proposed it looks like it could be problematic.
So I propose to do the Dragon abort trick, use the abort engine to land the 2nd stage in the event that no abort happens. That means if their is an abort the 2nd stage is sacrificed even if the fault is in the first, but your probably fleeing an explosion that was going to destroy it anyway so no big loss. Now the abort system isn't as parasitic because we MUST have smaller vernier engines to land a second stage, the main propulsion system of Vacuum Raptor engines is not going to cut it as their are too few engines and they can't throttle low enough or safely at Sea Level. Though certainly the amount that we need for abort is over kill over what we would for just landing.
If a Super-Draco like liquid engine in used you would need A LOT of them, like 120 to get the same thrust, around a 10 m core you should just manage to get that on the vehicle in a giant ring of engines packed side by side. In reality I think we will see some new engine several times more powerful then Super Draco (SuperDuber Dracos??) and the number will be more reasonable. Still a large number of engines will make touch-down very easy, no only can you come to halt really fast but with such a large depth of throttle it should be possible to put the stage down on very delicate tooth-pick landing gear which should save considerable mass.
I am not sure an interstage abort system would work. The mechanical loads on the MCT would be 10 or more times stronger during an abort than during normal operation. The MCT including fuel would need an acceleration of 5+gs as opposed to less than 1. Plus the force to overcome atmospheric pressure at max Q. The parasitic mass would be within the metal structure of MCT instead of fuel and abort angines. I am not sure that is lighter than a crew abort capsule. Also you can't abort from an exploding MCT during first ascend on earth or ascent on Mars. An ascent abort on Mars might be required when a colony is already established and an infrastructure for rescue on Mars is available.
[LAS lifeboat]
I'd like to offer a completely different MCT LAS for consideration:The question is implicitly about the power to weight ratio of rocket engines.
Considering that the MCT will certainly have high efficiency Methalox engine(s) for its main thrust (and perhaps landing?) the issue seems to be the delay time required to 'spin up' the turbines in the case of a mid flight abort. So this means that the MCT really only needs a short burst (<4-5 seconds?) of high thrust (5gs?) to pull it safely away from a disintegrating 1st stage. Correct?
So, how about placing an LAS set of engines, with associated fuel, in the interstage between the booster and the MCT? Under normal conditions, the MCT would stage, leaving the interstage attached to the booster, and the booster plus interstage/LAS would RTLS and be completely reused. Under abort conditions, the LAS engines would fire and the interstage/LAS would separate from the booster and propel the MCT a distance away. The MCT main engine(s) (Raptors?) would spin up during this LAS firing period, then the MCT would detach from the interstage and proceed to do an abort landing.
The LAS could be fast acting, high thrust (poor ISP) hypergols, or even (gulp) solids. This system wouldn't need to be very massive, given its short firing duration, but it would admittedly still be considerable parasitic mass hurting the RTLS effort.
I like this idea because (1) no parasitic mass going to orbit and beyond, penalizing the whole system, (2) entire system is reusable, and (3) unlike 'puller' LAS systems, you don't require that the LAS system cleanly separate from the manned portion as an additional staging event, lest you have LOM.
Thoughts?
Impaler, love the idea of using these abort engines as landing engines. That way, their fuel is not parasitic at all! There may be more engines than are needed for landing (so I guess the unused ones are 'parasitic'), but these can also just serve as redundant backups, as they are on the Dragon 2.
I'm in the camp that's expecting the BFR to be a single stage booster, with the MCT serving as its own embedded second stage. If this is the case, then the LAS engines could be used on the BFR stage 1 to assist landing, and I'll bet their not much oversized for that mission. Hmm, I think the BFR would still use a center Raptor to do the big initial 'boostback' burn though.
You come perilously close to one of my ideas: using said Super-Dragon for Earth Abort, Earth Return, as MAV, and as Mars landing propulsion, then returning the main habitat & cargo section of MCT a year later. I am not yet decided on incorporating that element: It becomes favored if the ISRU equipment has a low propellant mass production ratio, and I don't have firm numbers for that.[LAS lifeboat]
Essentially you've added the cost of developing a 100-person Super-Dragon capsule, with the added cost of integrating it into MCT like a matryoshka doll. (And you rarely actually use it. Most of the time, all of its independent systems are just dead weight. And it won't have the develop-use-update-use pattern that SpaceX prefers.)
If you're going to the expense of developing an extra vehicle, why not just use the 100-person "lifeboat" as a shuttle to ferry passengers to LEO when the MCTs are fuelled and ready to go? It doesn't need to go to Mars.
I agree that this leaves no LAS for the launch from Mars phase of the MCT. But as others have pointed out, it's debatable whether or not an LAS is needed/useful for an SSTO vehicle (which the MCT is at this point).Remembering that the point of the MCT is that there IS a colony on Mars, one would assume that there will be a support and rescue service of some form. One of the side effects of "wild west" multi-company independent and only semi-regulated growth in a colony is that it will create a risk level for some colonists that NASA would never accept (and the implications of that will be important in many ways, but not for this thread). I actually suspect that the rescue service will be a private venture too, that's pretty well the point of this colonisation model.
In a previous article, the same author, Richard Heidmann, seems to have fixed on the idea that BFR is an SSTO, that MCT has no main engine, just belly thrusters for landing. That's led to some weird conclusions, and hence the current article.Of course once you get to 50km or so, you could flip around and use those belly thrusters to achieve Earth Orbit or use them for Mars injection. I mean, a little dumb, but still possible.
This won't weigh down the vehicle when it is serving in a cargo only flight (which will be probably x10 more numerous then crewed flights) and it allows us to change and modify the crew carrying modules for evolving needs and numbers, from small crews which will initially live inside the MCT to future high volume passenger counts which will immediately disembark.
Does cargo really have the same mass-volume relationship as human "cargo"?
So long as the cargo-hold volume is sufficient for the lowest density thing you want to transport then there's no problem, the denser items will just leave some volume unoccupied.
That's like asking why do we have airplanes that are full of only people and ones that are full of only air-freight. That's the nature of every mature transportation system to make a very sharp distinction of between a cargo carrying trip and a passenger trip, it makes the logistics simpler and maximizes passenger comfort.Virtually all large passenger aircraft also carry cargo because it's nice source of income. The logistics are not difficult. Passengers are not inconvenienced by ULD (http://vrr-aviation.com/uld-intro)s that contain cargo. There are financial advantages to being a "cargo only" airline, dealing with humans is a hassle. You need to provide food, flight attendants, etc. But there is no financial advantage to being "passenger only" and the airline business is not one that can afford to leave money on the table.
...
The topic of a Mars ascend LAS is of course highly debatable. If there is no infrastructure to rescue people after an ascend abort, there is no point in aborting in the first place. But if there is already an established colony, that might be a different topic.
If we learned anything from the first shuttle failure, it is that we need to have an abort mechanism. It is not possible to have a system "secure by design" it doesn't work that way.
Cargo MCTs don't need life support.
That's like asking why do we have airplanes that are full of only people and ones that are full of only air-freight. That's the nature of every mature transportation system to make a very sharp distinction of between a cargo carrying trip and a passenger trip, it makes the logistics simpler and maximizes passenger comfort.I think that's a fair analogy but raises my reasoning because it doesn't respond to that.
In spacecraft going to mars we have a huge savings in propellants if we go slow but this is deleterious to human health so we send them fast and freight slow, this is all the incentive we need to segregate passengers and freight trips. And their is almost certainly a reduction in marginal mass needed as passenger counts grow due to the redundancy needed for probabilistic equipment failure and the ability to time-share public space and amenities, that's why the service is better on a 747 then on a piper-cub.
A top mounted thruster system like this would be inherrently stable, unlike landing on a tail engine (F9 booster)....like a helicopter is stable with it's thrust up on top and it's mass hanging below it.
I'm rather surprised that you aren't aware of the Pendulum Fallacy: https://en.wikipedia.org/wiki/Pendulum_rocket_fallacy
Nice concept. But quite frankly, I like the concept developed in L2 better.
If you aren't willing to show/share it, don't bring it up here. This kind of information sand boxing is why I'm not joining L2 anytime soon. Just my personal opinion...
Lobo: Is that L2 vehicle concept posted anywhere where us plebs can read it?
As I'm sure you remember I favor the 2nd of your scenarios, "Another is launching MCT dry on a 2nd stage rather than the integrated design, with just enough propellant to power it's LAS engines. So the whole dry MCT is aborted and lands. For a nominal launch, it docks with the depot, and the 2nd stage does it's own EDL and returns to launch site."
I'd like to offer a completely different MCT LAS for consideration:
Considering that the MCT will certainly have high efficiency Methalox engine(s) for its main thrust (and perhaps landing?) the issue seems to be the delay time required to 'spin up' the turbines in the case of a mid flight abort. So this means that the MCT really only needs a short burst (<4-5 seconds?) of high thrust (5gs?) to pull it safely away from a disintegrating 1st stage. Correct?
So, how about placing an LAS set of engines, with associated fuel, in the interstage between the booster and the MCT? Under normal conditions, the MCT would stage, leaving the interstage attached to the booster, and the booster plus interstage/LAS would RTLS and be completely reused. Under abort conditions, the LAS engines would fire and the interstage/LAS would separate from the booster and propel the MCT a distance away. The MCT main engine(s) (Raptors?) would spin up during this LAS firing period, then the MCT would detach from the interstage and proceed to do an abort landing.
The LAS could be fast acting, high thrust (poor ISP) hypergols, or even (gulp) solids. This system wouldn't need to be very massive, given its short firing duration, but it would admittedly still be considerable parasitic mass hurting the RTLS effort.
I like this idea because (1) no parasitic mass going to orbit and beyond, penalizing the whole system, (2) entire system is reusable, and (3) unlike 'puller' LAS systems, you don't require that the LAS system cleanly separate from the manned portion as an additional staging event, lest you have LOM.
Thoughts?
GORDAP - I like it! If abort is only practical for Earth launch, it is an elegant solution.
The MCT will somehow need to be able to use Methalox for its thrusters - not sure how that will work. Would there be smaller LOX and Methane containers that are pressurized at high pressure (continuously refilled from the main tanks) that feed the maneuvering thrusters? One could also imagine such a system could be scaled up for abort/landing thrusters.
[LAS lifeboat]
Essentially you've added the cost of developing a 100-person Super-Dragon capsule, with the added cost of integrating it into MCT like a matryoshka doll. (And you rarely actually use it. Most of the time, all of its independent systems are just dead weight. And it won't have the develop-use-update-use pattern that SpaceX prefers.)
If you're going to the expense of developing an extra vehicle, why not just use the 100-person "lifeboat" as a shuttle to ferry passengers to LEO when the MCTs are fuelled and ready to go? It doesn't need to go to Mars.
The difference is, that in an abort, the MCT (assuming it is its own second stage) has to be abortet including its entire mass of fuel. I think I remember that the dry mass of the MCT is about 20 % of its wett mass. So the 3g in normal operation would be reached near dry mass. The 5 g abort would be reached at wett mass. The force on the metal structure is F=m*a, mass times acceleration. Wett mass is 5 times dry mass, acceleration is higher, and air resistance is higher. So during about, the structure of MCT must take I would guess more than 10 times the force of normal operation. The metal structure must be sized accordingly.
In the one stage BFR your looking at the combo 2nd stage MCT vehicle is going to be considerably more massive then the smaller and un-fueled one I'm looking at, that will mean probably 3x more thrust needed for separation. Rather then the 9 million N I'm estimating for a 200 mT MCT your looking at 27 million N, the thrust of 12 raptor engines. I see a problem with rapidly initiating that much thrust INSIDE the interstate when the vehicle is still assembled, it would be like setting off a bomb. So the thrust must be vented laterally with Dragon like canted rocket nozzles along the perimeter of the inter-stage. With the cosine loss were looking at something like 400 Super Draco equivalent rockets. Do able probably by having multiple bands of nozzles such that the entire outer surface of the inter-stage is nozzles. But we should be aware that were increasing the number of parasitic engines, though they parasatizing the first stage rather then the 2nd which means they count for only 1/10th as much in final payload.
The decisive factor I think is that in a 2 stage BFR with the abort engines atop the 2nd stage and below a smaller MCT is that you can abort during both first AND second stage burn. You only give up the abort motor once full Earth orbit has been achieved. Where as with a one stage BFR your in a black zone after first stage separation and yet your still attached to a potential bomb which has 6-7 raptor engines which might set the thing off.
6 posts in a row? Is there some obscure forum record you are trying to break, Lobo? ;)
Ok let me just touch base on terminology and what looks to be a set of acronyms that will describe them.
Super Dragon Mars Colonial Transport (SDMCT): A dragon shaped vehicle which lands with a bottom side heat-shield and is powered by either side-wall engines or a heat-shield penetrating central engine. It launches on a 2 stage rocket which is essentially an enormous F9 colloquially known as the BFR. Least imaginative design, initially popular with many but now falling out of favor. Would abort as Dragon dose by pushing away the whole vehicle from the 2nd stage.
Separate Bi-conic Mars Colonial Transport (SBMCT): A obviously bi-conic vehicle with bottom mounted engines and a horizontal high lift entry orientation that keeps engines and other systems in the rear. Propellant would be in the nose of the vehicle and cargo in the base just above engines. Launches on top of a conventional 2 stage rocket comparable to the SDMCT. Would abort similar to SDMCT by pushing the whole vehicle away from the 2nd stage. Designed around single stage to Low Mars orbit and a 'semi-direct' like architecture. My currently preferred solution.
Integrated Bi-conic Mars Colonial Transport (IBMCT) A modification of the SBMCT which 'integrates' it with the 2nd stage of the rocket it would launch on such that their is a bi-conic nose and a cylindrical aft body, the tip of the nose is an abort capsule, the next section down is cargo and propellant fills the cylinder. Launches on a single stage rocket which is similar to the first stages of the alternative configurations. Designed specifically for Single stage Direct Earth return. Lobo and others currently preferred solution.
Dose that look like an accurate depiction of the current positions?
We can all agree that MCT is SSTO on MARS, and that means the only means of abort would be disassembly of the vehicle via something like a nose-cone escape crew capsule.
This should be avoided because we shouldn't even HAVE an integrated crew cabin, if the MCT simply has a large cargo hold in which cargo or crew modules are interchangeably placed then it becomes possible to make a powered escape module and place this into the vehicle when ever crew are on board.
This won't weigh down the vehicle when it is serving in a cargo only flight (which will be probably x10 more numerous then crewed flights) and it allows us to change and modify the crew carrying modules for evolving needs and numbers, from small crews which will initially live inside the MCT to future high volume passenger counts which will immediately disembark.
Also note that when aborting to the Mars surface your possible landing sights are very nearly a great circle around Mars as your point of landing varies becomes a track of the landing point of a sub-orbital flight, if it is late in the MCT assent phase and thrust is lost then your looking at a landing sight possibly on the opposite side of the planet from where you launched. Not a feasible distance from a base for rescue to pick up survivors even if they were landed without a scratch, an extremely difficult matter when your basically performing full EDL which is a process which requires considerable heat-shields and retro-propulsion.
No one is going to be unloading cargo by HAND, your going to be using forklifts, scissor-lifts movers and other wheeled equipment to take out large containers which would then be taken into garages or coupled to pressurized areas, this is what we do already on ISS for Christ sake. The lack of basic knowledge of logistical functions here is maddening sometimes.
MCT doesn't need life-support because their is only a need for ONE vehicle variant which is a universal carrying shell, you put pressurized habitat modules (which has all the life-support equipment in it) in when you want to carry passengers and you put cargo containers in it when you want to carry cargo. Why is this so hard to understand? It is infinity more flexible and efficient then any other configuration.
The returning MCT should only have a small crew though, as most would be colonists left there.
The working idea is that all concepts would have a Dragon 2 like nose hatch (maybe about 5m wide or so) under which could be a docking collar. Tanker and depot dock in LEO and proceed to do a slow spin to settle propellants in the bottom of the tanks where it can then be pumped from the tanker to the depot, and later from the depot to MCT.
That last is where a separable lifeboat LAS has an advantage. Since MCT is a SSTO for Mars ascent, if there's an explosive problem with the main propulsion system, the crew is lost. That'd be the case for full vehicle abort of the SDMCT and NIBMCT. The LAS lifeboat can get away and then "fly" down range to a landing spot within it's ascent trajectory. What they do when they land safely is another story.
If in the early exploration phase, the returning crew is small. If the lifeboat is of sufficient size to support them until they can get to a backup vehicle somewhere else on the surface, then that would be good. If later in colonization phase, there'd be a full colony with presumably long range rovers that could be dispatched for rescue. The returning MCT should only have a small crew though, as most would be colonists left there. So you really shouldnt' have to rescue 100 people, unless there's some abort from orbit and the lifeboat lands somewhere far away form the intended LZ
The IBMCT would have 3 Raptors, but only two are needed for Mars ascent (maybe only one). It has 3 because 3 are needed to get it from booster staging to LEO. So it has engine-out backup for Mars EDL (with LAS assist) and it has engine out backup for Mars ascent. And if the MPS exploded, it's LAS can abort the lifeboat away.
The SDMCT and NIBMCT have only two engines, as they don't do Earth ascent. If there's an engine out at Mars ascent, there's some issue of how to handle it. One engine thrusting at full throttle should keep it aloft long enough to burn off enough propellant so the LAS system can land it again for an emergency landing. It might not be able to get to orbit, but it should be able to do a controlled emergency landing. And it depends when the engine were to fail, far enough up just one engine should still get it home.
And you have nothing in it to be a tanker (perhaps stretched tanks within the same Outer Mold Lines). Just a big empty volume.
And you have some cryo-refrigeration equipment in it as a depot.
And you put a payload bay behind the nose to deploy FH and D4H class satellites to GTO. The vehicle then comes back around on the GTO eliptical orbit to Earth, where it does a small deorbit burn and does Earth EDL. Not expendable PLF required.
The working idea is that all concepts would have a Dragon 2 like nose hatch (maybe about 5m wide or so) under which could be a docking collar. Tanker and depot dock in LEO and proceed to do a slow spin to settle propellants in the bottom of the tanks where it can then be pumped from the tanker to the depot, and later from the depot to MCT. Propellant lines would run from the nose to the tanks. (unless there's a way to pump liquids in zero g without using centrifugal forces? That's a bit of a gray area.)
The crew MCT would also have an access tunnel for access to the flight deck and on through to the hab area below it.
The Sat launcher would have a payload bay under the nose cap instead of a docking collar or access tunnel. Something like Rocketplane Kistler K-1 (below). The sat launcher has no need to refuel, so no need for a docking collar installed.
OML's are the same.
This is where I prefer the IBMCT. Just one vehicle, rather than than two including the dedicated SII. It's not an expendable "dumb" stage, it has to be it's own LEO reusable spacecraft in it's own right. And may or may not share much with the booster and MCT. There's extra cost in that.
The IBMCT would have the same tank tooling, bulkheads, etc as the booster. So just 2 pieces with a lot of commonality. That wouldn't be the case for the Super Capsule. And may or may not be for NIBMCT depending on it's design. If a full biconic, then it won't share much with anything else. If a cylinder/biconic then it might, but it will run into issues that using booster sized tank tooling and domes will make the LOX tanks too big using 12m domes, because the tanks are smaller than IBMCT. So you might need custom tankage anyway.
In Early exploration any abort to surface is fatal because their is zero rescue infrastructure on the surface and no conceivable capsule could carry sufficient supplies to see them through more then a few days. So I consider this a pointless abort. By the time you have an infrastructure to do surface rescue your passenger count is much too high for the small capsule your proposing, it would need to be a large vehicle comprising a significant portion of the whole mass of the vehicle and would present great difficulty in landing as you going to be falling on a ballistic trajectory from a high altitude and need massive retro-propulsion to not impact the surface, it in no way resembles the kind of un-powered capsule landing that can be done on Earth.
I don't know if you saw it earlier but I did a mass brake down of my MCT concept. I'd like to see what the IBMCT comes out at.
Thermal protection at 5 kg/ m^2 over an area of 650 m^2: 5 mT
Tanks and Plumbing 5% of 300 mT propellant mass: 15 mT
Landing legs, 10% of touch down mass: 18 mT
4 Raptor Engines at 150:1 T:W ratio: 6 mT
Vernier Engines that can hover on landing at 100:1 T:W ratio: 1 mT
Solar, Radiator and computer systems: 5 mT
Structural skeleton, 1/3rd of dry mass: 25 mT
Total 75 mT dry mass
In Early exploration any abort to surface is fatal because their is zero rescue infrastructure on the surface and no conceivable capsule could carry sufficient supplies to see them through more then a few days. So I consider this a pointless abort. By the time you have an infrastructure to do surface rescue your passenger count is much too high for the small capsule your proposing, it would need to be a large vehicle comprising a significant portion of the whole mass of the vehicle and would present great difficulty in landing as you going to be falling on a ballistic trajectory from a high altitude and need massive retro-propulsion to not impact the surface, it in no way resembles the kind of un-powered capsule landing that can be done on Earth.
This. You put it better than I could have. This is why IMO the way to increase safety for the MCT is to make he whole thing abort capable, through added redundancies in propulsion and systems.
If a MCT has to abort during a Mars ascent and land far down-range, the crew can survive for an *extended* period in the MCT. Not in a small capsule where everyone is squeezed into. Designing in a pointless separable abort capsule leads you down the path of terrible engineering trade-offs.
In Early exploration any abort to surface is fatal because their is zero rescue infrastructure on the surface and no conceivable capsule could carry sufficient supplies to see them through more then a few days. So I consider this a pointless abort. By the time you have an infrastructure to do surface rescue your passenger count is much too high for the small capsule your proposing, it would need to be a large vehicle comprising a significant portion of the whole mass of the vehicle and would present great difficulty in landing as you going to be falling on a ballistic trajectory from a high altitude and need massive retro-propulsion to not impact the surface, it in no way resembles the kind of un-powered capsule landing that can be done on Earth.
Isn't LOX heavier than Liquid methane? Also it will take twice the lox to burn the methane. So shouldn't the Lox tank be on bottom with the Methane on top?
I've said all along that an elongated 2nd stage would do tanker duty, and depots are in my opinion unnecessary, the MCT will act as it's own depot taking on propellants from visiting 2nd stages until it is full. As MCT must depart Earth with some propellant for EDL at mars and must then hold significant amounts while on the martian surface (which while cold is still warmer then cryogenic LOX), so the MCT will have to have significant long-term cryo-storage capabilities likely through a combination of insulation and cryo-coolers, thus it makes an excellent depot.
These cryo-systems along with radiators and solar arrays are the only systems that I would integrated into the vehicle. Human habitats placed into the MCT are simply plugged into these utilities much like an RV.
One significant factor that needs to be taken into account for a biconic vehicle (which I advocate) is a the mass distribution. It needs to be have the proper balance when almost empty (normal atmospheric entry & landing), and it needs to be able to also have the proper balance for a near full propellant load. (Earth or Mars abort) Also, when near empty of propellant the vehicle must also be properly balanced for a full cargo load vs empty.
To handle this range, the layout that makes most sense (IMO) is to put the cargo/crew in the middle of the vehicle - with the LOX tank above, and Methane tank below. This would allow a balanced biconic sideways reentry with ANY cargo load, and ANY propellant load.
EDIT: See image below for how that might look, in this DC-Y(?) drawing:
Again, this a concept in reaction to some who feel strongly the LAS is necessary. (I argued against an LAS, but was out voted, heh) It could very easily be left off, with no Mars ascent abort option (for an explosive event, you can abort to orbit with just an engine out), and a separate LEO-taxi with LAS for Earth ascent. For those that favor that, I think that's viable too.
One significant factor that needs to be taken into account for a biconic vehicle (which I advocate) is a the mass distribution. It needs to be have the proper balance when almost empty (normal atmospheric entry & landing), and it needs to be able to also have the proper balance for a near full propellant load. (Earth or Mars abort) Also, when near empty of propellant the vehicle must also be properly balanced for a full cargo load vs empty.
To handle this range, the layout that makes most sense (IMO) is to put the cargo/crew in the middle of the vehicle - with the LOX tank above, and Methane tank below. This would allow a balanced biconic sideways reentry with ANY cargo load, and ANY propellant load. Or to balance cargo above and below propellant tanks.
EDIT: See image below for how that might look, in this DC-Y(?) drawing:
The IBMCT would have most of the dry mass in the nose and MPS as the aft (Engines, thrust structure, etc). Between the two would empty main propellant tanks, a cargo deck, and volumous, but relatively light Hab area. The cargo deck would be between the hab volume and the tanks, and may have a fairly heavy mass when loaded with surface cargo. So for Mars EDL, that would be about in the middle. So you have your greatest mass areas in the nose, in the tail, and [roughly] in the middle. So it shouldn't be too bad. There will be some residuals in the tanks to power Raptor from terminal velocity to hover.
I don't know you need two separate tanks top and bottom like in this concept.
Except, what do you want loitering in LEO while it's being filled up? A reusable unmanned depot, or your crewed MCT? I think the preference would be to not have MCT floating around up there any longer than necessary. Even if you sent the crew up later once it was tanked up, it's just that much more time in LEO to get struck by MMOD.
And above the weather for TMI burn.Except, what do you want loitering in LEO while it's being filled up? A reusable unmanned depot, or your crewed MCT? I think the preference would be to not have MCT floating around up there any longer than necessary. Even if you sent the crew up later once it was tanked up, it's just that much more time in LEO to get struck by MMOD.
There will have to be plenty of loiter time in LEO for MCTs. Why? Because of launch windows to Mars. You will likely want to launch a fleet of them in very close succession, and this will require lots of loitering to place, refuel, & prepare the MCT's in orbit. Having a couple of weeks system checkouts in the relative safety of LEO is also advantageous.
No. If you are truly scaling this to be able to deliver 100t of cargo, you really need to have it in the middle. Not up front with the hab volume. When you do atmospheric entry, the propellant tanks will be mostly empty, so then by placing the cargo up top you are now forcing yourself to have to have a substantial minimum cargo load or the thing won't fly right.
Think about it. 100t. And it could be there, or it could be empty. If you do a sideways re-entry, that DOES constrain you to a center placement of cargo. OR you need to split the cargo into two balanced areas, one below and one above propellant tanks.
A SEP transit vehicle that takes you to high earth orbit followed by a perigee burn near Earth to send you to Mars, the SEP vehicle flies independently to mars and you rendevoue with it on low mars orbit to return to a high Earth orbit where crew disembark on a Dragon capsule. The bi-conic just dose mars assent with a 25 mT habitat inside pluss a modest landing propellent reserve. The intent is to ultimatly be able to do a rapid cycle between mars surface and low orbit, loading cargo and orbit and unloading on the surface.
I'm rather surprised that you aren't aware of the Pendulum Fallacy (https://en.wikipedia.org/wiki/Pendulum_rocket_fallacy)But I don't know that my stated bit about inherent stability is incorrect. Just not exactly as I stated.
In that, would it not be similar to a helicopter? Why is a helicopter more stable
than something like the LLRV?
And in this way I think they'd act more like the F9 core grid fins, which are placed up high to provide force (via air resistance rather than a jet of thrust) to help stabilize the core for landing.
The lifeboat would be [....] A flight deck in the nose with a bulkhead and hatch between it and the rest of the hab. Not much more to it that that. Not quite like a separate spacecraft.
For early exploration missions Dragon 2 would be sufficient, no separate system needed as crews will only be maybe 6-7.
Others thought it needed it, as later during colonization they'd want to launch all 100 people on the MCT rather than have a separate ferry.
I argued against an LAS, but was out voted, heh
10 or 100 people per flight.
10 people means 10 times the number of manned flights. Which means 10 times the likelihood of a flight with crew loss. Will people say, oh well, it is only 10 people, not 100, that's OK? I doubt it.
10 or 100 people per flight.
10 people means 10 times the number of manned flights. Which means 10 times the likelihood of a flight with crew loss. Will people say, oh well, it is only 10 people, not 100, that's OK? I doubt it.
However, with extra crew ships on the same route, if there is a failure of a major system on one of the ships, those 10 people can be spread amongst the remaining 9 ships. Additionally, during the emergency, they'd have nearby external help operating out of safe, fully functional ships, rather than trying to save themselves from within the failing ship using failing systems (with the nearest advice operating behind several minutes comms lag.)
With a single crew, no chance of rescue, any major failure means LOM/LOC.
The advantage of that 9-fold backup, IMO, is worth its weight in diamonds.
I assume you mean that the early crews would ferry up to an MCT in LEO using D2? (Since D2 can't launch back off the surface of Mars. It will never be used for human missions to Mars.)
In which case, you are proposing an entirely different kind of MCT just for the first few missions. Your escape vehicle can't be retro-fitted to an existing MCT design. You can't just cut through a few joins connecting the flight-deck to the rest of the MCT and add some pyro-bolts. You have the design the entire MCT around the separation mechanism. That's not going to be an afterthought or upgrade.
You are still not getting what I'm trying to say. Let me try again.
The MCT must be able to launch, fly, and land with full cargo/crew. But ALSO when no cargo/crew is present.
The MCT must be able to launch, fly, and land with full propellant load. But ALSO with tanks nearly empty.
That places severe constraints on the placement of these elements on a biconic entry vehicle, and you can't just hand-wave that away by a "100t payload vs 100t cargo" semantic discussion. It doesn't matter.
I'm not sure I agree with that.
Ok, so let's say you have a 12.5m wide IBMCT (our working diameter for it and the booster). A separable nose would be 12.5m wide at the base, tapering down. So you'd have quite a lot of volume there for a small crew returning from Mars if they had to abort and land down range. For 100 people, yea, it's not going to keep them for very long, but for a crew of 5-7? Should be just fine. In fact, that may be the only hab space they need/have for exploration missions, with everything below for surface cargo. That's a volume twice as wide as Skylab op the base, and probably about the same height as the Skylab pressurized volume. There's no squeezing involved.
They would have provisions and supplies sufficient for the 4-6 month transit back to Earth, so they should be ok for quite awhile.
Then you have the question of what sort of contingency plan do you want to have in place to deal with them at that point. That's really a separate discussion. Maybe a remote operated large pressurized rover that could drive itself over to the lifeboat, to give them transportation to a supply cache somewhere pre-positioned for such a contingency?
In a situation where they are transporting 100 people to Mars, obviously 100 people won't be coming back home, so there will be far fewer people on it. Probably just some SpaceX employees or NASA personnel returning home after a tour of serving at the colony, and a few people who have either become more ill than can be treated on Mars, or have changed their minds and want to go home. But even if it were more people, with a colony on Mars, rescue could be dispatched anywhere on the globe, it's just a matter of how long it would take to get there, so the lifeboat would need to be set up to support X number of people of Y length of time needed to get rescue there.
As far as whole vehicle abort goes, there really is no such thing for Mars ascent. If the MCT MPS explodes, a separable lifeboat can save the crew. Whole vehicle abort would only work for Earth ascent.
If there's a non explosive failure, like an engine out, that's when having a redundant engine comes in.
So it's really all or nothing if you don't have a separable design. And that's ok, the LAS lifeboat is mainly for Earth ascent so the crew can get away from an exploding booster where you cannot abort the whole fueled stage. But with that comes the ability to abort on Mars if necessary. But there would need to be contingency plans to for the marooned crew obviously.In Early exploration any abort to surface is fatal because their is zero rescue infrastructure on the surface and no conceivable capsule could carry sufficient supplies to see them through more then a few days. So I consider this a pointless abort. By the time you have an infrastructure to do surface rescue your passenger count is much too high for the small capsule your proposing, it would need to be a large vehicle comprising a significant portion of the whole mass of the vehicle and would present great difficulty in landing as you going to be falling on a ballistic trajectory from a high altitude and need massive retro-propulsion to not impact the surface, it in no way resembles the kind of un-powered capsule landing that can be done on Earth.
Not necessarily. It would only be a smaller portion of the whole weight of the vehicle (Maybe 1/3 total dry mass or so?...but more importantly is it leaves all the propellant mass behind with just the LAS/landing propellant on board). It would leave behind the main tanks, engines, most of the TPS covering, etc. It wouldn't be insignificant, but it would be certainly less than the whole vehicle. The LAS engines and tanks would need to be sized not only for abort, but for propulsive landing.
Also it would be a biconic shape. So it can do a biconic EDL rather than ballistic. It would still need a large retro propulsion as any vehicle would, but again, that would have to be designed into the LAS system if you wanted it.
Again, this a concept in reaction to some who feel strongly the LAS is necessary. (I argued against an LAS, but was out voted, heh) It could very easily be left off, with no Mars ascent abort option (for an explosive event, you can abort to orbit with just an engine out), and a separate LEO-taxi with LAS for Earth ascent. For those that favor that, I think that's viable too.
It would probably still need landing thrusters of some sort to land on Earth, which would likely be pressure fed for fast reaction control and reliability. Otherwise a means of landing on Earth with a vacuum Raptor nozzle would need to be figured out. Something like a retractable nozzle extension, or a jettisonable nozzle extension, so that the Raptor thrust isn't too over expanded for sea level. I'm not an engine expert, but have been told by several that vacuum engines with large vacuum nozzles like M1D-Vac and RL-10B cannot operate at sea level due to their large high efficiency nozzles.
But Raptor would still have to be capable of quickly responsive throttle in order to be able to land, which it may not be being a big pump fed staged combustion main propulsion engine. Otherwise you are back to landing thrusters.
[separate launch "taxi"]
That means you have two spacecraft to design from the start rather than one, but that's certainly an option.
First, it won't be just a few missions. Probably all the missions for the first couple decades. There will be many exploration missions with just a small crew before they can possibly think about actual colonization with large numbers of colonists. You have to explore various potential location looking for a promising site with favorable conditions for a colony. Then you have to test out the new systems which the colony will use, and test out resource collection, etc. Not to mention I think it highly likely NASA would jump in bed with them as soon as it were to look likley they could land people on Mars. They'll provide fund which will be beneficial for SpaceX, but they'll have their own agenda of places they want to go too.
And you can only fly out every 2 years. So, after maybe 20 years of Mars mission, you are ready to put 100 colonists on an MCT, what do you do then?
That's the question.
Why wouldn't a Mars escape trajectory arc over some? Do you wait until Mars rotates exactly towards Earth and then launch? I'm being a little facetious, but you get the idea, you need to turn anyway, so why not reduce gravity losses while you are doing it.
Hans Koenigsmann .... "but at least 100 times is our goal."Should have asked earlier - assume it's BFR here due to thread. Is that right? (MCT can only launch every 18 months, so 100 uses is over 50 years)
http://www.iflscience.com/space/huge-spacex-announcement-coming-soon-could-be-mars-mission
Ok Chris, you knew we were going to get ahold of this eventually. ;D
First question: Bigger than a Breadbox?
Whoever said they would only make one MCT?Hans Koenigsmann .... "but at least 100 times is our goal."Should have asked earlier - assume it's BFR here due to thread. Is that right? (MCT can only launch every 18 months, so 100 uses is over 50 years)
Hans Koenigsmann .... "but at least 100 times is our goal."Should have asked earlier - assume it's BFR here due to thread. Is that right? (MCT can only launch every 18 months, so 100 uses is over 50 years)
Hans Koenigsmann .... "but at least 100 times is our goal."Should have asked earlier - assume it's BFR here due to thread. Is that right? (MCT can only launch every 18 months, so 100 uses is over 50 years)
One MCT launch to Mars requires a minimum of 3-4 launches. Each synod needs at least 2 MCT to maintain a station, more to expand it, say at least 3. That's 9-12 launches every synod of 26 months. That's a bare minimum. I am quite sure there will be more.
Hans Koenigsmann .... "but at least 100 times is our goal."Should have asked earlier - assume it's BFR here due to thread. Is that right? (MCT can only launch every 18 months, so 100 uses is over 50 years)
One MCT launch to Mars requires a minimum of 3-4 launches. Each synod needs at least 2 MCT to maintain a station, more to expand it, say at least 3. That's 9-12 launches every synod of 26 months. That's a bare minimum. I am quite sure there will be more.
Not that it matters I guess, but ... my question doesn't matter. It MUST be the BFR.
(Your answer generalises it to the Mars Colonial Transport system. Which means BFR and tankers etc)
If the goal is for the MCT component that goes to Mars to be reused 100 times, and that component launches once every synod, but Koenigsmann doesn't think it'll last 30 years, then obviously he's talking about something else. If 2 or 10 MCTs launch that's not the same as a single MCT being reused 2 or 10 times.... but it would reuse the BFR.
(I'll have a listen to the speech and see if there was something else)
My take on the statement was that he meant it will not fly daily for 30 years which would be over 10,000 uses.Ahh, so in a sense he was recontextualising reuse, almost to say that the MCT's version of reuse is still much less than a plane.
It is in german and there was not much else in it for us space fans. It was an interview in a general news magazine.
I think Mr Musk was quite clear that he would do all possible for the MCT to return to Earth during the same synod. This would enable the same MCT to be serviced and reused during the next synod. Possible or not that is the goal.
You are still not getting what I'm trying to say.
It has to be a separate spacecraft, or it can't operate as a escape vehicle. Such a double-vehicle would become hideously complex.
[separate launch "taxi"]
That means you have two spacecraft to design from the start rather than one, but that's certainly an option.
But you propose two spacecraft in one. The LAS you propose is a separate spacecraft, according to you capable of full EDL. But it must also be integrated back into the MCT in a way that doesn't just allow it to break away as a LAS, doesn't just require it to be integrated tightly enough to handle re-entry while acting as the nose of the MCT (while still being able to break-away instantly), but its engines must also be able to serve as the landing engines for the whole MCT, both on Mars and on Earth.
(Plus you want another non-LAS "explorer" MCT for smaller crews. Plus yet another MCT design for cargo. So four spacecraft.)
I can't fathom how you think that is going to be somehow easier to nest two spacecraft inside each other like matryoshka dolls than to build them independently.
To use an analogy, do you think is would be cheaper/easier to design a car whose drivers-side wheels, seat and side panels split off to become a motorbike, or to just design a normal car and a separate normal motorbike?
You are assuming the major risk is in the cruise phase. I assume it is during launch, TMI and landing
Particularly if you are trying the use the escape vehicle's engines as landing engines for the entire MCT. The forces on the connectors, which must be instantly separable during launch-abort, would be ridiculous.
It has to be a separate spacecraft, or it can't operate as a escape vehicle. Such a double-vehicle would become hideously complex.
Yes. That's probably why such a cylinder-cone vehicle with a separable nose cone with a crew hab has never been thought of before.
;-)
It has to be a separate spacecraft, or it can't operate as a escape vehicle. Such a double-vehicle would become hideously complex.
Yes. That's probably why such a cylinder-cone vehicle with a separable nose cone with a crew hab has never been thought of before.
;-)
Doesn't even need to be cylinder/cone:
[images of Apollo CM+SM and Dragon+trunk]
It's not really two spacecraft in one, like MCT+ S2 would be two completely separate spacecraft. It's more a spacecraft, with an additional section. Since everything below the lifeboat cannot function on it's own as it own spacecraft. So unlike your analogy, it's more like a semi tractor and trailer.
But the point being, to say the the pieces being joined where they can separate in an emergency is "hideously complex" may be a tad over stated.
I think the Separate Bi-conic is attractive in this regard because rocket stages are cheaper to develop per unit of dry weight then manned capsules
But the point being, to say the the pieces being joined where they can separate in an emergency is "hideously complex" may be a tad over stated.
Do you not understand that a vessel that must function independently and as a deeply integrated functional part of a larger vessel (your lifeboat) is going to be much more complex than a vessel that only has to function independently (a LEO taxi)? And that a larger vessel that must be designed around a major piece that separates is going to be much more complex than a similar sized vessel that doesn't come apart?
11. it needs to survive on Mars for about 550 days - longer if it does not return the next synod
My point is that MCT is already horrendously difficult to design, adding a lifeboat makes it far more difficult.
Those first ~3 MCT will serve as a monument for many generations of martians.
Quote11. it needs to survive on Mars for about 550 days - longer if it does not return the next synod
They can waive that requirement if they have to. I always anticipated that the first few MCT which have long surface times on Mars may never return. That would include the first passenger MCT that serves as habitat for at least one synod on Mars.
Those first ~3 MCT will serve as a monument for many generations of martians. :)
Later MCT will be unloaded, refuelled, checked then fly back to earth after a short stay. They will land on better prepared landing sites too.
The really hard requirements are 5) 7) 8) 9) 13) 14) But that is already plenty I agree.
Waive a requirement to last for 550 days on Mars, replace with requirement to last indefinitely ?!?!
And the crew needs to get back somehow - first few missions won't have colonists.
Waive a requirement to last for 550 days on Mars, replace with requirement to last indefinitely ?!?!
And the crew needs to get back somehow - first few missions won't have colonists.
Waive the requirement to fly back after that time. The ISS shows that a habitat can be maintained for decades. That would be more valid for a habitat that can at some point be connected to an external ECLSS specifically designed for Mars and an expanding local power supply.
Part of the crew will go back with the passenger MCT that brings the crew that will man the station for the next synod and flies back after a few weeks. There will likely not be a life boat MCT. What would it be good for? It cannot go before the return window opens and that will be when the next MCT has arrived.
The Short-Stay Mission - often referred to as an opposition-class mission, this mission profile provides Mars stay times of 30 to 90 days with a round trip total time of 400 to 650 days. This mission class requires a large amount of energy to be expended in transit, even after taking advantage of either a Venus swingby (on either the inbound or outbound leg) or a deep space propulsive maneuver in order to limit Mars and Earth entry speeds.
So it is still a requirement to fly back one synod later. Not much of a gain there, at the cost of adding extra requirements about long duration surface stays and connection to external ECLSS.
So it is still a requirement to fly back one synod later. Not much of a gain there, at the cost of adding extra requirements about long duration surface stays and connection to external ECLSS.
Sounds almost like you are deliberately misunderstanding me.
With you having been given advanced sight of the SpaceX MCT architecture, every word you post on this "speculation" thread from now onwards will be scrutinised with a fine tooth coombe until the grand reveal ! not that I'm suggesting in anyway that you should refrain from posting on this thread until the reveal, PLEASE keep on posting :)
As far as we know all MCT, both cargo and passenger, will go on trajectories that allow them to return after a short stay on Mars and return in the same synod. At least that is what Elon Musk has set as a goal for his transport system to reuse them every synod instead of every second synod.
That means the return leg will be significantly longer than the leg earth-mars. You are introducing a new requirement for shorter return flights that would mean that manned MCT could not be reused every synod. That may or may not be the case. IMO it is just a reason to reduce the number of people who return to earth to a minimum, maybe have more water as shielding for at least a part of the crew space for the return leg. But that is problematic as the mass budget available for return is much smaller.
I think that's a good point about the difference between NASA Mars and SpaceX Mars. NASA Mars is very much focused on rotation. SpaceX Mars is to create a colony.
I doubt they would be short of customers who would be willing to spend some cash (probably affordable to a lot of people via selling their home) and up sticks and become a resident of Mars.....and not return (the element of increasing the population, as opposed to several years stays and coming back).
Reloading a returned MCT with cargo is an issue. If modular, cargo could be transferred by BFR and a tug.
Landing on Earth is not required for maintenance.
The MCTs could be staged somewhere in cis-lunar space instead of LEO to reduce outgoing and return delta V requirements. Makes it more difficult to stage, but could be worth the effort.
Reloading a returned MCT with cargo is an issue. If modular, cargo could be transferred by BFR and a tug.
Landing on Earth is not required for maintenance.
The MCTs could be staged somewhere in cis-lunar space instead of LEO to reduce outgoing and return delta V requirements. Makes it more difficult to stage, but could be worth the effort.
Why bring back the MCT to cis-lunar space or LEO if you do not land on Earth? Just keep it at Mars. Use it to shuttle cargo/personnel from Mars orbit to the surface.
Why doesn't SpaceX just build the MCT with a metholox plug nozzle engine and strap 8 Falcon 9's around it to launch it off earth? They already have the Falcon 9's.
But the point being, to say the the pieces being joined where they can separate in an emergency is "hideously complex" may be a tad over stated. ;-)
But the point being, to say the the pieces being joined where they can separate in an emergency is "hideously complex" may be a tad over stated. ;-)
(https://upload.wikimedia.org/wikipedia/en/thumb/1/17/CEV_Lockheed_Martin.jpg/220px-CEV_Lockheed_Martin.jpg)
I think that's a good point about the difference between NASA Mars and SpaceX Mars. NASA Mars is very much focused on rotation. SpaceX Mars is to create a colony.By the time people have payed off their home or have enough equity in their home to pay for the couple to move to Mars they would have already raised their kids. Not good to start a colony with people who have already raised their kids. More likely the first colonist will be sent there by sponsors.
I doubt they would be short of customers who would be willing to spend some cash (probably affordable to a lot of people via selling their home) and up sticks and become a resident of Mars.....and not return (the element of increasing the population, as opposed to several years stays and coming back).
By the time people have payed off their home or have enough equity in their home to pay for the couple to move to Mars they would have already raised their kids. Not good to start a colony with people who have already raised their kids. More likely the first colonist will be sent there by sponsors.
I think the Separate Bi-conic is attractive in this regard because rocket stages are cheaper to develop per unit of dry weight then manned capsules which are very nearly the most expensive things in aerospace development,
I estimate a bi-conic would have a mass of around 75 mT, and the 2nd stage would be 72 mT. I expect the Integrated version would have a mass greater then 75 but probably less the the raw additive 150 of the two separate vehicles due to some savings on redundancies, but it will probably have a development cost that is per pound equal to the smaller bi-conic I'm looking at. This will wipe out the advantage of not developing a 2nd stage and makes the total cost greater unless the integrated vehicle has an exceedingly low mass or the cost ratio between stages/capsules is extremely low perhaps due to the difficulty of 2nd stage reuse engineering.
I would really like to hear some exact mass number from Lobo about the whole vehicle stack, dry masses for the first stage, the integrated bi-conic dry mass (without abort systems if that's your preference) and propellant loads in each so I can plug them into the launch vehicle performance calculator I've been using http://www.silverbirdastronautics.com/LVperform.html and do an apples to apples comparison to see how gross take off weight differs.
But the point being, to say the the pieces being joined where they can separate in an emergency is "hideously complex" may be a tad over stated. ;-)
(https://upload.wikimedia.org/wikipedia/en/thumb/1/17/CEV_Lockheed_Martin.jpg/220px-CEV_Lockheed_Martin.jpg)
And your point with this image is...?
See Lars's post too. It may be impractical, but I'm not the only one to propose it. Heidmann's is more "CST-100" than "Dragon 2". Which was our first thought on it too. But we need a method to land on Earth which cannot be done one Vacuum Raptor as it sits. So the idea of moving them external and landing on Earth on them as double duty was considered, vs. another set of landing engines at the base.
And its not just a lifeboat. That would be one thing. In the system Lobo proposes, the "lifeboat" is a major section of the re-entry system for the MCT and the primary propulsion for landing.
A short return trip adds few extra requirements onto the MCT as there already are requirements for a short trip to Mars. However a long return trip which goes into the orbit of Venus adds requirements that are not needed on other phases of the mission. For cargo it adds a few extra requirements, but for crew it adds far more, especially as the MCT will be limited to 25% payload on the return trip.
....
I think Elon will be satisfied if the 90% cargo missions can be reflown in 1 synod while the 10% crew missions are reflown in 2 synods. Crew MCT probably need far more refurbishment than cargo MCT so it would be a push to get them reflown the next synod anyway.
An emplty MCT would have a greater delta v capability than one that is carring cargo or crew. This could reduce the travel time back to Earth and or have more return windows.A short return trip adds few extra requirements onto the MCT as there already are requirements for a short trip to Mars. However a long return trip which goes into the orbit of Venus adds requirements that are not needed on other phases of the mission. For cargo it adds a few extra requirements, but for crew it adds far more, especially as the MCT will be limited to 25% payload on the return trip.
....
I think Elon will be satisfied if the 90% cargo missions can be reflown in 1 synod while the 10% crew missions are reflown in 2 synods. Crew MCT probably need far more refurbishment than cargo MCT so it would be a push to get them reflown the next synod anyway.
If you can do a short trip to Mars with a longer return trip, to fit into 1 synod, I would assume you could do a long trip to Mars with a short return trip.
From an Earth perspective a human-return MCT would return from Mars just a few months after the main fleet leaves for Mars. The people would offload, and it would be refurbed and leave with cargo only (having missed the short trip window) for a 1 year trip to Mars.
From the Martian perspective you'd have that cargo MCT arriving a year after the passenger MCTs (which have been sent back with cargo on a slow return already). Offload the cargo, load with passengers, and launch back on a quick trajectory again.
It allows each MCT to be used once per synod - but I apologise in advance that I don't have a good enough grasp of the orbital mechanics involved. Is there any reason we hear more about a short trip to Mars, 1 month stay, with long return - and not the opposite long trip to Mars, 1 month stay, with short return?
As far as we know all MCT, both cargo and passenger, will go on trajectories that allow them to return after a short stay on Mars and return in the same synod. At least that is what Elon Musk has set as a goal for his transport system to reuse them every synod instead of every second synod.I do not believe we have shown whether this is possible or not, in terms of how much delta V is required for what overall mission duration & entry velocities.
That means the return leg will be significantly longer than the leg earth-mars. You are introducing a new requirement for shorter return flights that would mean that manned MCT could not be reused every synod. That may or may not be the case. IMO it is just a reason to reduce the number of people who return to earth to a minimum, maybe have more water as shielding for at least a part of the crew space for the return leg. But that is problematic as the mass budget available for return is much smaller.
Reloading a returned MCT with cargo is an issue. If modular, cargo could be transferred by BFR and a tug.
Landing on Earth is not required for maintenance.
The MCTs could be staged somewhere in cis-lunar space instead of LEO to reduce outgoing and return delta V requirements. Makes it more difficult to stage, but could be worth the effort.
Why bring back the MCT to cis-lunar space or LEO if you do not land on Earth? Just keep it at Mars. Use it to shuttle cargo/personnel from Mars orbit to the surface.
I'm assuming a reusable interplanetary vehicle would be useful. Also direct entry to Mars instead of going into Mars orbit. I believe that is what Elon has in mind. Of course, other options are possible.
I'm strongly in favor of SEP and see considerable use for it making VERY FAST transit possible.
How is this possible? Wouldn't the system need to be huge and have magical power sources and such (insert anything Zubrin has ever said). NO, you can get a Fast transit on the order of 100 days to Mars with a slow, low power SEP system.
The trick is you use your SEP to move your Mars bound vehicle with propellants up to high Earth orbit and then then drop by the Earth for a huge Oberth assisted burn. For 2 km/s you should leave Earth with huge escape velocity and reach mars in 100 days (average).
Now the problem is capturing at mars, the answer is Magneto-Plasma Aerocapture, this lets us avoid expensive propulsive capture and is then followed by about a week of Plasma assisted aerobraking which lets the eventual EDL be from a gentle 4 km/s. So we get to have both fast transit and easy low speed EDL.
The SEP system has not even left Earth yet in this scenario, so you can do either one of two things, bring it back down to LEO for refueling and do it again (basically making it a Cis-lunar tug), or send it to mars by the conventional slow method of spiraling out from the Earths SOI (the SEP is too delicate to take the high thrust of the Oberth maneuver). In the latter case your going to arrive much later then the manned capsule but if this is a conjunction mission the crew will be spending around 600 days on mars so their is plenty of time for the SEP to arrive before it is needed for departure which is what I favor.
The MCT would only need to reach low Mars orbit and would then rendezvous with the SEP and head for Earth, this return transit is made reasonably short by the fact the MCT is a completely dry shell now of only 100 mT (75 vehicle mass + 25 return cargo) and the SEP is nearly dry too so power to weight ratios are increased, also were not aiming to match Earth's orbit and capture gently, were going to simply intersect it on an elliptical orbit around the sun, that cuts the DeltaV needed. At Earth we used the Magneto to capture again and bring both SEP and the MCT down to LEO (they probably need to separate to do this as the SEP is more delicate and would slow the process down for the MCT). The crew can be retrieved via a Dragon capsule now, and we need to send another tanker to LEO to put landing propellants into the MCT, if we use enough the MCT can do a lot of retro-propulsion on entry and bring it's entry speed down from the 7.7 km/s of orbit down to the range of 4 km/s which matches it's mars entry speed, so all the thermal protection systems can be designed for this low performance point.
IMLEO is estimated at 570 mT of which 100 mT is the cargo load, 75 mT is the MCT dry mass, 200 mT is chemical propellant in the MCT (2 tanker loads of 100 mT each), 155 is SEP propellant, 15 mT is the SEP tank and 22 mT is the SEP hardware which has a power output of 4.5 MW which corresponds to an alpha value of 5 kg/kw.
BTW Using a braking system like Magneto Plasma is the only way I can see an Integrated Bi-conic and direct Earth return being viable, without it the entry conditions are too extreme to meet the low dry mass fractions that it's advocates are proposing.
That sounds far too ambitious. First of all, why VERY FAST transit? What's the point? Not worth the effort IMO, not in any near future.
You probably want some chemical propulsion on your SEP stage, at least for the crew transfer, but not anything on the scale of MCT.
Magneto-Plasma Aerocapture? Are you sure that's not only for aerobraking? Either way, probably far from ready.
Why take the Lander back to LEO if you do not need it for aerocapturing? If they can't "refurbish" it on Mars at the beginning, you might as well expend it.
4.5MW is huge and 5kg/kw is far below the numbers I've seen.
The SEPs can cycle between LEO/HEO and HEO/Mars, that way they are back a lot faster and you need less power.
375t is a freaking huge payload for SEP. You can divide that into cargo/hab/lander and get 100t+ pieces. Again, less power.
The trick is you use your SEP to move your Mars bound vehicle with propellants up to high Earth orbit and then then drop by the Earth for a huge Oberth assisted burn.
You probably want some chemical propulsion on your SEP stage, at least for the crew transfer, but not anything on the scale of MCT.
the thrust level of chemical propulsion would require the SEP to have far higher rigidity they it otherwise would and it's mass would essentially be parasitic when doing the chemical burn as we intend to leave Earth with all the escape velocity needed to reach mars.
The trick is you use your SEP to move your Mars bound vehicle with propellants up to high Earth orbit and then then drop by the Earth for a huge Oberth assisted burn.You probably want some chemical propulsion on your SEP stage, at least for the crew transfer, but not anything on the scale of MCT.the thrust level of chemical propulsion would require the SEP to have far higher rigidity they it otherwise would and it's mass would essentially be parasitic when doing the chemical burn as we intend to leave Earth with all the escape velocity needed to reach mars.
Doing a drop'n'go Oberth manoeuvre means that your SEP is a high thrust type. (You can't spend days doing a HEO-LEO-MTO Oberth burn.)
That implies that your SEP-ship is already quite rigid. (Doing aerobraking into Mars orbit also implies that your ship is pretty rigid.)
Yep. SEP is a non-starter for a Mars transit. At a minimum it needs to be augmented by significant chemical propulsion for most of the effective delta-V. So why bother?
And if you treat the ship as essentially a cycler (in high orbits on both ends), then you are dealing with the same problem that cycler-advocates keep forgetting... In order to reach your SEP "battlestar galactica", you have to use chemical propulsion to accelerate all the crew, supplies, cargo, AND lander propellant to that high (almost escape) orbit for rendezvous. This last element is what SEP and cycler advocates seem to always forget about.
No you misunderstand, the SEP pushes the lander up to HEO first, then the Lander and SEP SEPARATE at HEO, only the lander with it's chemical propulsion system and inherent ability to tolerate high acceleration goes through the Oberth maneuver.
The SEP spirals away from the Earth on low thrust without any Oberth benefits, but it isn't carrying ANY cargo and isn't needed at mars for nearly 2 years so it can travel slowly and at high efficiency.
No one who favors SEP for heliocentric HEO->Mars transfer would neglect using it for going between LEO->HEO
No one who favors SEP for heliocentric HEO->Mars transfer would neglect using it for going between LEO->HEO
They might. The radiation during a slow LEO-HEO spiral is apparently hell on solar panels. That means your efficiency is already degraded by the time you do the long Mars flight.
Musk has had time of flight goals that are very aggressive as he seems to want to basically outrun the radiation danger and have a transit time that most people would find acceptable when crammed in small spaces.
I don't believe it is necessary to make a hybrid drive SEP stage, the thrust level of chemical propulsion would require the SEP to have far higher rigidity they it otherwise would...
It is low TRL now but it is in active development, a demonstration cube sat will be launched soon by NASA.
And it is needed here to carry the return habitat from mars surface to orbit and it carries the primary Magneto system to brake with upon return to Earth. The SEP would have a brake too but much smaller as it's more delicate and doesn't have the urgency because it doesn't carry the crew.
It's a conservative value consistent with Hall thrusters with power conversion units and solar at 300 W/kg. Perhaps you misread it as kw/kg, it is the other way.
The potential does exist to use solar
power for exploration within the inner solar
system. Solar array technology continues
to improve in cell efficiencies and in
system alpha. The expanded use of electric
prolusion, and the synergistic benefit of
electric propulsion and increasing power
availability is driving commercial and
government towards higher power systems.
The SOA solar arrays are currently 15 - 20
kg/kW. The DARPA Fast Access
Spacecraft Testbed (FAST) program has
progressed the SOA has with near-term
projected goals of 8 kg/kW at 1 AU.19
NASA studies have far-term predictions
approaching 3.3 kg/kW including the array,
based on Stretched Lens Array Square
Rigger (SLASR) technology with advanced
cells, gimbals, booms, power cabling, etc.20
A point design for a 232kW system had a current best estimate (CBE) system mass of 781 kg.21 Using an advanced
cell, a 100 kW End-of-Life, after 10 years at GEO, stretched lens array design has a predicted mass performance of
1.85 kg/kW
Yep. SEP is a non-starter for a Mars transit. At a minimum it needs to be augmented by significant chemical propulsion for most of the effective delta-V. So why bother?
Yep. SEP is a non-starter for a Mars transit. At a minimum it needs to be augmented by significant chemical propulsion for most of the effective delta-V. So why bother?..."Non-starter"?? That clearly doesn't mean what you think it means. SEP is baselined for Mars transit in NASA's current Mars exploration plan, and SpaceX has mentioned it's being traded for MCT.
Yep. SEP is a non-starter for a Mars transit. At a minimum it needs to be augmented by significant chemical propulsion for most of the effective delta-V. So why bother?..."Non-starter"?? That clearly doesn't mean what you think it means. SEP is baselined for Mars transit in NASA's current Mars exploration plan, and SpaceX has mentioned it's being traded for MCT.
That makes SEP pretty much the opposite of "non-starter."
With SEP you fall back into just another "Battlestar Galactica" mission mold (massive mother ships).
In order to reach your SEP "battlestar galactica", you have to use chemical propulsion to accelerate all the crew, supplies, cargo, AND lander propellant to that high (almost escape) orbit for rendezvous. This last element is what SEP and cycler advocates seem to always forget about.
No most of the effective acceleration is still coming from the SEP, to simply have the Lander perform depart form LEO with this much Escape velocity would require nearly 5.8 km/s DeltaV which would mean an IMLEO of 1000 mT for the same TMI mass.
A SEP transit vehicle is not at all like a cycler, it is making a controlled propulsive orbital insertion or braking maneuver at each planet, that means your not blasting off at escape velocity to catch the thing as it flies by on a hyperbolic trajectory each time.
No one who favors SEP for heliocentric HEO->Mars transfer would neglect using it for going between LEO->HEO as each of these is about equal in DeltaV for a low thrust system. Only the crew in a small capsule needs to be sent via chemical propulsion for radiation avoidance, but all cargo, all consumables, all propellants and the vehicle hardware would obviously be a positioned in HEO by SEP leaving <1% of the mass to be moved by Chem.
With SEP you fall back into just another "Battlestar Galactica" mission mold (massive mother ships).
Which is evidently wrong, because SEP vehicles tend to be a lot more lightweight. What makes them appear massive are the big solar arrays.
HEO (or even worse, a highly elliptical LEO-HEO transfer orbit) is a terrible assembly/rendezvous point, for a variety of reasons. Some that immediately come to mind:
- Time to get there for SEP cargo runs - you need a LOT of SEP transfer vehicles to preposition cargo for a flotilla of your Mars transfer vehicles.
- The number of runs you need to take through the Van Allen belts with a crew is significant, even if you dock and depart ASAP. It's not great for all the cargo spirals either.
So Im curious, what exactly do you mean by "HEO"?
Again what you propose does not have much left in common with what we know of the MCT architecture, so the discussion should probably move.
I'm strongly in favor of SEP and see considerable use for it making VERY FAST transit possible.Do you have math to back this mission plan up?
How is this possible? Wouldn't the system need to be huge and have magical power sources and such (insert anything Zubrin has ever said). NO, you can get a Fast transit on the order of 100 days to Mars with a slow, low power SEP system.
The trick is you use your SEP to move your Mars bound vehicle with propellants up to high Earth orbit and then then drop by the Earth for a huge Oberth assisted burn. For 2 km/s you should leave Earth with huge escape velocity and reach mars in 100 days (average).
Now the problem is capturing at mars, the answer is Magneto-Plasma Aerocapture, this lets us avoid expensive propulsive capture and is then followed by about a week of Plasma assisted aerobraking which lets the eventual EDL be from a gentle 4 km/s. So we get to have both fast transit and easy low speed EDL.
The SEP system has not even left Earth yet in this scenario, so you can do either one of two things, bring it back down to LEO for refueling and do it again (basically making it a Cis-lunar tug), or send it to mars by the conventional slow method of spiraling out from the Earths SOI (the SEP is too delicate to take the high thrust of the Oberth maneuver). In the latter case your going to arrive much later then the manned capsule but if this is a conjunction mission the crew will be spending around 600 days on mars so their is plenty of time for the SEP to arrive before it is needed for departure which is what I favor.
The MCT would only need to reach low Mars orbit and would then rendezvous with the SEP and head for Earth, this return transit is made reasonably short by the fact the MCT is a completely dry shell now of only 100 mT (75 vehicle mass + 25 return cargo) and the SEP is nearly dry too so power to weight ratios are increased, also were not aiming to match Earth's orbit and capture gently, were going to simply intersect it on an elliptical orbit around the sun, that cuts the DeltaV needed. At Earth we used the Magneto to capture again and bring both SEP and the MCT down to LEO (they probably need to separate to do this as the SEP is more delicate and would slow the process down for the MCT). The crew can be retrieved via a Dragon capsule now, and we need to send another tanker to LEO to put landing propellants into the MCT, if we use enough the MCT can do a lot of retro-propulsion on entry and bring it's entry speed down from the 7.7 km/s of orbit down to the range of 4 km/s which matches it's mars entry speed, so all the thermal protection systems can be designed for this low performance point.
IMLEO is estimated at 570 mT of which 100 mT is the cargo load, 75 mT is the MCT dry mass, 200 mT is chemical propellant in the MCT (2 tanker loads of 100 mT each), 155 is SEP propellant, 15 mT is the SEP tank and 22 mT is the SEP hardware which has a power output of 4.5 MW which corresponds to an alpha value of 5 kg/kw.
BTW Using a braking system like Magneto Plasma is the only way I can see an Integrated Bi-conic and direct Earth return being viable, without it the entry conditions are too extreme to meet the low dry mass fractions that it's advocates are proposing.
HEO (or even worse, a highly elliptical LEO-HEO transfer orbit) is a terrible assembly/rendezvous point, for a variety of reasons. Some that immediately come to mind:
- Time to get there for SEP cargo runs - you need a LOT of SEP transfer vehicles to preposition cargo for a flotilla of your Mars transfer vehicles.
- The number of runs you need to take through the Van Allen belts with a crew is significant, even if you dock and depart ASAP. It's not great for all the cargo spirals either.
So Im curious, what exactly do you mean by "HEO"?
Again what you propose does not have much left in common with what we know of the MCT architecture, so the discussion should probably move.
Our SEP tug is reusable and we expect to be launching continually from Earth to build up a High Earth Orbit, aka something near escape velocity, could be Lagrange point, Lunar Distant Retrograde, highly elliptical orbits etc etc, basically something out at the very edge of Earth's Sphere of influence, the DeltaV is similar for them all.
- High Earth orbit is FAR better then LEO in every way, it is free of dangerous orbital junk, it is colder meaning propellant boil-off is minimized, the SEP won't be shaded as it would be in LEO meaning it can depart at full power, their is no appreciable drag which would require constantly re-boosting stuff.
- Time to conduct the raising of cargo is inconsequential because we expect a continual non-stop launch campaign on Earth putting cargo into a stream going to the assembly point, the SEP tugs and their propellants will be less massive then the equivalent propellants needed to move cargo to HEO and do TMI from their then doing it from LEO on chem even at the slowest transfer.
- Why do I have to keep shooting down this straw-man, I have told you on several occasions that crew are sent up to the assembly point by a capsule on chemical propulsion, they do not make multiple passes through the Belt and I have never heard any evidence that the belt will be harmful to cargo. Amorphous silicon solar collectors are basically immune to radiation degradation and can make plenty of passes through the belt as well.
Your INTERPRETATION of Musk's comments is different then my own, your interpretation is common, even dominant I grant that but it basically amounts to a single stage from LEO being then going all the way to mars surface and then back to Earth surface with refueling at LEO and mars surface. I've rejected this as physically impossible as it makes a SSTO vehicle on Earth look easy in comparison. I interpret his landing comment to refer to just the lander, aka the lander dose not shed any parts during EDL, a transit vehicle in mars orbit will be necessary to return to Earth. I interpret any ambiguous comments from Musk in the most conservative manor possible and expect that system will be more complex then even Musk's original plans call for as he aims for the simplest solution first and has to compromise once he gets flight experience.
Magnetic capture/deceleration may be included in the plan sooner than we think. It sounds like it is real and has very major advantages. We can assume though that it is not part of the plan now. There is much to do, lets wait it out.
In every way? ::) Oh boy. Not for accessibility, that's for sure. Nor travel time for supplies. Nor for solar flare protection. There are good some valid trade-offs to be made by staging there, but pretending to that there are no down sides just makes your argument look silly and it taints the rest of your arguments.
Inconsequential? ;D As said by everyone when real world practicality gets in the way. But I suppose being so generic allows you to hand-wave away all concerns. And the number of SEP tugs that would be needed. You haven't even settled on a HEO staging point, so how can we challenge your math assumptions for a SEP tug and how many you will need, and how large such SEP tugs would be?
I only mentioned it to make you define your HEO staging orbit. And since you have not - above you wrote that "highly elliptical orbits" were a possible staging place, and those could certainly intersect with Van Allen belts. (again, specificity about a staging orbit would help avoid us bringing up these "straw-men" arguments) As for the effect on computers, arrays, and other equipment, I'm not sure what the point would be to bring it up to you, since you have already hand-waved away such concerns.
Well, now that *you* have rejected it... ;) So you admit that you are basically using this thread now as a soap box for your own Mars architecture, now that you have rejected the aspects of it that are unique? And are there any more exotic (theoretical with no actual work done on them) technologies like Magneto Plasma Aerocapture that you want to throw in? NEP?
Whatever shape SpaceX's Mars plans actually solidify into will likely follow their current paradigm. Better, more practical use of existing technologies instead of chasing the very cutting edge of performance. It will be something that is more affordable rather than exotic. Taking something like a Mars SSTO lander and realizing that with some added performance and propellant transfer it can be used for Mars transit. And it could boost itself into LEO. Who would have thunk? Your ideas are almost the opposite of the SpaceX approach. Multiple staging points. Multiple kinds of vehicles. Are you *trying* to make this as expensive as possible by mutating it into a NASA-ish Mars program?
SEP tugs themselves could be refueled in LEO.
A refueling assembly station at L1. This would be a fuel depot and cargo storage facility.
SEP tugs would refuel and pick up cargo and fuel and transfer this to L1 on a continuous basis.
MCT would launch, refuel at a LEO station, go to L1, refuel and load cargo containers, then go to Mars.
Yep. SEP is a non-starter for a Mars transit. At a minimum it needs to be augmented by significant chemical propulsion for most of the effective delta-V. So why bother?I think you're misunderstanding the concept, which seems sound to me, other than the notion that a technology as immature as MAC could be in the critical path of the plan without SpaceX having already purchased MSNW. I think you guys are talking past each other because you may not understand the concept.
And if you treat the ship as essentially a cycler (in high orbits on both ends), then you are dealing with the same problem that cycler-advocates keep forgetting... In order to reach your SEP "battlestar galactica", you have to use chemical propulsion to accelerate all the crew, supplies, cargo, AND lander propellant to that high (almost escape) orbit for rendezvous. This last element is what SEP and cycler advocates seem to always forget about.
Yep. SEP is a non-starter for a Mars transit. At a minimum it needs to be augmented by significant chemical propulsion for most of the effective delta-V. So why bother?I think you're misunderstanding the concept, which seems sound to me, other than the notion that a technology as immature as MAC could be in the critical path of the plan without SpaceX having already purchased MSNW. I think you guys are talking past each other because you may not understand the concept.
And if you treat the ship as essentially a cycler (in high orbits on both ends), then you are dealing with the same problem that cycler-advocates keep forgetting... In order to reach your SEP "battlestar galactica", you have to use chemical propulsion to accelerate all the crew, supplies, cargo, AND lander propellant to that high (almost escape) orbit for rendezvous. This last element is what SEP and cycler advocates seem to always forget about.
I understand why the crew will require chemical propulsion to reach high orbit rendezvous, but what about everything else? Why can't they use SEP tugs to climb in and out of the gravity well?
What's your alternative that you are implying exists?
That's quite fascinating... Because from my point of view, "DeltaV-only thinking" seems to be what *you* are doing, making you turn to SEP, interesting trajectories, and Magneto Plasma Aerocapture. Again, pursuing performance at any cost is not what SpaceX is known for doing. They prefer a more brute force but cost effective approach.
Yep. SEP is a non-starter for a Mars transit. At a minimum it needs to be augmented by significant chemical propulsion for most of the effective delta-V. So why bother?..."Non-starter"?? That clearly doesn't mean what you think it means. SEP is baselined for Mars transit in NASA's current Mars exploration plan, and SpaceX has mentioned it's being traded for MCT.
That makes SEP pretty much the opposite of "non-starter."
Yep. SEP is a non-starter for a Mars transit. At a minimum it needs to be augmented by significant chemical propulsion for most of the effective delta-V. So why bother?..."Non-starter"?? That clearly doesn't mean what you think it means. SEP is baselined for Mars transit in NASA's current Mars exploration plan, and SpaceX has mentioned it's being traded for MCT.
That makes SEP pretty much the opposite of "non-starter."
I don't know if "non-starter" is the right term, but the numbers don't look very good for SEP on a 100 day trip.
Why did NASA baseline it? I didn't read said plan (I try to stick to non-fiction) but NASA makes a lot technological decisions that are not ideal.
Why is NASA still sticking with H2 and solids? Why parachutes?
NASA is a) not a singular body but rather made up of many "selfish" bodies, b) is influence by many irrelevant and political factors, and ) sometimes makes honest to goodness mistakes.
Yep. SEP is a non-starter for a Mars transit. At a minimum it needs to be augmented by significant chemical propulsion for most of the effective delta-V. So why bother?..."Non-starter"?? That clearly doesn't mean what you think it means. SEP is baselined for Mars transit in NASA's current Mars exploration plan, and SpaceX has mentioned it's being traded for MCT.
That makes SEP pretty much the opposite of "non-starter."
I don't know if "non-starter" is the right term, but the numbers don't look very good for SEP on a 100 day trip.
Why did NASA baseline it? I didn't read said plan (I try to stick to non-fiction) but NASA makes a lot technological decisions that are not ideal.
Why is NASA still sticking with H2 and solids? Why parachutes?
NASA is a) not a singular body but rather made up of many "selfish" bodies, b) is influence by many irrelevant and political factors, and ) sometimes makes honest to goodness mistakes.
NASA isn't looking at 100 day transits much. SpaceX is. And SEP does reduce IMLEO dramatically for a typical 6-7 month trajectory that NASA likes to look at. And NASA is planning surface rendezvous, so most flights to Mars will carry cargo (which doesn't benefit much from fast transit).
SEP actually trades quite well, and technology improvements allow Isp improvements, not being limited in Isp like chemical and NTR by needing to keep your combustion chamber from melting.
SEP allows you to reduce transit time (versus an equivalent-IMLEO chemical-only system) if you have good enough solar array and thruster technology, but its chief benefit is usually a dramatic reduction in IMLEO. And SpaceX could leverage that just for hauling propellant from LEO to HEO if they so desired. Remember, SpaceX's LEO Constellation will use SEP tech, so SpaceX will be quite capable technically of building SEP tugs (and will be producing Megawatts' worth of SEP arrays/thrusters anyway just for the constellation).
I'll also add that in addition to the trip duration, there's the "land the whole thing" mentality. If you're talking about ISS-esque solar array and radiator array, the idea of stowing them before EDL is not practical.
If instead you have a "Hermes"-type orbit-to-orbit ship in mind, then now SEP starts making sense.
But again, this is about MCT.
With SEP you can basically get the first ~3 km/s of your LEO departure speed done outside of the vehicle actually being sent TMI which is achieving the remaining ~2 km/s vastly easier if your doing it via chemical propulsion because your propellant is now less massive then the payload.
We KNOW SpaceX is considering SEP for Mars, and as it is impossible to land a SEP, that mean IPSO FACTO that they are have ALSO considered a Semi-Direct architecture of a transit-vehicle and a separate landing vehicle. And that such a vehicle trades very well against a single massive direct vehicle.
No. If you are truly scaling this to be able to deliver 100t of cargo, you really need to have it in the middle. Not up front with the hab volume. When you do atmospheric entry, the propellant tanks will be mostly empty, so then by placing the cargo up top you are now forcing yourself to have to have a substantial minimum cargo load or the thing won't fly right.
Think about it. 100t. And it could be there, or it could be empty. If you do a sideways re-entry, that DOES constrain you to a center placement of cargo. OR you need to split the cargo into two balanced areas, one below and one above propellant tanks.
It's 100mt of payload. That may be 100mt of pure cargo down the road at some point, but for quite some time it will be mixed cargo and crew/hab.
So you have your IBMCT. It has it's fairly heavy MPS on the bottom. Above that you have two stacked cylindrical tanks that go up to about the center of the overall vehicle (when measured from tip to tail)...up a little over half way up the cylindrical portion. Above that you'll have a cargo deck for mixed flights that is maybe 3 meters tall. As the cargo will have to be lowered to the surface, none of it can be -too- large in one piece, but it will be heavy overall. Above that you will have a pressurized hab volume filling the rest of the cylinder, and then another tapered pressurized hab volume in the nose. If an LAS is required, the nose will have LAS/landing engines it it as well, along with small pressurized tanks to fuel them. Overall, it will have a fair amount of mass in it.
So again, you'll have a mass area in the tail, amidships, and in the nose.
With 100mt of pure cargo, and no hab area at all, that may be a little more tricky if all that 100mt is between the nose and the tanks. They'd probably stow the heaviest pieces just above the tanks, with lighter and lighter pieces above that to help with weight distribution.
But, it's obviously something that'd have to be looked at in more detail by actual SpaceX engineers during the actual design process to see how the real weight distribution will interact with the EDL profile.
This divided concept would have the problem of being nose-light. The mass will be concentrated in the middle, and at the aft in the MPS. But given it's tapered overall OML, maybe that still makes for a feasible distribution for EDL? It's a little above my area. :-)
With the SDMCT, obviously that's not a problem because it's always vertical.
To go back to balancing the Biconic for a bit, would it be possible to have a configuration similar to the Phoenix (see attached), except applied to a Biconic and with some sort of escape system in the nose (sorry, really like the idea of the nose section being a dedicated 'lifeboat'. I just see it as a way to solve so many problems) It would still have the central column, as a way to move down to the main pressurized area.
Not sure what this would do to the mass in total, but wouldn't it allow for a more evenly distributed mass during EDL?
One additional benefit to a tunnel through a fuel tank: It makes the best solar storm shelter you can buy on the mass required for the tunnel segment and some electric blankets.To go back to balancing the Biconic for a bit, would it be possible to have a configuration similar to the Phoenix (see attached), except applied to a Biconic and with some sort of escape system in the nose (sorry, really like the idea of the nose section being a dedicated 'lifeboat'. I just see it as a way to solve so many problems) It would still have the central column, as a way to move down to the main pressurized area.
Not sure what this would do to the mass in total, but wouldn't it allow for a more evenly distributed mass during EDL?
I do like that image, it is close to what I have in mind for an MCT. But I think a lengthy tunnel through a cryogenic tank might have issues.
If you insist on have a nose-mounted LAS, another way around it might be to have just a giant central tank (split between oxidizer and fuel), and then have a lower cargo bay (just above engines), and an upper cargo bay in the nose. It would force some level of cargo balance, but that would allow you to put all the pressurized volume up top. And as a bonus the unpressurized/heavy cargo would be easily offloaded near the surface.
Please list some new technologies/systems you actually support developing other then huge boosters.
I'm still shocked that anyone can consider HALL thrusters EXOTIC propulsion in this day and age, they have existed for decades, are on hundreds of com-sats. The bloody Ruskies have had them as baseline propulsion systems for their Mars mission plans for decades (first nuclear powered now solar powered), and they are kind of know for using 'brute force' technologies. Just because NASA and American Aerospace only 'discovered' them 10 years ago dosn't mean they are some finniky bleeding edge tech.
Small-thruster SEP is now routine on commsats for stationkeeping. Hall Effect Thrusters have recently solved the erosion problems that limited their lifetime. Large arrays of small thrusters are trivially easy. The design of large arrays of solar panels require only moderate scaleup efforts relative to the monetary outlays discussed in Mars mission planning; The only reason we don't have them yet seems to be that we're not pushing much money into those efforts. The current generation of commsat builders are all selling SEP orbit-raising solutions for GTO -> GSO, the "all-electric propulsion satellite"; The business model seems to be "Two satellites for the launch price of one" rather than "A satellite twice as big" for now.With SEP you can basically get the first ~3 km/s of your LEO departure speed done outside of the vehicle actually being sent TMI which is achieving the remaining ~2 km/s vastly easier if your doing it via chemical propulsion because your propellant is now less massive then the payload.
Again you completely hand-wave away complexity of a massive SEP booster stage. The fully loaded and fueled Mars transit vehicle/lander would mass what? 200t? 400t? More? What delta-V of departure burn from the assembly point (EML1) is necessary? How large would your super-duper-mega SEP need to be to accelerate the assembled Mars craft to a 3 km/s boost at perigee? How long would it take to build that up the 3 km/s? Please... Some specifics. I dare you. (as a bonus, estimate the cost of establishing such an infrastructure of smaller SEP tugs and massive SEP boosters - one needed for every MCT!)
Let's work with this. Let's say that after the BFR+MCT unknown is answered, we can get 100 tons of large-diameter freight to the high end of LEO in a completely reusable manner, with nothing left in orbit.With SEP you can basically get the first ~3 km/s of your LEO departure speed done outside of the vehicle actually being sent TMI which is achieving the remaining ~2 km/s vastly easier if your doing it via chemical propulsion because your propellant is now less massive then the payload.
Again you completely hand-wave away complexity of a massive SEP booster stage. The fully loaded and fueled Mars transit vehicle/lander would mass what? 200t? 400t? More? What delta-V of departure burn from the assembly point (EML1) is necessary? How large would your super-duper-mega SEP need to be to accelerate the assembled Mars craft to a 3 km/s boost at perigee? How long would it take to build that up the 3 km/s? Please... Some specifics. I dare you. (as a bonus, estimate the cost of establishing such an infrastructure of smaller SEP tugs and massive SEP boosters - one needed for every MCT!)
You seem to be taking a Zubrinist position, that no new technology is needed and simple 'brute force' can achieve all of SpaceX's mars goals. Cost was no object to Zubrin he didn't bat an eyelash at 100 Billion to perform a single mission to mars which would have been flags and footprints.
Has anyone suggested the possibility of SpaceX's announcement including "a reusable, nuclear (fission, fusion, either or) upper stage for the BFR" yet?
Has anyone suggested the possibility of SpaceX's announcement including "a reusable, nuclear (fission, fusion, either or) upper stage for the BFR" yet?
I think a nuclear fusion engine using magnetic control of the plasma and having one end expel the plasma for propulsion with the other end feeding liquid hydrogen in and continuously making plasma for propulsion would work. It would still require a lot of power just to use lasers to create the fusion at one end.Yeah... these are tangents that really don't belong in this thread. Fusion engineering is comparable in scope to the most expensive Mars programs proposed. Fusion propulsion is way, way out there, quite possibly postdating practical fusion power; Not that it shouldn't be explored, but this is a thing that Elon Musk wants to fly in a decade, not in half a century.
However if you are going to use some other type of power source to create the fusion for the engine, one still needs a lot of power, and you have to have heavy hydrogen to operate. How much I don't know?
So it seems SEP with very large arrays and a cluster of a lot of ion engines and a large tank of propellant could also make a very large in space, higher speed cargo or even human run to Mars.
A small NEP reactor, sealed, use for several years, and throw away, by shooting it to the Sun. It might be somewhat smaller and could make several Mars runs before depleting its fuel. Technology for small sealed reactors should be improved first.
I think the biconic shape is the way to go. I do think concentric fuel/oxygen tanks running the entire length of the MCT in the center would allow for more balance. Say with liquid methane in the center tube and the lox in a donut tube around the methane. Cargo, solar panels, landing legs, thrusters and thruster fuel, etc could be around this central tanking system. The picture on post 733 is what I like except with concentric center fuel/lox tanking. One can still dock on the nose, as tank dome could stop short of the nose. Heavy cargo could be stored near the bottom for easier unloading. Instead of the plug nozzle engine as shown, standard Raptors vacuum engines could be on the bottom. The interstage could double as heat protection flaps when coming back through the atmosphere.
It's not, in fact, impossible to land an SEP on Mars. Just kind of odd and mass inefficient. But with a sufficiently good solar array, you could do it. I should point out that an "ISS-esque" solar array is, in fact, retractable!
I'll also add that in addition to the trip duration, there's the "land the whole thing" mentality. If you're talking about ISS-esque solar array and radiator array, the idea of stowing them before EDL is not practical.
If instead you have a "Hermes"-type orbit-to-orbit ship in mind, then now SEP starts making sense.
But again, this is about MCT.
We KNOW SpaceX is considering SEP for Mars, and as it is impossible to land a SEP, that mean IPSO FACTO that they are have ALSO considered a Semi-Direct architecture of a transit-vehicle and a separate landing vehicle. And that such a vehicle trades very well against a single massive direct vehicle.
Has anyone suggested the possibility of SpaceX's announcement including "a reusable, nuclear (fission, fusion, either or) upper stage for the BFR" yet?
It's not, in fact, impossible to land an SEP on Mars. Just kind of odd and mass inefficient. But with a sufficiently good solar array, you could do it. I should point out that an "ISS-esque" solar array is, in fact, retractable!
To go back to balancing the Biconic for a bit, would it be possible to have a configuration similar to the Phoenix (see attached), except applied to a Biconic and with some sort of escape system in the nose (sorry, really like the idea of the nose section being a dedicated 'lifeboat'. I just see it as a way to solve so many problems) It would still have the central column, as a way to move down to the main pressurized area.
Not sure what this would do to the mass in total, but wouldn't it allow for a more evenly distributed mass during EDL?
I do like that image, it is close to what I have in mind for an MCT. But I think a lengthy tunnel through a cryogenic tank might have issues.
If you insist on have a nose-mounted LAS, another way around it might be to have just a giant central tank (split between oxidizer and fuel), and then have a lower cargo bay (just above engines), and an upper cargo bay in the nose. It would force some level of cargo balance, but that would allow you to put all the pressurized volume up top. And as a bonus the unpressurized/heavy cargo would be easily offloaded near the surface.
It's not, in fact, impossible to land an SEP on Mars. Just kind of odd and mass inefficient. But with a sufficiently good solar array, you could do it. I should point out that an "ISS-esque" solar array is, in fact, retractable!
Just on the Lifeboat idea, I know it was mentioned back-thread that it would be making things a lot more complex, but I think it would be worth it. Why? Not only can you Abort during Launch and EDL both at Earth AND Mars anywhere through the flight,
but it allows for something I've been worried about, Abort during transit.
If we assume these things are traveling in a fleet (say 10 MCT's for 1000 colonists to Mars?), if one suffers a mishap during flight, then the colonists can load up and punch out from the failing MCT and rendezvous with another nearby one. Granted, depending on the mishap, the 'other' MCT could just come to them anyway, but there may be issues necessitating a quick departure, for example during Mars approach and entry...I'm pretty sure you can't keep 100 people alive on an MCT for 2 years waiting for free-return to Earth if you miss.
The other MCT would have to put up with 200 people for a short time, while they can be distributed amongst the rest of the fleet, giving you 9 MCT's with 110 people each for the rest of the trip.
Just on the Lifeboat idea, I know it was mentioned back-thread that it would be making things a lot more complex, but I think it would be worth it. Why? Not only can you Abort during Launch and EDL both at Earth AND Mars anywhere through the flight,
Don't make the mistake of thinking that all an abort capsule needs to do is taking passengers away from an imminent explosion, to just survive the imminent danger. That works on Earth, because rescue is nearby, and the environment is survivable.
But that is NOT enough for Mars. An abort capsule is USELESS if it just dooms you to death a couple of hours later. To be a proper abort/rescue vehicle it basically needs to be a fully independent spacecraft/lander, built into another spacecraft/lander. By adding enough rescue/survival capability, you get into a dark spiral that eventually leads you to conclude that your escape capsule needs another escape capsule.
Being able to abort/rescue during transit is a completely separate aspect from whether or not you have a launch abort capsule. Any reasonable MCT proposal would have a fleet of them going during a launch window, so that rescue would be available even without abort capsules.
And no, I don't think you need to have a rapid way to escape your MCT during transit. At some point you just have to accept that there will be riskier times during a trip. Those who cannot accept risks can stay behind.
urm, that is quite a logic leap.
I'll also add that in addition to the trip duration, there's the "land the whole thing" mentality. If you're talking about ISS-esque solar array and radiator array, the idea of stowing them before EDL is not practical.
If instead you have a "Hermes"-type orbit-to-orbit ship in mind, then now SEP starts making sense.
But again, this is about MCT.
We KNOW SpaceX is considering SEP for Mars, and as it is impossible to land a SEP, that mean IPSO FACTO that they are have ALSO considered a Semi-Direct architecture of a transit-vehicle and a separate landing vehicle. And that such a vehicle trades very well against a single massive direct vehicle.
The fact that SpaceX were looking into SEP could mean they were looking at an all-in-one vehicle with both SEP and chemical propulsion.
MCT will remain a "land the entire thing" architecture, no matter how disagreeable you find the prospect of it.
My question is, why not send the crew up in traditional capsules to dock and load into the MCT.
How would these children fare if they return to earth with heavier gravity?
My question is, why not send the crew up in traditional capsules to dock and load into the MCT.
Because that'd be way expensive cumulatively than launching the MCT itself, and cost is a factor.
But surely if we're assuming a mostly developed Mars Colony ready to accept 1000 new immigrants, then it is reasonable to assume that they could mount a quick rescue of an abort capsule? Something like a dedicated MCT with enough fuel to do a sub orbital hop, on station and ready to go could make it to the capsules landing site in a few hours, and if the capsule was close enough, you could even go over land. That's why it would only need to sustain the colonists for a few hours, not days or weeks. That keeps the capsule more simplistic. Simple CO2 scrubbers, maybe a dedicated radiator if the base hull rejection rates are naturally too low and batteries for power. Add some pressure fed hypergolic rockets, or methalox ones that are made so the tanks can be 'charged' from the main tanks when they are likely to be needed (for those who don't like hypergols) and a parachute or two to help slow down. Again, I'm not saying it'll be 'cheap' either in terms of mass or money, I'm saying that it is worth it. It would have to cost more than $1.5 billion to save 100 lives for it to be considered 'Grossly improportional' and for SpaceX to have a reasonable argument that is was too expensive.
I was making an example of how the statement that they were "looking into SEP" could be understood in a different manner, not an actual plan.The fact that SpaceX were looking into SEP could mean they were looking at an all-in-one vehicle with both SEP and chemical propulsion.
As was already mentioned in the thread, that would require the large solar arrays not only somehow be retracted into the MCT itself but also be deployed again after re-launch from Mars, and thus be so reliably shielded from re-entry (and issues on the surface like dust) that you can guarantee they will redeploy for the return trip to Earth. There's simply no advantage to justify that complexity.
He wasn't comparing traditional mars architectures to MCT, he was saying that he was not intending on starting off with a cycler or using Dragons adapted for landing. When asked what vehicle he would use, he replied that "you should just land the entire thing."MCT will remain a "land the entire thing" architecture, no matter how disagreeable you find the prospect of it.
The original land-the-entire-thing comment was made in the context of the typical multi-stage launch/landing architectures. It doesn't mean that Musk has ruled out having a SEP stage that remains in orbit.
And given that he's happy to do EOR refuelling, to avoid the restrictions of monolithic launches that are typical of Mars architectures, I can't see why anyone would think he'd arbitrarily rule out such an option just to satisfy one interpretation of a throw away quote.
We might find out in a month or so. Until then, these options are on the table... "No matter how disagreeable you find the prospect."
Musk’s architecture for this human Mars exploration effort does not employ cyclers, reusable spacecraft that would travel back and forth constantly between the Red Planet and Earth — at least not at first.source (http://www.space.com/18596-mars-colony-spacex-elon-musk.html)
"Probably not a Mars cycler; the thing with the cyclers is, you need a lot of them," Musk told SPACE.com. "You have to have propellant to keep things aligned as [Mars and Earth’s] orbits aren’t [always] in the same plane. In the beginning you won’t have cyclers."
Musk also ruled out SpaceX's Dragon capsule, which the company is developing to ferry astronauts to and from low-Earth orbit, as the spacecraft that would land colonists on the Red Planet. When asked by SPACE.com what vehicle would be used, he said, "I think you just land the entire thing."
Asked if the "entire thing" is the huge new reusable rocket — which is rumored to bear the acronymic name MCT, short for Mass Cargo Transport or Mars Colony Transport — Musk said, "Maybe."
around 4:25 "So we're looking at solar-electric propulsion. I think we're gonna umm, look at some other interesting IN-space propulsion technologies..."
The fact that SpaceX were looking into SEP could mean they were looking at an all-in-one vehicle with both SEP and chemical propulsion.
As was already mentioned in the thread, that would require the large solar arrays not only somehow be retracted into the MCT itself but also be deployed again after re-launch from Mars, and thus be so reliably shielded from re-entry (and issues on the surface like dust) that you can guarantee they will redeploy for the return trip to Earth. There's simply no advantage to justify that complexity.MCT will remain a "land the entire thing" architecture, no matter how disagreeable you find the prospect of it.
The original land-the-entire-thing comment was made in the context of the typical multi-stage launch/landing architectures. It doesn't mean that Musk has ruled out having a SEP stage that remains in orbit.
And given that he's happy to do EOR refuelling, to avoid the restrictions of monolithic launches that are typical of Mars architectures, I can't see why anyone would think he'd arbitrarily rule out such an option just to satisfy one interpretation of a throw away quote.
We might find out in a month or so. Until then, these options are on the table... "No matter how disagreeable you find the prospect."
His original understanding based on the interview he gave was that they would try to eliminate all complexity by doing a mission without any orbital rendezvous: Take off from Earth, go directly to Mars (in 3-4 months), come directly back.
The requirements for an MCT vehicle which does perfect massless aerocapture and refuels only in LEO and on the Martian surface with ~100d transfers are still over 10km/s in that one stage.
The requirements for an MCT vehicle which does perfect massless aerocapture and refuels only in LEO and on the Martian surface with ~100d transfersare still over 10km/s in that one stage.are still 7.6km/s in that one stage.
Are you sure you are adding up the right numbers? Because with the most efficient (not fastest) Hohmann transfer, i get ~9.5 km/s from LEO to the Martian surface, but that is with NO aerobraking at Mars. With aerobraking that should come down to less than 8km/s, right?
Yes, a lot of the ideas that Musk have thrown out for "MCT" seem almost impossible (or actually impossible) if one assumes they are all true, at once. Many of them I interpret as "nice to have" for future versions, and one of them I see that way is the "fast transit". I don't expect that to happen until a propellant depot is established in LMO.
I'm strongly in favor of SEP and see considerable use for it making VERY FAST transit possible.Do you have math to back this mission plan up?
How is this possible? Wouldn't the system need to be huge and have magical power sources and such (insert anything Zubrin has ever said). NO, you can get a Fast transit on the order of 100 days to Mars with a slow, low power SEP system.
The trick is you use your SEP to move your Mars bound vehicle with propellants up to high Earth orbit and then then drop by the Earth for a huge Oberth assisted burn. For 2 km/s you should leave Earth with huge escape velocity and reach mars in 100 days (average).
Now the problem is capturing at mars, the answer is Magneto-Plasma Aerocapture, this lets us avoid expensive propulsive capture and is then followed by about a week of Plasma assisted aerobraking which lets the eventual EDL be from a gentle 4 km/s. So we get to have both fast transit and easy low speed EDL.
The SEP system has not even left Earth yet in this scenario, so you can do either one of two things, bring it back down to LEO for refueling and do it again (basically making it a Cis-lunar tug), or send it to mars by the conventional slow method of spiraling out from the Earths SOI (the SEP is too delicate to take the high thrust of the Oberth maneuver). In the latter case your going to arrive much later then the manned capsule but if this is a conjunction mission the crew will be spending around 600 days on mars so their is plenty of time for the SEP to arrive before it is needed for departure which is what I favor.
The MCT would only need to reach low Mars orbit and would then rendezvous with the SEP and head for Earth, this return transit is made reasonably short by the fact the MCT is a completely dry shell now of only 100 mT (75 vehicle mass + 25 return cargo) and the SEP is nearly dry too so power to weight ratios are increased, also were not aiming to match Earth's orbit and capture gently, were going to simply intersect it on an elliptical orbit around the sun, that cuts the DeltaV needed. At Earth we used the Magneto to capture again and bring both SEP and the MCT down to LEO (they probably need to separate to do this as the SEP is more delicate and would slow the process down for the MCT). The crew can be retrieved via a Dragon capsule now, and we need to send another tanker to LEO to put landing propellants into the MCT, if we use enough the MCT can do a lot of retro-propulsion on entry and bring it's entry speed down from the 7.7 km/s of orbit down to the range of 4 km/s which matches it's mars entry speed, so all the thermal protection systems can be designed for this low performance point.
IMLEO is estimated at 570 mT of which 100 mT is the cargo load, 75 mT is the MCT dry mass, 200 mT is chemical propellant in the MCT (2 tanker loads of 100 mT each), 155 is SEP propellant, 15 mT is the SEP tank and 22 mT is the SEP hardware which has a power output of 4.5 MW which corresponds to an alpha value of 5 kg/kw.
BTW Using a braking system like Magneto Plasma is the only way I can see an Integrated Bi-conic and direct Earth return being viable, without it the entry conditions are too extreme to meet the low dry mass fractions that it's advocates are proposing.
Edit: For the case of the highly simplified circular Mars / circular Earth orbit rendezvous, based on:
http://forum.nasaspaceflight.com/index.php?topic=37536.msg1371984#msg1371984
If you start at high Earth orbit and descend to perigee, then you can burn for 2241m/s to raise aphelion to 3.31AU, which drops time of flight to 100.03 days. Then you come in to the Mars approach at 12880m/s, and have to burn off 9275m/s in order to capture into a highly elliptical orbit. This is fairly difficult (capture intensity goes up faster than linearly because you have a shorter chord of atmosphere to cut through, not just less time in that atmosphere), and exceeds the capabilities of any chemical propulsion capacity in this planning exercise by a large measure.
-snip- (see original post) -snip-Yeah, I made an error there (I added escape velocity twice). 7.6km/s to get to Mars from LEO in 100 days assuming that perfect aerobraking accepts all excess velocity for free. But you go too far on the other point: Without aerobraking you'rs looking at 9.3km/s just for capture, making the total stage capability more like 16.9km/s. If you assume a direct descent to a Vacuum Mars Analog, the figure is 12.9km/s plus gravity losses, making for a total stage capability of 20.5km/s.
The tradeoff is: you can keep the BFR as a 4.5km/s vehicle instead of making it a (3.2km/s earth escape) + (2.4km/s Impaler Short Transit injection) + (~2km/s or whatever you need for EDL) ~= a 7.6km/s vehicle.
You're also going to need comparable efforts to get the thing home on a fast transit, though.
My question is, why not send the crew up in traditional capsules to dock and load into the MCT. If the MCT has to refuel in LEO. The refueling station could be where the crew transfers in. No need to put launch escape system on MCT.
My question is, why not send the crew up in traditional capsules to dock and load into the MCT.Because that'd be way expensive cumulatively than launching the MCT itself, and cost is a factor.
Now, you may argue that this is me reading too deeply into quotes
But nowhere did I find any indication that they are stepping away from a monolithic MCT and looking into having dedicated in-space elements, which means the original statement by Elon that their design goal for the MCT is "land the whole thing" remains a valid criterion in ascertaining whether or not a marsbound architecture can be considered an MCT architecture or not.
Now, you may argue that this is me reading too deeply into quotes and that any design posted in this thread is, as it were, cut from whole cloth. While that is all true (Especially the former, as I love reading too deeply into things), the purpose of this thread is to design a vehicle and accompanying architecture based entirely on criteria extracted from the fragments of information we are able to scrounge up.
I like Impaler's proposed architecture. However, I disagree on it being a viable MCT.
It is 'grossly improportional' if it also cripples your architecture. And I still don't think you are understanding how capable your abort capsule would have to be to avoid black zones on Mars. And parachutes won't help much for something this size. (they don't scale up well, and Mars atmosphere is thin to begin with) Consider this... Delta-V to reach low Mars orbit is over 3 km/s. So if you have a problem when you are *almost* in orbit, you now have ~2(?) km/s of propulsive capability that this capsule needs to have in order to land softly - and you will land far over the horizon, on the other side of the planet. Any capsule capable of 2+ km/s that is large enough to hold 100(!) people is going to be massive. MASSIVE. Your simple capsule just won't cut it.
Sometimes abort capability is impractical, and you mitigate risks other ways. You make your vehicle capable of abort, by having engine out capabilities and redundant systems. But you cannot remove all risks. Airliners still sometimes fall out of the sky, yet we don't have escape pods or parachutes.
But I'm tired of arguing these points over and over. For some, launch abort seems to be an essential that they just cannot comprehend going without.
The requirements for an MCT vehicle which does perfect massless aerocapture and refuels only in LEO and on the Martian surface with ~100d transfersare still over 10km/s in that one stage.are still 7.6km/s in that one stage.
Are you sure you are adding up the right numbers? Because with the most efficient (not fastest) Hohmann transfer, i get ~9.5 km/s from LEO to the Martian surface, but that is with NO aerobraking at Mars. With aerobraking that should come down to less than 8km/s, right?
Yes, a lot of the ideas that Musk have thrown out for "MCT" seem almost impossible (or actually impossible) if one assumes they are all true, at once. Many of them I interpret as "nice to have" for future versions, and one of them I see that way is the "fast transit". I don't expect that to happen until a propellant depot is established in LMO.I'm strongly in favor of SEP and see considerable use for it making VERY FAST transit possible.Do you have math to back this mission plan up?
How is this possible? Wouldn't the system need to be huge and have magical power sources and such (insert anything Zubrin has ever said). NO, you can get a Fast transit on the order of 100 days to Mars with a slow, low power SEP system.
The trick is you use your SEP to move your Mars bound vehicle with propellants up to high Earth orbit and then then drop by the Earth for a huge Oberth assisted burn. For 2 km/s you should leave Earth with huge escape velocity and reach mars in 100 days (average).
Now the problem is capturing at mars, the answer is Magneto-Plasma Aerocapture, this lets us avoid expensive propulsive capture and is then followed by about a week of Plasma assisted aerobraking which lets the eventual EDL be from a gentle 4 km/s. So we get to have both fast transit and easy low speed EDL.
The SEP system has not even left Earth yet in this scenario, so you can do either one of two things, bring it back down to LEO for refueling and do it again (basically making it a Cis-lunar tug), or send it to mars by the conventional slow method of spiraling out from the Earths SOI (the SEP is too delicate to take the high thrust of the Oberth maneuver). In the latter case your going to arrive much later then the manned capsule but if this is a conjunction mission the crew will be spending around 600 days on mars so their is plenty of time for the SEP to arrive before it is needed for departure which is what I favor.
The MCT would only need to reach low Mars orbit and would then rendezvous with the SEP and head for Earth, this return transit is made reasonably short by the fact the MCT is a completely dry shell now of only 100 mT (75 vehicle mass + 25 return cargo) and the SEP is nearly dry too so power to weight ratios are increased, also were not aiming to match Earth's orbit and capture gently, were going to simply intersect it on an elliptical orbit around the sun, that cuts the DeltaV needed. At Earth we used the Magneto to capture again and bring both SEP and the MCT down to LEO (they probably need to separate to do this as the SEP is more delicate and would slow the process down for the MCT). The crew can be retrieved via a Dragon capsule now, and we need to send another tanker to LEO to put landing propellants into the MCT, if we use enough the MCT can do a lot of retro-propulsion on entry and bring it's entry speed down from the 7.7 km/s of orbit down to the range of 4 km/s which matches it's mars entry speed, so all the thermal protection systems can be designed for this low performance point.
IMLEO is estimated at 570 mT of which 100 mT is the cargo load, 75 mT is the MCT dry mass, 200 mT is chemical propellant in the MCT (2 tanker loads of 100 mT each), 155 is SEP propellant, 15 mT is the SEP tank and 22 mT is the SEP hardware which has a power output of 4.5 MW which corresponds to an alpha value of 5 kg/kw.
BTW Using a braking system like Magneto Plasma is the only way I can see an Integrated Bi-conic and direct Earth return being viable, without it the entry conditions are too extreme to meet the low dry mass fractions that it's advocates are proposing.
Edit: For the case of the highly simplified circular Mars / circular Earth orbit rendezvous, based on:
http://forum.nasaspaceflight.com/index.php?topic=37536.msg1371984#msg1371984
If you start at high Earth orbit and descend to perigee, then you can burn for 2241m/s to raise aphelion to 3.31AU, which drops time of flight to 100.03 days. Then you come in to the Mars approach at 12880m/s, and have to burn off 9275m/s in order to capture into a highly elliptical orbit. This is fairly difficult (capture intensity goes up faster than linearly because you have a shorter chord of atmosphere to cut through, not just less time in that atmosphere), and exceeds the capabilities of any chemical propulsion capacity in this planning exercise by a large measure.-snip- (see original post) -snip-Yeah, I made an error there (I added escape velocity twice). 7.6km/s to get to Mars from LEO in 100 days assuming that perfect aerobraking accepts all excess velocity for free. But you go too far on the other point: Without aerobraking you'rs looking at 9.3km/s just for capture, making the total stage capability more like 16.9km/s. If you assume a direct descent to a Vacuum Mars Analog, the figure is 12.9km/s plus gravity losses, making for a total stage capability of 20.5km/s.
The tradeoff is: you can keep the BFR as a 4.5km/s vehicle instead of making it a (3.2km/s earth escape) + (2.4km/s Impaler Short Transit injection) + (~2km/s or whatever you need for EDL) ~= a 7.6km/s vehicle.
You're also going to need comparable efforts to get the thing home on a fast transit, though.
These numbers are so high because 100 day transfers require lots of extra energy that 200 day transfers do not.
If you're content with 200 day transfers...
-numbers to come, WIP-
Burninate,
Thanks for those numbers. The wet to dry mass numbers are fascinating as well.
I certainly favor the slower, more delta-V conservative approaches. And I think it is the best hope of making an "MCT" work, without a LMO propellant depot.
Burninate,
Thanks for those numbers. The wet to dry mass numbers are fascinating as well.
I certainly favor the slower, more delta-V conservative approaches. And I think it is the best hope of making an "MCT" work, without a LMO propellant depot.
Oh, I think there's still room for an LMO prop depot even in the most conservative case. There's still a question of whether slow-transit, LEO depot, plus LMO depot is *enough* for a vehicle with integrated ISRU capacity on Mars to do a first mission to a new site reusably. A lot of it falls on what the mass return ratio for the ISRU system ends up being.
If it's not high enough, you either *have* to resort to a nested MAV (and return the big'un next synod), or you just have to throw away a whole MCT for every site you land at, as a permanently landed asset.
Burninate,
Thanks for those numbers. The wet to dry mass numbers are fascinating as well.
I certainly favor the slower, more delta-V conservative approaches. And I think it is the best hope of making an "MCT" work, without a LMO propellant depot.
Oh, I think there's still room for an LMO prop depot even in the most conservative case. There's still a question of whether slow-transit, LEO depot, plus LMO depot is *enough* for a vehicle with integrated ISRU capacity on Mars to do a first mission to a new site reusably. A lot of it falls on what the mass return ratio for the ISRU system ends up being.
If it's not high enough, you either *have* to resort to a nested MAV (and return the big'un next synod), or you just have to throw away a whole MCT for every site you land at, as a permanently landed asset.
That has always been my assumption - that every new site would have not just one, but several unmanned MCT land to set up ISRU and other equipment. Most of those initial unmanned "pathfinder" MCT would not be returned, I would expect - They would instead become the first outpost habitats, storage sheds, and MCT spare part depots. ;)
This is not flags and footprints that are planned here. The goal is to create a permanent manned outpost. But the first crews would not be launched until the next launch window after ISRU production is well underway.
That has always been my assumption - that every new site would have not just one, but several unmanned MCT land to set up ISRU and other equipment. Most of those initial unmanned "pathfinder" MCT would not be returned, I would expect - They would instead become the first outpost habitats, storage sheds, and MCT spare part depots. ;)
This is not flags and footprints that are planned here. The goal is to create a permanent manned outpost. But the first crews would not be launched until the next launch window after ISRU production is well underway.
The threshold you have to pass, then, is being able to set up and run an ISRU water-mining workflow without any human assistance, perhaps without any realtime telerobotics, if we don't go for a landing.
You also need it to last a decade.
Burninate, I still see a potential for SEP instead of the magneto hydrodynamic breaking in what you describe.
Imagine going with say a 120 day transfer to a point where Mars would have been about 20 days before the craft gets there, but for the full length of the 150 day flight SEP is running at 0.4 mm/s2 to alter the orbit so that it actually arrives at Mars later, a little further along on Mars orbit, but at almost the same speed for a relatively low energy capture (say around 1km/s for capture).
and yes that is just roughed out, I don't have Andy Weir's continuous thrust orbit model software. So really it is a matter of optimizing for the mass of the SEP system including solar arrays against how much faster than a standard half ellipse Hohmann you want to fly.
That has always been my assumption - that every new site would have not just one, but several unmanned MCT land to set up ISRU and other equipment. Most of those initial unmanned "pathfinder" MCT would not be returned, I would expect - They would instead become the first outpost habitats, storage sheds, and MCT spare part depots.
It's not, in fact, impossible to land an SEP on Mars. Just kind of odd and mass inefficient. But with a sufficiently good solar array, you could do it. I should point out that an "ISS-esque" solar array is, in fact, retractable!
I'll also add that in addition to the trip duration, there's the "land the whole thing" mentality. If you're talking about ISS-esque solar array and radiator array, the idea of stowing them before EDL is not practical.
If instead you have a "Hermes"-type orbit-to-orbit ship in mind, then now SEP starts making sense.
But again, this is about MCT.
We KNOW SpaceX is considering SEP for Mars, and as it is impossible to land a SEP, that mean IPSO FACTO that they are have ALSO considered a Semi-Direct architecture of a transit-vehicle and a separate landing vehicle. And that such a vehicle trades very well against a single massive direct vehicle.
This is my preference as well. Absent a good characterization of MAC, let's keep its use at a manageably small level. We have a good characterization of SEP, on the other hand, and SEP can be used for much of capture. The simplified case that's most conservative on MAC would be to go with the 180-day version I calculated above, but plan on burning for about 2 months while approaching aphelion. There are other benefits to using modest amounts of SEP; It makes things like Mars-orbit habs and propellant much easier to deal with.
It's not, in fact, impossible to land an SEP on Mars. Just kind of odd and mass inefficient. But with a sufficiently good solar array, you could do it. I should point out that an "ISS-esque" solar array is, in fact, retractable!
I'll also add that in addition to the trip duration, there's the "land the whole thing" mentality. If you're talking about ISS-esque solar array and radiator array, the idea of stowing them before EDL is not practical.
If instead you have a "Hermes"-type orbit-to-orbit ship in mind, then now SEP starts making sense.
But again, this is about MCT.
We KNOW SpaceX is considering SEP for Mars, and as it is impossible to land a SEP, that mean IPSO FACTO that they are have ALSO considered a Semi-Direct architecture of a transit-vehicle and a separate landing vehicle. And that such a vehicle trades very well against a single massive direct vehicle.
RB. You just used the word "in fact" along with stating you can build an engine with a T/W of 0.25, and an ISP in the thousands.
Forget Mars... With your engine you can roam the solar system at will!
You need a sanity check sometimes. You can't just add brochure numbers for this and that (solar panels, thruster) and arrive at meaningful numbers.
Would anyone like to do the work of reversing that spreadsheet so we could look at more refined estimates of return delta V? I might try it at some point, but not today.
That has always been my assumption - that every new site would have not just one, but several unmanned MCT land to set up ISRU and other equipment. Most of those initial unmanned "pathfinder" MCT would not be returned, I would expect - They would instead become the first outpost habitats, storage sheds, and MCT spare part depots.
Hmmm, in theory, only the first site or two would need that. Once you have the initial 3-6 MCTs left at those 1-2 sites, they could be used as suborbital hoppers for ferrying equipment (even people) to secondary sites to prep them for the incoming (fully reusable) MCTs.
[images of Apollo CM+SM and Dragon+trunk]
Apollo didn't and Dragon doesn't do EDL with those modules attached. The level of integration your require is an order of magnitude more complex.
Just on the Lifeboat idea, I know it was mentioned back-thread that it would be making things a lot more complex, but I think it would be worth it. Why? Not only can you Abort during Launch and EDL both at Earth AND Mars anywhere through the flight,
Don't make the mistake of thinking that all an abort capsule needs to do is taking passengers away from an imminent explosion, to just survive the imminent danger. That works on Earth, because rescue is nearby, and the environment is survivable.
But that is NOT enough for Mars. An abort capsule is USELESS if it just dooms you to death a couple of hours later. To be a proper abort/rescue vehicle it basically needs to be a fully independent spacecraft/lander, built into another spacecraft/lander. By adding enough rescue/survival capability, you get into a dark spiral that eventually leads you to conclude that your escape capsule needs another escape capsule.
But there is another way. Build safety margins and redundancy into your MCT/lander and forego an abort system that is effectively useless.
[Dragon's trunk] It is 2/3 of the way to a lifeboat on a biconic MCT.
So, how often have liquid boosters went "boom" in the US?
How important is it to protect against that one failure mode?
They're reversible, but you're forgetting the Oberth effect: because on the way from Earth to Mars, you can dump your exhaust in a deeper gravity well than Mars to Earth, it takes less delta-v.Would anyone like to do the work of reversing that spreadsheet so we could look at more refined estimates of return delta V? I might try it at some point, but not today.
Orbits are time reversible so don't we just need to look at the 'Vinf at mars' and calculate the DeltaV needed to archive that escape velocity from mars surface, then we would (if we pointed ourselves in the right direction) be headed back down the equivalent half of the outbound orbit and we should reach Earth in the specified time and with the specified Vinf so we know what we need to do to aerocapture at Earth as well.
It would be nice to have this done for us on the spreadsheet though.
So, how often have liquid boosters went "boom" in the US?
How important is it to protect against that one failure mode?
Antares, October 28, 2014.
Falcon 9, June 28, 2015.
That's two in the past 12 months.
At least launching from Earth, it's pretty important. As you mentioned, for early MCT missions with a small crew, send the crew up on a Dragon 2.
So, how often have liquid boosters went "boom" in the US?
How important is it to protect against that one failure mode?
Antares, October 28, 2014.
Falcon 9, June 28, 2015.
That's two in the past 12 months.
At least launching from Earth, it's pretty important. As you mentioned, for early MCT missions with a small crew, send the crew up on a Dragon 2.
Not quite. I was referring to exploding boosters.
As I understand, F9's was the 2nd stage which blew up causing the booster to then fail, which is different than the booster itself exploding, and trying to abort away form it. If that were an Integrated MCT, then the LAS lifeboat would be within a few meters of the exploding tank. Could an LAS could save it with such close proximity? (Which was one of the problems with putting Orion on a side mounted SDHLV if I recall correctly. Orion would be right next to the ET, so even with an LAS system, it's unlikely the crew could get away from a sudden ET explosion).
The F9 v1.1 booster failure from a few years ago was an engine out, which as I said, would be accounted for.
Antares was also an engine out. But with just 2 engines, there was no engine out redundancy so they terminated the LV, as I understand. Had Antares had engine out redundancy, there's not reason to think it wouldn't have made staging nominally.
But, yes, for that first several dozen flights of MCT, the crew can just go up on a simple F9/D2. There will be 4 active Falcon pads in operation after one, 3 on the East Coast. One of those launched per Mars mission seems pretty reasonable.
By the time you'd be looking at putting 100 people on it, there'll be many exploration missions over decades. You'll probably have a pretty good idea of that point if it's reasonable to put people on MCT for launch, or go another route like a "Big Dragon" on a FH or something that can hold 100 people, and have a Earth LAS system. Or some other route.
So, how often have liquid boosters went "boom" in the US?
How important is it to protect against that one failure mode?
Antares, October 28, 2014.
Falcon 9, June 28, 2015.
That's two in the past 12 months.
At least launching from Earth, it's pretty important. As you mentioned, for early MCT missions with a small crew, send the crew up on a Dragon 2.
Just on the Lifeboat idea, I know it was mentioned back-thread that it would be making things a lot more complex, but I think it would be worth it. Why? Not only can you Abort during Launch and EDL both at Earth AND Mars anywhere through the flight,
Don't make the mistake of thinking that all an abort capsule needs to do is taking passengers away from an imminent explosion, to just survive the imminent danger. That works on Earth, because rescue is nearby, and the environment is survivable.
But that is NOT enough for Mars. An abort capsule is USELESS if it just dooms you to death a couple of hours later. To be a proper abort/rescue vehicle it basically needs to be a fully independent spacecraft/lander, built into another spacecraft/lander. By adding enough rescue/survival capability, you get into a dark spiral that eventually leads you to conclude that your escape capsule needs another escape capsule.
But there is another way. Build safety margins and redundancy into your MCT/lander and forego an abort system that is effectively useless.
Lars,
You make some good points here.
My initial thought on the IBMCT is just that, to not have an LAS at all. To use Dragon 2 as a LEO taxi for exploration crews for the first couple of decades until MCT builds up a track record of reliability. Then go with the Space Shuttle or Airliner model when it's time to put a full 100 passengers on board. MCT should have dozens of launches under it's belt by then.
It'd have full booster engine out capability, as well as full IBMCT engine out capability. So an engine out during ascent on either stage would still result in a nominal LEO insertion.
It doesn't protect against one failure mode, an exploding booster. It's popular to look at the booster explosions on the N-1 tests and take that to mean you must protect against that, but what really needs to be looked at is how often to American liquid boosters actually explode once they are out of test phase and in to actual production? I don't know the answer to that exactly, but I don't think it's very often. Especially in modern times. The Challenger flight is one of the most famous examples, but that really wasn't a liquid booster failure. It was an o-ring joint failure from a segmented SRB than then caused a failure of the liquid booster. Had the O-ring failed on the outboard side so it didn't burn into the side of the ET, the stack likely wouldn't have exploded. Once detected they probably would have done an emergency orbiter abort and glided to a contingency landing site. Or the booster may have lasted to SRB staging.
There was a Delta II that failed spectacularly not so long ago, but I think that was an SRB failure as well, not of the liquid core. I don't think there was a booster failure of the a production Saturn 1B or Saturn V, or Atlas II, III, or V, or Delta IV, III, or II outside of the SRB failure.
And I think most of the Titan IV failures were SRB issues, and not the core (which wasn't really a booster, but a 2nd stage that ignited after SRB sep.)
So, how often have liquid boosters went "boom" in the US?
How important is it to protect against that one failure mode?
They're reversible, but you're forgetting the Oberth effect: because on the way from Earth to Mars, you can dump your exhaust in a deeper gravity well than Mars to Earth, it takes less delta-v.Would anyone like to do the work of reversing that spreadsheet so we could look at more refined estimates of return delta V? I might try it at some point, but not today.
Orbits are time reversible so don't we just need to look at the 'Vinf at mars' and calculate the DeltaV needed to archive that escape velocity from mars surface, then we would (if we pointed ourselves in the right direction) be headed back down the equivalent half of the outbound orbit and we should reach Earth in the specified time and with the specified Vinf so we know what we need to do to aerocapture at Earth as well.
It would be nice to have this done for us on the spreadsheet though.
Impaler: I don't think Musk mentioned 100 day trajectory on the return trip, just on the way there. Also, exactly 100 is a little bit of, um, spurious precision. 102 days is essentially the same thing. (I know you know this, just want to point it out.)The 102 day transfer above is very different from the 100 day transfer. I included it because it's the author's default values on the spreadsheet; The decision is made to go with a perihelion below 1AU in an attempt to balance propulsive capture with propulsive transfer injection, unlike in the 100 day transfer.
Not sure SpaceX would start at LEO. A high orbit seems more realistic, as it allows you to leverage SEP without actually including SEP on MCT directly. SEP would be just used to haul up propellant from LEO. This helps reduce IMLEO a LOT.
But yeah, I still think MCT will start on the surface of Mars and go straight to Earth, with a much-longer-than-100-day trajectory.
So, how often have liquid boosters went "boom" in the US?
How important is it to protect against that one failure mode?
Antares, October 28, 2014.
Falcon 9, June 28, 2015.
That's two in the past 12 months.
At least launching from Earth, it's pretty important. As you mentioned, for early MCT missions with a small crew, send the crew up on a Dragon 2.
You say that a six month transfer to Mars is intolerable for passengers, so you guys are proposing that we go to these great lengths to avoid it. And the identical six month transfer back from Mars is... more tolerable?
Dragon survived the Falcon 9 second stage failure, that was not a "boom".
You say that a six month transfer to Mars is intolerable for passengers, so you guys are proposing that we go to these great lengths to avoid it. And the identical six month transfer back from Mars is... more tolerable?
100 people out, mostly colonists, 15-20 back, mostly crew. Different rules.
Great work, but could you please break out your math a bit more rather than listing totals? For example, what dV do you assume for Mars Ascent leg?
Also: 300 tons of propellant in Mars orbit is not a very high amount, given that you can use SEP all the way and the gear ratio is excellent on that; If you want a short transit you have to send more launches worth of propellant to deep elliptical earth orbit for the outbound leg than to LMO for the inbound leg (using your SEP propellant tug idea), which in turn is far fewer than if you avoided DEO rendezvous altogether and stuck to LEO rendezvous only.
Impaler: I don't think Musk mentioned 100 day trajectory on the return trip, just on the way there. Also, exactly 100 is a little bit of, um, spurious precision. 102 days is essentially the same thing. (I know you know this, just want to point it out.)
Not sure SpaceX would start at LEO. A high orbit seems more realistic, as it allows you to leverage SEP without actually including SEP on MCT directly. SEP would be just used to haul up propellant from LEO. This helps reduce IMLEO a LOT.
But yeah, I still think MCT will start on the surface of Mars and go straight to Earth, with a much-longer-than-100-day trajectory.
You say that a six month transfer to Mars is intolerable for passengers, so you guys are proposing that we go to these great lengths to avoid it. And the identical six month transfer back from Mars is... more tolerable?
and yet people here easily accept that MCT is going to land dozens of people and/or dozens of tons of cargo?
SpaceX having succeeded in launching resupply missions to the ISS as well as satellites to LEO and GTO doesn't make statements about "Mars colonization plans" (which are vastly more ambitious and difficult) any more credible.and yet people here easily accept that MCT is going to land dozens of people and/or dozens of tons of cargo?Because of his successes, Musk's statements, even his off the cuff remarks, receive far more credibility than they might otherwise.
If Interorbital Systems were proposing the MCT it would be greeted with snorts of derision. But Musk has succeeded so often in the past, and is saying something we all fervently want to be so, that we sometimes forget that Musk is only human after all.
SpaceX having succeeded in launching resupply missions to the ISS as well as satellites to LEO and GTO doesn't make statements about "Mars colonization plans" (which are vastly more ambitious and difficult) any more credible.and yet people here easily accept that MCT is going to land dozens of people and/or dozens of tons of cargo?Because of his successes, Musk's statements, even his off the cuff remarks, receive far more credibility than they might otherwise.
If Interorbital Systems were proposing the MCT it would be greeted with snorts of derision. But Musk has succeeded so often in the past, and is saying something we all fervently want to be so, that we sometimes forget that Musk is only human after all.
SpaceX having succeeded in launching resupply missions to the ISS as well as satellites to LEO and GTO doesn't make statements about "Mars colonization plans" (which are vastly more ambitious and difficult) any more credible.and yet people here easily accept that MCT is going to land dozens of people and/or dozens of tons of cargo?Because of his successes, Musk's statements, even his off the cuff remarks, receive far more credibility than they might otherwise.
If Interorbital Systems were proposing the MCT it would be greeted with snorts of derision. But Musk has succeeded so often in the past, and is saying something we all fervently want to be so, that we sometimes forget that Musk is only human after all.
SpaceX having succeeded in launching resupply missions to the ISS as well as satellites to LEO and GTO doesn't make statements about "Mars colonization plans" (which are vastly more ambitious and difficult) any more credible.
NASA has already landed several rovers on Mars. SpaceX still hasn't even landed their own demonstration payload there yet.
[Dragon's trunk] It is 2/3 of the way to a lifeboat on a biconic MCT.
It really, really isn't.
NASA has already landed several rovers on Mars, so they have at least some experience with Mars missions. SpaceX still hasn't even landed their own demonstration payload there yet.To add to what others have said, they are also the first and only entity so far to do supersonic retro propulsion. In fact NASA learning from SpaceX with respect to this aspect of a Mars landing.
An innovative partnership between NASA and SpaceX is giving the U.S. space agency an early look at what it would take to land multi-ton habitats and supply caches on Mars for human explorers, while providing sophisticated infrared (IR) imagery to help the spacecraft company develop a reusable launch vehicle
......then you aren't describing a real colonization effort. A colony isn't just a "bigger base." Either there's a much larger net flux toward Mars than away /or/ MCT won't be needing to carry 100 passengers anyway because not enough people will be going. And then you shouldn't call it MCT.
Agreed, and I do not think this issue can be brushed away by claiming that their are few (or no) return passengers. All trips to mars will come with FREE return according to Musk, and I fully expect all personnel to rotate in and out at a rate that is so close to 1:1 that it might as well be that....
...What, exactly, are you talking about? Humans can withstand 6 gees for at least 10 minutes, which is 36km/s of delta-v, usually while still doing simple tasks (unnecessary in this case, since capture/entry/reentry would be automated). Such a long period of fairly constant deceleration is possible using a lifting (re)entry, actually including negative lift. Once captured, your next trip through the atmosphere can be for landing (possibly with a skip). So you just need to get rid of your hyperbolic velocity on the first pass because the next pass will take care of actual (re)entry.
I don't think we will actually see 100 day transits mainly do to capture g-force limitations...
I find the whole 'NASA risk aversion' myth so annoying and dumb. People seem to forget all the failed NASA missions and development efforts that were so crazy ambitious. Dose anyone remember the X-33 which was in all likelihood a less ambitious vehicle/mission profile then MCT? NASA isn't caviler with astronaut lives and neither is SpaceX nor will they ever be if they want to actually sell tickets to normal people.
NASA is not monolithic. Nor is NASA unchanging through time. Some elements of current NASA are not risk averse, but other elements are. But risk aversion is not bad by itself, unless it is taken to the extreme. (Human spaceflight being part of that - Witness Orion/SLS, the "Apollo revived with Shuttle parts" mixture, designed to be as conservative as possible)
A side note: Your X-33 example would carry more weight if it actually flew. ;)
...What, exactly, are you talking about? Humans can withstand 6 gees for at least 10 minutes, which is 36km/s of delta-v, usually while still doing simple tasks (unnecessary in this case, since capture/entry/reentry would be automated). Such a long period of fairly constant deceleration is possible using a lifting (re)entry, actually including negative lift. Once captured, your next trip through the atmosphere can be for landing (possibly with a skip). So you just need to get rid of your hyperbolic velocity on the first pass because the next pass will take care of actual (re)entry.
I don't think we will actually see 100 day transits mainly do to capture g-force limitations...
I don't think this is any more radical than anything else SpaceX is doing. And isn't more radical than Shuttle's reentry. Sure, we need a good estimate of Mars' atmospheric state, but that is also a solvable problem.
36km/s of hyperbolic velocity is much more than enough for a 100 day transfer.
...lifting (re)entry is well within the state-of-the-art for both Mars and Earth. SpaceX already does this fairly regularly with Dragon. SpaceX also has a very good heatshield material (PICA-X) that they're very familiar with. I don't have any clue why they would shoot themselves in the foot by not leveraging these things to their full potential.
[Dragon's trunk] It is 2/3 of the way to a lifeboat on a biconic MCT.
and yet people here easily accept that MCT will be capable of landing dozens of people in the same vehicle that was used for the interplanetary transfer, or up to 100 tons of cargo?
The limitation I see with the magneto is that it doesn't produce lift. That makes the g-forces much higher because your doing a single short burst of drag.
The limitation I see with the magneto is that it doesn't produce lift. That makes the g-forces much higher because your doing a single short burst of drag.
I recall from the relevant thread that the possibility was mentioned that producing lift by making the magnetic braking device unsymmetric should be possible. In that case it even might be able to do a lot of steering and adjusting to different conditions of the local atmosphere.
and yet people here easily accept that MCT will be capable of landing dozens of people in the same vehicle that was used for the interplanetary transfer, or up to 100 tons of cargo?
There are at least 4 threads on just this topic, in just this section, of people not merely accepting that.
The limitation I see with the magneto is that it doesn't produce lift. That makes the g-forces much higher because your doing a single short burst of drag.
I recall from the relevant thread that the possibility was mentioned that producing lift by making the magnetic braking device unsymmetric should be possible. In that case it even might be able to do a lot of steering and adjusting to different conditions of the local atmosphere.
The same thought had occurred to me, but I realized that the combined vehicle would simply pivot to bring the center of mass on the end of the tether into line with the center of drag. The effect is analogous to deploying two different sized round non-lift producing parachutes at the same time, they would create different amounts of drag but your line angles would simply be skewed such that the load is directly under the center of drag and you would descend vertically. Or at least this is my mental picture of what would happen when the plasma is not deflecting any air laterally which by definition this can't do because it's drag is entirely from ion collisions which scatter in all directions and end with the incoming gas just being swallowed into the existing plasma vortex, without deflection you can't have lift because you would be violating conservation of angular momentum.
I'm sorry, but where exactly would those be? All I seem to see are people speculating on how MCT/BFR would work, and taking such a system for granted (i.e. they don't seem to acknowledge that EDL for heavy payloads is very hard, unlike every other discussion on Mars mission architectures)and yet people here easily accept that MCT will be capable of landing dozens of people in the same vehicle that was used for the interplanetary transfer, or up to 100 tons of cargo?
There are at least 4 threads on just this topic, in just this section, of people not merely accepting that.
You're still speaking entirely in generalities. What sort of g-load do you believe is a limit? You can't make a claim like you're making (that g-loads are prohibitively high for 100 day transits) without giving some numbers....What, exactly, are you talking about? Humans can withstand 6 gees for at least 10 minutes, which is 36km/s of delta-v, usually while still doing simple tasks (unnecessary in this case, since capture/entry/reentry would be automated). Such a long period of fairly constant deceleration is possible using a lifting (re)entry, actually including negative lift. Once captured, your next trip through the atmosphere can be for landing (possibly with a skip). So you just need to get rid of your hyperbolic velocity on the first pass because the next pass will take care of actual (re)entry.
I don't think we will actually see 100 day transits mainly do to capture g-force limitations...
I don't think this is any more radical than anything else SpaceX is doing. And isn't more radical than Shuttle's reentry. Sure, we need a good estimate of Mars' atmospheric state, but that is also a solvable problem.
36km/s of hyperbolic velocity is much more than enough for a 100 day transfer.
...lifting (re)entry is well within the state-of-the-art for both Mars and Earth. SpaceX already does this fairly regularly with Dragon. SpaceX also has a very good heatshield material (PICA-X) that they're very familiar with. I don't have any clue why they would shoot themselves in the foot by not leveraging these things to their full potential.
I never said crew was the limit. The limits are mostly on the vehicle and it's structure, remember that our aerocapture g's in a bi-conic vehicle are negative, aka nose bleed g's so they represent a 2nd direction that the vehicle needs to tolerate, and the more lift their is the more lateral g you get.
Aerocapture of a large mass at mars from a hohmann trajectory (low incoming speed and low DeltaV needed) would be possible with a bi-conic shape, but 100 day transits result in MUCH larger deceleration needs which means you simply need more surface area then the vehicle itself can provide. So some kind of 'expander' is needed, either an inflatable device (which is disposable) or the magneto which looks to be fully reusable. The limitation I see with the magneto is that it doesn't produce lift. That makes the g-forces much higher because your doing a single short burst of drag.
If magnetic deceleration can be used it would help a lot with reusability. If a first pass can achieve capture then a second or third pass can achieve landing. Replacing the heatshield after every flight will cost time and money. I don't see magnetic as a requirement but i hope for it. A metallic heatshield will not do especially for EDL back on earth.Even a PICA-X shield can take multiple entries (Musk said 100, which is an overstatement but is a lot more than MCT will need). And yes, a metallic shield would work, you just have to do it gently enough. But I suspect SpaceX is looking at PICA-X. A tanker going to LEO, most certainly, could use a metallic TPS without any hard limits on numbers of reuses (thousands are possible).
If magnetic deceleration can be used it would help a lot with reusability. If a first pass can achieve capture then a second or third pass can achieve landing. Replacing the heatshield after every flight will cost time and money. I don't see magnetic as a requirement but i hope for it. A metallic heatshield will not do especially for EDL back on earth.Even a PICA-X shield can take multiple entries (Musk said 100, which is an overstatement but is a lot more than MCT will need). And yes, a metallic shield would work, you just have to do it gently enough. But I suspect SpaceX is looking at PICA-X. A tanker going to LEO, most certainly, could use a metallic TPS without any hard limits on numbers of reuses (thousands are possible).
You're still speaking entirely in generalities. What sort of g-load do you believe is a limit? You can't make a claim like you're making (that g-loads are prohibitively high for 100 day transits) without giving some numbers.
As a side note, a Falcon 9 first stage already has 8 deployed aerosurfaces that generate drag. You think SpaceX will not be able to have ANY deployed aerosurfaces for MCT?
SpaceX, on the other hand, has never mentioned magnetohydrodynamic drag devices. You are basically handicapping them by disallowing any of their current ways of solving the problem so you can introduce your own pet solution.
I am already assuming separate aerocapture and entry at both ends and the DeltaV I stated were the minimum to capture, aka you go into a big elliptical orbit.
Given the size of the MCT and the mass it has to work with I think 2 gs's is the best that can be reasonably hopped for and it will still in fact be hard to make the vehicle that strong. Your not taking into account the square-cube law that makes a large object proportionally weaker then a small one.
NASA did studies for Inspiration Mars. They came to the conclusion that direct reentry from Mars at earth return would be the way to go assuming PicaX as a heatshield material. A skip reentry would increase heatshield stress.
I was considering using high pressure to stabilize MCT during reentry. Maybe 40 to 50 PSI or 3 times earth atomospheric pressure. This pressure during the minutes of reentry will not cause problems to crew given what we know from diving.
Is there any concept of a reverse gravity slingshot?- e.g. Use the gravity of the moon to slow down before re-entering earth's atmosphere.
NASA did studies for Inspiration Mars. They came to the conclusion that direct reentry from Mars at earth return would be the way to go assuming PicaX as a heatshield material. A skip reentry would increase heatshield stress.Yes, it's true that if you're just trying to minimize heatshield mass using an ablative heatshield, you often just want to go hot and fast. But in our discussion, it's peak acceleration that we're attempting to minimize, not heatshield mass. 6 gees is a pretty good peak acceleration to use. 5 gees has also been used. Nobody that I'm aware of has used 2 gees in any serious analysis. Heck, a /regular/ reentry uses more than 2 gees. So I find that disingenuous.
...I was considering using high pressure to stabilize MCT during reentry. Maybe 40 to 50 PSI or 3 times earth atomospheric pressure. This pressure during the minutes of reentry will not cause problems to crew given what we know from diving.Clever idea. I like it. This gives an important point: there is room for innovation in improving structural mass for a given load. Better materials and better manufacturing techniques are developed regularly. I don't think it's wise to sand-bag peak acceleration to something ridiculously low like 2 gees.
Is there any concept of a reverse gravity slingshot?- e.g. Use the gravity of the moon to slow down before re-entering earth's atmosphere.
First the moon's orbit is both slow & long and worse yet inclined off plane so the chance of the moon being where you need it to be is tine.
But given that the moon's mass (1/80th Earth) is too small to shed any meaningful velocity.
NASA did studies for Inspiration Mars. They came to the conclusion that direct reentry from Mars at earth return would be the way to go assuming PicaX as a heatshield material. A skip reentry would increase heatshield stress.Yes, it's true that if you're just trying to minimize heatshield mass using an ablative heatshield, you often just want to go hot and fast. But in our discussion, it's peak acceleration that we're attempting to minimize, not heatshield mass. 6 gees is a pretty good peak acceleration to use. 5 gees has also been used. Nobody that I'm aware of has used 2 gees in any serious analysis. Heck, a /regular/ reentry uses more than 2 gees. So I find that disingenuous.
The shuttle orbiter is about the same mass as I expect a dry MCT will be, but is in a shape that I don't expect MCT to be. Those wings are very heavy structurally. If it's the side of a pressure vessel, and the whole thing is pressurized to some extent or another, then MCT could be far more structurally efficient than Shuttle was. Additionally, MCT will have access to superior materials and manufacturing techniques.NASA did studies for Inspiration Mars. They came to the conclusion that direct reentry from Mars at earth return would be the way to go assuming PicaX as a heatshield material. A skip reentry would increase heatshield stress.Yes, it's true that if you're just trying to minimize heatshield mass using an ablative heatshield, you often just want to go hot and fast. But in our discussion, it's peak acceleration that we're attempting to minimize, not heatshield mass. 6 gees is a pretty good peak acceleration to use. 5 gees has also been used. Nobody that I'm aware of has used 2 gees in any serious analysis. Heck, a /regular/ reentry uses more than 2 gees. So I find that disingenuous.
Bigger spacecraft can tolerate less g forces, that is a clear pattern. If I tried to say a small sample return capsule of under 1 m couldn't tolerate more then 6 g's then THAT would be sandbagging as such capsules and missiles routinely tolerate >20 g's.
The Shuttle was the largest re-entry vehicle ever flown and it maxed out at 3 g's for launch and 2 for landing while being half the likely entry mass of a MCT. When I say 2 g's I am making accommodation for the vehicles size rather then just hand-waving away a WELL understood engineering relationship between size and strength and substituting the g-forces experienced by a craft of a completely different scale and claiming it as a legitimate basis for comparison.
Launch g's are experienced in one orientation, and I specified that I'm worried about the 'nose bleed' negative g's that a bi-conic would experience. If your a fan of a 'Super Dragon' type MCT they you can claim the launch g's and aerocapture g's are oriented the same way and the same structures resist both forces and that could get you up to ~4 g's without any additional structure over what's needed for launch. But I would then point out that your a low lift configuration with high ballistic coefficient which would increase g-forces during entry well beyond 4 g's.
This is a key point, flipping the whole vehicle over and applying forces from the opposite direction is huge. It's the difference between designing a normal skyscraper and one that still stands when upside down.
...and heck, even at your sand-bagged 2 gees, that still doesn't rule out 100 day transfers with multi-pass aerocapture/entry at the end. With a lifting entry, you can spread out that acceleration over a long time.
That was also true for each and every Shuttle and Apollo reentry. The reentry corridor was narrow. But if one MCT out of 500 ships did skip out on a heliocentric trajectory, you'd just have to devote one or two stripped-down MCTs in orbit or on the surface of Mars to be ready to mount a rescue mission (or wouldn't even have to be devoted to that purpose, just ready to be fueled up when the MCTs are arriving). In the grand scheme of things, that would be fairly cheap insurance....and heck, even at your sand-bagged 2 gees, that still doesn't rule out 100 day transfers with multi-pass aerocapture/entry at the end. With a lifting entry, you can spread out that acceleration over a long time.
Aerocapture is hard precisely because it must come all at once, lest the craft shoot out on a modified heliocentric orbit; Consequences for unpredictable failure usually involve loss of crew.
'Lifting entry' is primarily useful after capture has occurred, when you have plenty of time-in-atmosphere to work with.No, it's also useful during aerocapture because you can adjust for uncertainties in the atmosphere's density. And you can utilize negative lift to give you more time-in-atmosphere.
100 day transfers don't just involve high amounts of velocity to dissipate, more and more of that velocity has to be dissipated on the first pass.Sure. But not more than needs to be dissipated for lunar reentry.
If you're limited to just 2 gees on reentry, that's a good argument for base-first entry, same load-path.
That was also true for each and every Shuttle and Apollo reentry. The reentry corridor was narrow. But if one MCT out of 500 ships did skip out on a heliocentric trajectory, you'd just have to devote one or two stripped-down MCTs in orbit or on the surface of Mars to be ready to mount a rescue mission (or wouldn't even have to be devoted to that purpose, just ready to be fueled up when the MCTs are arriving). In the grand scheme of things, that would be fairly cheap insurance.
No, it's also useful during aerocapture because you can adjust for uncertainties in the atmosphere's density. And you can utilize negative lift to give you more time-in-atmosphere.
Quote100 day transfers don't just involve high amounts of velocity to dissipate, more and more of that velocity has to be dissipated on the first pass.Sure. But not more than needs to be dissipated for lunar reentry.
If magnetic deceleration can be used it would help a lot with reusability. If a first pass can achieve capture then a second or third pass can achieve landing. Replacing the heatshield after every flight will cost time and money. I don't see magnetic as a requirement but i hope for it. A metallic heatshield will not do especially for EDL back on earth.Even a PICA-X shield can take multiple entries (Musk said 100, which is an overstatement but is a lot more than MCT will need). And yes, a metallic shield would work, you just have to do it gently enough. But I suspect SpaceX is looking at PICA-X. A tanker going to LEO, most certainly, could use a metallic TPS without any hard limits on numbers of reuses (thousands are possible).
A metallic heatshield for reentry from interplanetary speeds? I doubt it. PicaX reusable yes. But from interplanetary speeds? To Mars and back, yes, thats two uses, the second very heavy. But again? Maybe two return flights, maybe not.
Edit: I hope I am wrong.
Metallic heatshield for the tankers, that would help a lot, I agree. Many tank flights without refurbishing the heatshield. But it would mean a completely different design for the tanker. The tanker coming back from LEO can probably make a significant number of flights before refurbishing with PicaX.
Do they wan to have two different heat shields for the tanker version and the Mars version? Or just use Pica-X on both for commonality, and just replace it after may one round trip from Mars, or like 10 [for example] trips to LEO? Probably need to thoroughly go through a tanker after 10 round trips anyway and look for other wear items.
Again like in other discussions we run into the situation that we need to differentiate between different phases. I would assume that initially the tanker too would have PicaX for ease of development. If and when the number of flights really takes off it can be worth it to redesign the tanker to be more reusable without changing the heatshield. Then metallic or even better magnetic would look quite attractive.
why is MCT always considered a single vehicle launched from Earth? Has Musk in any way indicated that?
isn´t there any possibility MCT could be a space only craft in permanent transit between Mars and Earth, picking people on Earth and dropping them on Mars?
that way, MCT could be faster and bigger. Because you wouldn´t need to accelerate an habitat for 100 people to live for 6 months (or less if possible) EVERY trip.
Just accelerate (to catch up with MCT and transfer people) a "small" capsule where you have 100 people crammed in a tight space with their luggage.
The idea of a big transport to Mars for people to live in for 6 months being launched EVERY trip sounds to me somewhat like launching half the space station into space every time a new astronaut goes to it.
why is MCT always considered a single vehicle launched from Earth? Has Musk in any way indicated that?
how shorter does he want the transit times?
from here https://en.wikipedia.org/wiki/Mars_cycler it seems plenty of them are under 3 months?
but anyway, isn´t there a middle term? Cyclers hardly use any fuel at all. What about a cycler which also has fuel assist BUT still would use much less fuel (for it´s size) than launching from Earth?
why is MCT always considered a single vehicle launched from Earth? Has Musk in any way indicated that?
Musk made some comments about "landing the whole thing" and "using MCT as the surface habitat initially." This has lead to people assuming there won't be multiple vehicle that land, launch, and transit separately like Mars Semi-direct or something.
It's not much to go on, and SpaceX's eventual actual MCT concept could have completely changed since those comments. But it's all we've had to go on so far. :-)
While SpaceX clearly has experience with PICA and their own formulations of such, I don't think they would be afraid of pursuing metallic if they felt it made good engineering and cost sense. The biggest factor in favor of metallic TPS is that is it the only way you get to a true 'gas-and-go' vehicle and we know that's ultimately what they want. The hypothetical Tanker-2-LEO vehicle is clearly going to have a rapid launch cadence, much more so then the vehicle which actually transits to and from mars so it's the logical the first place something like metallic TPS would be useful.
The question is given all the other things that will be needed to turn a vehicle around if the process actually takes a week or two (still a huge improvement over Shuttle) dose the replacement of an ablative heat-shield really impact your turn around time significantly, especially if it is replaced only infrequently. I suspect this will be the case, that the best vehicle that can presently be built won't really need metallic TPS for it's turn around time so it won't initially have it. But if the turn around process is improved, perhaps after several years of doing it to the point that heat-shield replacement becomes a bottleneck then I can see this being a desirable upgrade.
I am very sure that we will not see PICAX directly applied to propellant tankage though as some have speculated, the bonding of a brittle material to a tank wall that will become cryogenic strikes me as unfeasible as ablatives are attached with adhesives which are notoriously bad when cold, nor dose it offer any means of repair or replacement as ablation takes place. You need a structural frame that holds the ablative and allows it to be removed, this would also be the default means of using metallic TPS in large panels, though I can conceive of a monocoque design in which the a multi-layer sandwiching of metals creates the tank wall, the insulation layer and the outer skin.
why is MCT always considered a single vehicle launched from Earth? Has Musk in any way indicated that?
isn´t there any possibility MCT could be a space only craft in permanent transit between Mars and Earth, picking people on Earth and dropping them on Mars?
that way, MCT could be faster and bigger. Because you wouldn´t need to accelerate an habitat for 100 people to live for 6 months (or less if possible) EVERY trip.
Just accelerate (to catch up with MCT and transfer people) a "small" capsule where you have 100 people crammed in a tight space with their luggage.
The idea of a big transport to Mars for people to live in for 6 months being launched EVERY trip sounds to me somewhat like launching half the space station into space every time a new astronaut goes to it.
I'm wilderness bound. MUsk's statement was a while ago, and things definitely change. I think its going to be a lot more radical than people have been speculating so far.why is MCT always considered a single vehicle launched from Earth? Has Musk in any way indicated that?
isn´t there any possibility MCT could be a space only craft in permanent transit between Mars and Earth, picking people on Earth and dropping them on Mars?
that way, MCT could be faster and bigger. Because you wouldn´t need to accelerate an habitat for 100 people to live for 6 months (or less if possible) EVERY trip.
Just accelerate (to catch up with MCT and transfer people) a "small" capsule where you have 100 people crammed in a tight space with their luggage.
The idea of a big transport to Mars for people to live in for 6 months being launched EVERY trip sounds to me somewhat like launching half the space station into space every time a new astronaut goes to it.
Not, always, I'm one of a few voices in the wilderness that still see semi-direct like architectures as most likely and fully compatible with Musk's goals and statements so far. And we should NOT be taking even clear statements of intent as evidence that alternatives have been taken off the table as their are FAR too many instances of early statements of intent in the engineering minutia by SpaceX that fall through.
I'm wilderness bound. MUsk's statement was a while ago, and things definitely change. I think its going to be a lot more radical than people have been speculating so far.
Things do change and I can't wait for the great reveal. I like MCT going all the way and back for a simple reason, not only because Elon Musk said so. It is operationally simple, straightforward and elegant. Making MCT ready for the next flight is a lot simpler down on earth than in space. Getting the weight back up into orbit is not an expensive problem given the capability of BFR.
...Thank you Captain Obvious.
That was also true for each and every Shuttle and Apollo reentry. The reentry corridor was narrow. But if one MCT out of 500 ships did skip out on a heliocentric trajectory, you'd just have to devote one or two stripped-down MCTs in orbit or on the surface of Mars to be ready to mount a rescue mission (or wouldn't even have to be devoted to that purpose, just ready to be fueled up when the MCTs are arriving). In the grand scheme of things, that would be fairly cheap insurance.
No, Shuttle did not do Aerocapture because it was not going anywhere near escape velocity on atmospheric entry and could not leave the Earth if it over-shot
it did controlled skips and banked turns along the atmosphere to dissipate energy but that is not the same as aerocapture.Nope, but the reentry corridor is quite narrow in order to keep the reusable and fragile TPS intact, which was my point: narrow reentry corridor.
When you under-shoot aerocapture you end up back on a Heliocentric orbit leaving the planet entirely.Of course. If you screw up the Shuttle reentry corridor, you don't just end up on a heliocentric orbit, you end up completely dead. I'd rather be on a heliocentric trajectory, waiting on a rescue mission, than completely dead. But the narrow reentry corridor didn't kill anyone on either Shuttle or Apollo, and we have more sophisticated guidance and navigation than we did then.
The rescue idea is ridiculous, the lost vehicle would be heading for a aphelion well into the asteroid belt and even if another MCT on mars had the DeltaV to rendezvous with it they would have not have the DeltaV to return to mars or likely any other planet on a time scale of less then years.Um, yes it would. A stripped down MCT in elliptical Mars orbit could have capability for FAR more C3 than the MCT that didn't quite aerocapture enough. It could launch almost immediately after the failed aerocapture, easily catching up to the MCT in question. You aren't stuck waiting for orbits to come back around because you would have plenty of delta-v to work with. On the scale that SpaceX is dealing with, this sort of capability would be cheap. It's not "ridiculous" if this is the only reason keeping you from attempting aerocapture.
Sure, but even a typical capsule shape can generate significant lift. And with greater entry precision, that amount of lift may be all you really need (more later).No, it's also useful during aerocapture because you can adjust for uncertainties in the atmosphere's density. And you can utilize negative lift to give you more time-in-atmosphere.
Yes this is correct, the more lift you can generate the wider your aerocapture corridor is and the more adjustment can be made in real time due to atmospheric conditions. It is an argument in favor of the bi-conic shape when doing aerocapture.
I pretty clearly was talking about velocity difference (i.e. acceleration times time), not energy.QuoteQuote100 day transfers don't just involve high amounts of velocity to dissipate, more and more of that velocity has to be dissipated on the first pass.Sure. But not more than needs to be dissipated for lunar reentry.
Apollo entry was at 10.77 km/s, contrary to popular belief they could not skip off the atmosphere and be lost in space forever, they would simply have re-entered later after a large elliptical orbit during which they would run out of oxygen because they were below escape velocity.
Burnate already did the numbers, on the 102 day transit your perigee at mars is 9.17 km/s and you need to shed 4.7 km/s. That's about half your velocity, and about 50% of the energy of the Apollo return.
The 100 day transit is even worse (it saves Earth Escape DeltaV by raising velocity at mars), cause your velocity is 13.77 km/s of which you must shed 9.2 km/s, which is 88% of your total energy. In fact you need to shed more energy then the entire Apollo lunar return, about 60% more.
I'm very skeptical that the cord through the martian atmosphere can be long enough to provide this much deceleration within even your g limits. The faster your going to less time you spend in the atmosphere and your deceleration force needs to increase exponentially. And as entry velocities increase the corridor gets narrower and narrower, according to http://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19930003532.pdf even with 5 g limit the entry velocity had to kept to under 9 km/s to keep the corridor to 1 degree wide with a 0.3 L/D vehicle like a capsule.So we'll have to get within less than a degree wide. I don't see the problem; we're much better at Mars entry precision than we were when that document was written (1992), particularly with the Curiosity rover's entry, and clearly some sort of areospatial positioning system could be deployed to enhance the precision to within a tiny fraction of a degree (on Earth, you can get within centimeters). But even with purely optical aids, you can do much better than a single degree: "Other investigators have shown that with the use of onboard optical sightings of the Martian moons, the error in the atmospheric entry angle could be reduced to +/-0.25 degrees without the need for secondary spacecraft to serve as navigational aides (Ref. 45)." (In fact, we already have spacecraft around Mars that could be used for this purpose.)
why is MCT always considered a single vehicle launched from Earth? Has Musk in any way indicated that?
isn´t there any possibility MCT could be a space only craft in permanent transit between Mars and Earth, picking people on Earth and dropping them on Mars?
that way, MCT could be faster and bigger. Because you wouldn´t need to accelerate an habitat for 100 people to live for 6 months (or less if possible) EVERY trip.
Just accelerate (to catch up with MCT and transfer people) a "small" capsule where you have 100 people crammed in a tight space with their luggage.
The idea of a big transport to Mars for people to live in for 6 months being launched EVERY trip sounds to me somewhat like launching half the space station into space every time a new astronaut goes to it.
Google "The MCV: A Mars Colonization Vehicle". Two kids at the Mars Society (Musk was there too!). Slightly different concept, but enough like Aldrin's cycler in ways that he was verbally jousting with them.why is MCT always considered a single vehicle launched from Earth? Has Musk in any way indicated that?
isn´t there any possibility MCT could be a space only craft in permanent transit between Mars and Earth, picking people on Earth and dropping them on Mars?
that way, MCT could be faster and bigger. Because you wouldn´t need to accelerate an habitat for 100 people to live for 6 months (or less if possible) EVERY trip.
Just accelerate (to catch up with MCT and transfer people) a "small" capsule where you have 100 people crammed in a tight space with their luggage.
The idea of a big transport to Mars for people to live in for 6 months being launched EVERY trip sounds to me somewhat like launching half the space station into space every time a new astronaut goes to it.
Because cyclers are inherently a ... dumb idea. IMO. You've got to accelerate all the crew, crew consumables, cargo, and propellant for the lander to a Mars transfer orbit anyway. Just to catch up to the cycler. All while stuck in VERY cramped conditions. This has some very serious failure modes - what if you cannot rendezvous with the cycler?
Do the math, look at the size of the "capsule" you need to transport all that to the cycler, and you'll realize that your small transfer vessel is almost at the size you need for normal transit, so you might as well just scale up the crew area and skip the cycler altogether. (And that larger crew area will be needed on Mars as well)
And that doesn't even beging to deal with the limited and few launch windows a cycler offers. And how you deal with a fleet of ships going in the same launch window.
No, I maintain that "cyclers" make no sense. And I'm dumbfounded why otherwise brilliant people such as Aldrin keep advocating for them.
Has anyone considered the possibility that the upcoming reveal will just be the Rocket, aka BFR, and not in fact the whole system? I really suspect that SpaceX will be doing just BFR development over the next decade while the actual vehicle to mars will continue to be studied and refined without a firm mission architecture until they actually have to cross that bridge.Possibly but it's also possible that reusable second stage and Mars ambitions combine to make BFR/MCT the next step.
SpaceX has always been a builder of spacecraft as well as a builder of launch vehicles.
I have a hard time picturing a payload other than MCT for the first BFR test flight.
SpaceX has always been a builder of spacecraft as well as a builder of launch vehicles.
That's true. We knew about F9 and Dragon pretty much at the same time when both were in development. Actually, didn't dragon development start before F9 did?
I don't imagine this will be prescient. BFR will either come at the same time as the MTC or slightly before. One prescient I do believe in: there won't be a BFR Falcon V analogue, they'll go full monty, and just iteratively upgrade the full size rocket/space craft over time like they're oft to do with F9.
Has anyone considered the possibility that the upcoming reveal will just be the Rocket, aka BFR, and not in fact the whole system? I really suspect that SpaceX will be doing just BFR development over the next decade while the actual vehicle to mars will continue to be studied and refined without a firm mission architecture until they actually have to cross that bridge.
Has anyone considered the possibility that the upcoming reveal will just be the Rocket, aka BFR, and not in fact the whole system? I really suspect that SpaceX will be doing just BFR development over the next decade while the actual vehicle to mars will continue to be studied and refined without a firm mission architecture until they actually have to cross that bridge.
Because cyclers are inherently a ... dumb idea. IMO. You've got to accelerate all the crew, crew consumables, cargo, and propellant for the lander to a Mars transfer orbit anyway. Just to catch up to the cycler. All while stuck in VERY cramped conditions. This has some very serious failure modes - what if you cannot rendezvous with the cycler?
Do the math, look at the size of the "capsule" you need to transport all that to the cycler, and you'll realize that your small transfer vessel is almost at the size you need for normal transit, so you might as well just scale up the crew area and skip the cycler altogether. (And that larger crew area will be needed on Mars as well)
And that doesn't even beging to deal with the limited and few launch windows a cycler offers. And how you deal with a fleet of ships going in the same launch window.
No, I maintain that "cyclers" make no sense. And I'm dumbfounded why otherwise brilliant people such as Aldrin keep advocating for them.
Has anyone considered the possibility that the upcoming reveal will just be the Rocket, aka BFR, and not in fact the whole system? I really suspect that SpaceX will be doing just BFR development over the next decade while the actual vehicle to mars will continue to be studied and refined without a firm mission architecture until they actually have to cross that bridge.Possibly but it's also possible that reusable second stage and Mars ambitions combine to make BFR/MCT the next step.
SpaceX has always been a builder of spacecraft as well as a builder of launch vehicles.
Has anyone considered the possibility that the upcoming reveal will just be the Rocket, aka BFR, and not in fact the whole system? I really suspect that SpaceX will be doing just BFR development over the next decade while the actual vehicle to mars will continue to be studied and refined without a firm mission architecture until they actually have to cross that bridge.Possibly but it's also possible that reusable second stage and Mars ambitions combine to make BFR/MCT the next step.
SpaceX has always been a builder of spacecraft as well as a builder of launch vehicles.
Anything is possible. We won't know until they see.
But, "BFR" is a term I think only we've used here on the forums. I don't think SpaceX or Musk has ever differentiated a HLV and MCT as separate items. It's always been 'MCT' as I recall seeing.
So I see it as being more like "STS" than "Saturn V" and "Apollo" or "Falcon" and "Dragon" or even "Energia" and "Buran"...as separate systems.
The orbiter was integral to STS, not optional. You couldn't really unveil the launch vehicle without the orbiter too.
As always, I might be completely misreading EM's intent. heh. ;-)
I'm pretty sure the term BFR has at least, been used by SpaceX if they didn't coin the term.
I'm pretty sure the term BFR has at least, been used by SpaceX if they didn't coin the term.
Yes EM's said it. I think it was something like "you need a really big f'ing rocket, a BFR". But if that memory is correct, then he wasn't really calling it a BFR at the time... just describing it.
Since then some publications have called it a "Big Falcon Rocket". Maybe BFR will stick.
I really suspect that SpaceX will be doing just BFR development over the next decade while the actual vehicle to mars will continue to be studied and refined without a firm mission architecture until they actually have to cross that bridge.
Things do change and I can't wait for the great reveal. I like MCT going all the way and back for a simple reason, not only because Elon Musk said so. It is operationally simple, straightforward and elegant. Making MCT ready for the next flight is a lot simpler down on earth than in space. Getting the weight back up into orbit is not an expensive problem given the capability of BFR.
I agree, with a proviso:
As you add more and more little things to make it work, a straightforward plan can have terrible complexities. Many refueling dockings, transferring people from several Dragons to the MCT, etc... there will be a point where making an easy system work isn't easy.
But otherwise yes... if they can make it work in the simple way envisaged it will make the trip cheaper (which is ultimately the SpaceX goal). And 20 Synods later they'll have a different vision for higher numbers.
Has anyone considered the possibility that the upcoming reveal will just be the Rocket, aka BFR, and not in fact the whole system? I really suspect that SpaceX will be doing just BFR development over the next decade while the actual vehicle to mars will continue to be studied and refined without a firm mission architecture until they actually have to cross that bridge.
That is precisely my thinking.
The BFR is congruent with SX's core competency and is a necessary precursor. Even it will undergo changes after hopefully SX succeeds in recovering and re-flying F9 cores.
I think Musk will speak to the MCT issue, but still in vague terms. Even were he somewhat specific I don't think many observers here would expect the details to hold true during what I expect would be a longer than a decade gestation for the Block One MCT, which I expect to be quite different from its folllow-on.
There are too many outside SX technology developments ongoing* to fixate and spend R&D $ on a MCT in this decade, and probably for the early part of the next.
*SEP panels & engines, plasma propulsion, magneto aero capture, life support systems tech, etc.
Never mind extreme wildcards like one of the myriad small fusion projects actually working.
You seem to forget that Dragon is clearly also SpaceX's core competency. Dragon is at least as impressive as F9 is. And what will the second stage of BFR be? MCT and BFR's second stage are solving almost all the same problems. The ones that are left are ones that Dragon is helping to solve.Has anyone considered the possibility that the upcoming reveal will just be the Rocket, aka BFR, and not in fact the whole system? I really suspect that SpaceX will be doing just BFR development over the next decade while the actual vehicle to mars will continue to be studied and refined without a firm mission architecture until they actually have to cross that bridge.
That is precisely my thinking.
The BFR is congruent with SX's core competency and is a necessary precursor. ...
...Um, SpaceX is absolutely working SEP panels/engines/plasmapropulsion and life support systems tech right now. Magneto aerocapture is a totally unproven idea that is only really being pursued by one company (with nearly no funding), and it will remain to be seen if it's either practical or worthwhile. Regardless, evangelists for the tech (mainly Jon Goff here ;) ) seem to emphasize its ability to allow /existing/ tech to utilize it, so SpaceX need not put off developing MCT even if it's a totally real thing. And I would put even less faith in the small fusion projects.
There are too many outside SX technology developments ongoing* to fixate and spend R&D $ on a MCT in this decade, and probably for the early part of the next.
*SEP panels & engines, plasma propulsion, magneto aero capture, life support systems tech, etc.
Never mind extreme wildcards like one of the myriad small fusion projects actually working.
What Musk says isn't "holy writ." You don't have to like it, but if you're going to discuss MCT, you have to actually discuss /MCT/ (which means what little we have to go on), not make up your favorite architecture in its place. You can do that in the Mars subsection of the more general forum.Has anyone considered the possibility that the upcoming reveal will just be the Rocket, aka BFR, and not in fact the whole system? I really suspect that SpaceX will be doing just BFR development over the next decade while the actual vehicle to mars will continue to be studied and refined without a firm mission architecture until they actually have to cross that bridge.
That is precisely my thinking.
The BFR is congruent with SX's core competency and is a necessary precursor. Even it will undergo changes after hopefully SX succeeds in recovering and re-flying F9 cores.
I think Musk will speak to the MCT issue, but still in vague terms. Even were he somewhat specific I don't think many observers here would expect the details to hold true during what I expect would be a longer than a decade gestation for the Block One MCT, which I expect to be quite different from its folllow-on.
There are too many outside SX technology developments ongoing* to fixate and spend R&D $ on a MCT in this decade, and probably for the early part of the next.
*SEP panels & engines, plasma propulsion, magneto aero capture, life support systems tech, etc.
Never mind extreme wildcards like one of the myriad small fusion projects actually working.
I sure we will get some nice video of a giant rocket taking off. It might just have some totally ambiguous payload-fairing on it leaving everything beyond the launch unspecified much the way Vulcan was introduced, this would be a clear sign that they are shopping BFR around to commercial customers, but even I would find that a bit conservative given the hype that has been built up.
More likely were shown some kind of large mars vehicle reaching LEO on top of the BFR. But then the video might just cut to said vehicle landing on mars with all the details of refueling, trajectory, aerocapture, entry etc etc left out. Or on the opposite end of the spectrum we could see a nearly full 'architecture' which covers all these steps and Musk will be open with the numbers.
But even the most minimal video will give some ideas as to what the present preferred architecture is and one of the three camps of 'Super Dragon', 'Integrated Bi-conic' and 'Separate Bi-conic' will likely get to crow, despite the fact that the shown architecture is still in flux, given the kind of 'Holy writ' that people ascribe to offhand statements of Musk their will be no point arguing with a video.
The speculation would then narrow down to other operational details of how the vehicle actually dose the mission and what the ultimate cost will be, how fast if at all SpaceX can put the vehicle into service, if SLS is doomed and when first landing on mars might happen.
Yes EM's said it. I think it was something like "you need a really big f'ing rocket, a BFR". But if that memory is correct, then he wasn't really calling it a BFR at the time... just describing it.
If that was the only mention of it, then the context still doesn't seem to indicate differentiation between a launch vehicle and a spacecraft f he was saying to land 100mt on Mars, you need a really big f'ing rocket...for example...there's nothing in that to indicate a stand alone rocket, and a separate spacecraft.
I'm more saying, was there a reference to "BFR" and "MCT" being referred to separately in the same comment, like how "Dragon" and "Falcon" are referred to differently in the same comment. If so, then there could be a reasonable chance EM will just be announcing the LV. Although even then, it's main purpose would be to loft MCT to space, just like the main purpose of the Saturn V was to loft the Apollo CSM and LEM through TLI. The spacecraft were in the "unveilings" from the beginning. Even when it was originally Direct Ascent instead of LOR.
So I would expect something like that for this unveiling. At least a working concept for the spacecraft, even if there is a stand alone LV.
But good to know that EM actually did use the term. I'd not been aware of that. :-)
This might be useful LoboWe have a thread that is just for information like this, where you'll find more information. Please put that in here if it isn't already: http://forum.nasaspaceflight.com/index.php?topic=37839.0
[Question about getting to Mars.] I don't think the Moon is a necessary step, but I think if you've got a rocket and spacecraft capable of going to Mars, you might as well go to the Moon as well - it's along the way. That's like crossing the English Channel, relative to Mars. So, it's like, if you have these ships that could cross the Atlantic, would you cross the English Channel? Probably. It's definitely not necessary, but you'd probably end up having a Moon base just because, like, why not, ya know. It terms of the key technologies, obviously it would be great to have some sort of fundamental new thing that's never existed before and pushes the boundaries of physics, that'd be great, but as far as the physics that we know today, I actually think we've got the basic ingredients - they're there. I mean, if you do a densified liquid methalox rocket with on-orbit refueling, so like you load the spacecraft into orbit and then you send a whole bunch of refueling missions to fill up the tanks and you have the Mars colonial fleet - essentially - that gets built up during the time between Earth-Mars synchronizations, which occur every 26 months, then the fleet all departs at the optimal transfer point. I think we have - we don't need any sort of thing that people don't already know about, I believe. I believe we've got the building blocks, but the mass efficiency is extremely important. So, having better heat shields, that obviously are reusable.
http://shitelonsays.com/transcript/elon-musk-at-mits-aeroastro-centennial-part-2-of-6-2014-10-24
and this from 2005:
In past talks Musk has hinted at the development of something called the “BFR” (where B stands for “big” and R for “rocket”), a heavy-lift vehicle far larger than the Falcon family of vehicles. At SpaceVision2005 Musk disclosed that the BFR, in its current iteration, would use “multiple” Merlin 2 engines. The BFR would be able to place 100 tons in low Earth orbit, putting it in competition with NASA’s planned shuttle-derived heavy-lift launcher. The BFR is so big, Musk said, that it’s too large for the BFTS at their Texas test site...
http://www.thespacereview.com/article/497/1
Has anyone considered the possibility that the upcoming reveal will just be the Rocket, aka BFR, and not in fact the whole system? I really suspect that SpaceX will be doing just BFR development over the next decade while the actual vehicle to mars will continue to be studied and refined without a firm mission architecture until they actually have to cross that bridge.
Lobo is on the right track here. Before jumping to design solutions, we need to find the requirements for BFR and MCT first. Thats like engineering 101. That hasn't been happening to the extend needed to define an MCT or BFR. Lobo just started that process and I am very glad he did it.
Lobo is on the right track here. Before jumping to design solutions, we need to find the requirements for BFR and MCT first. Thats like engineering 101. That hasn't been happening to the extend needed to define an MCT or BFR. Lobo just started that process and I am very glad he did it.Agree that's necessary - but I believe he and a few others with the appropriate knowledge have been getting into those requirements for quite some time - and from that trying to ascertain what could work that still fits within what EM has said.
Lobo is on the right track here. Before jumping to design solutions, we need to find the requirements for BFR and MCT first. Thats like engineering 101. That hasn't been happening to the extend needed to define an MCT or BFR. Lobo just started that process and I am very glad he did it.Agree that's necessary - but I believe he and a few others with the appropriate knowledge have been getting into those requirements for quite some time - and from that trying to ascertain what could work that still fits within what EM has said.
(Or did I miss some new approach?)
Sleeping your way to mars.
http://www.parabolicarc.com/2015/10/23/snooze-mars/
Musk has worked out just enough about potential "MCT" approaches that he has a generic requirement for his BFR launcher. I am certain that he has explored the solution space for Mars transport and wants a launch capability that can support solutions from all chemical solution to SEP/chemical hybrid transport and other possible more exotic solutions. Technology over a >10 year timeframe does not stay still. The 1st and 2nd stage need only be able to get the 100mT cargo plus MCT dry weight into appropriate LEO. If all his broad brush Mars transport solution space fits within these parameters and their evolutionary improvements (e.g.Merlin to Merlin FT), he's good to go.But why build BFR first? BFR has the most infrastructure requirements. Additionally, you're making the implicit assumption that MCT isn't essentially BFR's second stage. I really, REALLY don't expect SpaceX to make the same mistake NASA is currently making by building a super-expensive-to-develop-and-maintain launch vehicle without really anything to launch.
... So, the Block One MCT will likely again be different from the MCT that lands the first crew on Mars June 2033 following the un-crewed MCT bringing the ISRU equipment May 2031.I have no doubt MCT will evolve, but your timeline is not the same as SpaceX's timeline. They expect crewed missions much earlier. Which makes sense, as it doesn't make sense to develop a capability and then essentially just let it languish, sucking up money while nothing is accomplished (another mistake NASA is making, though this is mostly Congress's fault).
Musk has worked out just enough about potential "MCT" approaches that he has a generic requirement for his BFR launcher. I am certain that he has explored the solution space for Mars transport and wants a launch capability that can support solutions from all chemical solution to SEP/chemical hybrid transport and other possible more exotic solutions. Technology over a >10 year timeframe does not stay still. The 1st and 2nd stage need only be able to get the 100mT cargo plus MCT dry weight into appropriate LEO. If all his broad brush Mars transport solution space fits within these parameters and their evolutionary improvements (e.g.Merlin to Merlin FT), he's good to go.But why build BFR first? BFR has the most infrastructure requirements. Additionally, you're making the implicit assumption that MCT isn't essentially BFR's second stage. I really, REALLY don't expect SpaceX to make the same mistake NASA is currently making by building a super-expensive-to-develop-and-maintain launch vehicle without really anything to launch.
philw1776: No, I see the MCT as the BFR stage 2. Just its specifics beyond engines & tankage at TBD. I see several cargo refueler MCTs as does Musk to fuel up the transit MCT.Quote... So, the Block One MCT will likely again be different from the MCT that lands the first crew on Mars June 2033 following the un-crewed MCT bringing the ISRU equipment May 2031.I have no doubt MCT will evolve, but your timeline is not the same as SpaceX's timeline. They expect crewed missions much earlier. Which makes sense, as it doesn't make sense to develop a capability and then essentially just let it languish, sucking up money while nothing is accomplished (another mistake NASA is making, though this is mostly Congress's fault).
Musk has worked out just enough about potential "MCT" approaches that he has a generic requirement for his BFR launcher. I am certain that he has explored the solution space for Mars transport and wants a launch capability that can support solutions from all chemical solution to SEP/chemical hybrid transport and other possible more exotic solutions. Technology over a >10 year timeframe does not stay still. The 1st and 2nd stage need only be able to get the 100mT cargo plus MCT dry weight into appropriate LEO. If all his broad brush Mars transport solution space fits within these parameters and their evolutionary improvements (e.g.Merlin to Merlin FT), he's good to go.But why build BFR first? BFR has the most infrastructure requirements. Additionally, you're making the implicit assumption that MCT isn't essentially BFR's second stage. I really, REALLY don't expect SpaceX to make the same mistake NASA is currently making by building a super-expensive-to-develop-and-maintain launch vehicle without really anything to launch.
Quote... So, the Block One MCT will likely again be different from the MCT that lands the first crew on Mars June 2033 following the un-crewed MCT bringing the ISRU equipment May 2031.I have no doubt MCT will evolve, but your timeline is not the same as SpaceX's timeline. They expect crewed missions much earlier. Which makes sense, as it doesn't make sense to develop a capability and then essentially just let it languish, sucking up money while nothing is accomplished (another mistake NASA is making, though this is mostly Congress's fault).
... It is also unfeasible on technical grounds but no one seems willing to admit this.
You can't have it both ways, either SpaceX is a lean and efficient vertically integrated company that's vastly more efficient then 'Old-Space' capable of making a rocket without a huge standing army or they aren't. If BFR takes huge infrastructure to build and maintain like SLS dose then the prospect of getting to mars is already busted.
I've said a dozen times I expect SpaceX to sell normal commercial flights on BFR for a good long time before it is use for mars, that will build precious flight history before people fly on it. Two or three large com satellites at a time, Bigelow type space stations, the LEO satellite swarms, and any NASA missions that would have flown on SLS are all potential things to fill the rocket with. And because these things require a normal 2nd stage that is the configuration that makes sense, it allows a product that can fend for itself in the marketplace and pay for it's own infrastructure without putting that whole cost onto the back of the mars ticket. Also it doesn't risk completely bankrupting the company in getting to mars turns out to be harder then expected to develop or customers are not ready to buy when the ride is offered.
As NASA is the only conceivable customer for a first mission they need to be courted to create a mission utilizing SpaceX as the primary contractor. But NASA will need a decade or more to develop and ready the ground systems. But to fund that SLS needs to be canceled first and that not going to come until a proven alternative has been fully demonstrated to be reliable.
Were already seeing spiral development from SpaceX when they layed down considerable time and money for Raptor development which was clearly started without the full mars architecture in place, and probably no more then a vague sense of how large BFR would be either. They just set out to the make the best engine possible and then build the best launch vehicle with and around it.
Which is exactly why an integrated 2nd stage is a bad idea, it makes the launcher too narrow and unable to serve any other role efficiently.
It is also unfeasible on technical grounds but no one seems willing to admit this.
Which is exactly why an integrated 2nd stage is a bad idea, it makes the launcher too narrow and unable to serve any other role efficiently. It is also unfeasible on technical grounds but no one seems willing to admit this.
People have been collectively running math here for years and there's nothing theoretically showstopping about it. Yes, it's tricky engineering, but it's not physics violating engineering. This isn't so much of a quantum leap as the R7 was from the V2.
People have been collectively running math here for years and there's nothing theoretically showstopping about it. Yes, it's tricky engineering, but it's not physics violating engineering. This isn't so much of a quantum leap as the R7 was from the V2.
Designing MCT for Mars EDL will be tricky and in the end they may fail on that. Success is not guaranteed. But I am optimistic about it. As you said, there are years of engineering behind it by top designers. It is not like Elon Musk was on it alone and it is all pipe dreams.
Musk has worked out just enough about potential "MCT" approaches that he has a generic requirement for his BFR launcher. I am certain that he has explored the solution space for Mars transport and wants a launch capability that can support solutions from all chemical solution to SEP/chemical hybrid transport and other possible more exotic solutions. Technology over a >10 year timeframe does not stay still. The 1st and 2nd stage need only be able to get the 100mT cargo plus MCT dry weight into appropriate LEO. If all his broad brush Mars transport solution space fits within these parameters and their evolutionary improvements (e.g.Merlin to Merlin FT), he's good to go.But why build BFR first? BFR has the most infrastructure requirements. Additionally, you're making the implicit assumption that MCT isn't essentially BFR's second stage. I really, REALLY don't expect SpaceX to make the same mistake NASA is currently making by building a super-expensive-to-develop-and-maintain launch vehicle without really anything to launch.
You can't have it both ways, either SpaceX is a lean and efficient vertically integrated company that's vastly more efficient then 'Old-Space' capable of making a rocket without a huge standing army or they aren't. If BFR takes huge infrastructure to build and maintain like SLS dose then the prospect of getting to mars is already busted.
Year | Element | Phase | Notes |
2014 | Raptor | initial investigations | |
Q4 2015 | Architecture | release | |
Q1 2016 | Raptor | PDR | |
2016-7 | BFR+MCT | initial designs | several iterations towards a PDR in 2018 |
Q4 2017 | Raptor | CDR | |
Q1 2018 | BFR | PDR | depends on Raptor PDR |
2018 | Red dragon | site survey | surface element demos |
2019 | MCT | PDR | depends on Raptor PDR |
2020 | BFR | CDR | |
2020 | Red dragon(s) | site survey(s) | surface element demos & subsequent synods |
2022 | MCT | CDR | tanker and cargo versions |
2023 | MCT | CDR | crew versions |
2023 | BFR | First flight | 1st stage only |
2024 | MCT | Demo flight | bare bones tanker version |
2025 | BFR | Demo flight(s) | cargo version & refueling |
2026 | BFR | Demo flight(s) | crew version - LEO (unmanned) |
2027-31 | BFR | Demo flight(s) | crew version - LEO long duration (manned) |
2028 | BFR | Demo flight(s) | cargo version - Mars + crew version - cis-lunar |
2031 | BFR | Cargo flight(s) | cargo version - Mars |
2033 | BFR | Crew flight(s) | crew+cargo versions - Mars |
First the idea of a high maintenance cost is inconsistent with SpaceX operations, they shown the ability to build and launch with much less manpower. It also ignores the intended re-usable nature of BFR, they do not need to manufacture them at a rate of one a month like Falcon, though they will likely make Raptor engines at a few a week with a comparable labor input and assembly-line like production like Merlin manufacturing.SpaceX will lose a huge advantage in per-launch labor efficiency if they only build, say, 1 BFR a year. That is exactly why SLS is ridiculously expensive. SpaceX may beat the industry-standard labor efficiency for rocket manufacture and development by 2-3x, MAYBE, but their main advantage is picking reasonable solutions, i.e. not building a ridiculously expensive rocket using technology half a century old and that will only launch a couple times per year at most.
RB is arguing that the whole top to bottom mars transportation system has to come into being simultaneously.Ignoring for the moment that you're putting words in my mouth: You mean with Mars transfer vehicle, a lander/ascent vehicle, SEP propulsion, depots and everything? Am I arguing that? No. I'm arguing that a Mars transport system of BFR and MCT as the second stage is basically the same thing as a TSTO RLV BFR.
...The safe strategy ...Yeah, no, there's nothing safe about BFR or MCT or SpaceX's Mars plans. If they were merely playing safe, they would still be at Falcon 1 or maybe a Falcon 5. BFR/MCT is not at all a safe strategy.
The only thing that might make sense for an HLV is if you had some good rationale for flying enough payloads per year that you could take FULL advantage of a full manufacturing line (multiple shifts) /and/ full advantage of full reuse (both first and second stages). To keep a production line busy, you probably need at least 10 flights per year. To do that with multiple shifts, probably about 40 first stage cores per year (and more upper stages). To make first stage reuse make sense, you need to reuse the stage at least ten times. To make full reuse make sense, you need to reuse the first stage at least 100 times and the upper stage 10-30 times. That's roughly a thousand launches per year at a minimum. Partial reuse can get by with maybe 100 launches per year, 40 per year if some are expendable. Otherwise, it's not really the economic optimum. Unless you get to 1000 launches per year, though, you probably aren't launching enough to gain anything by using an HLV. (In fact, you'd reduce your economic efficiency since you wouldn't be reusing as much.)
This is my guess at a MCT plan. Note I've separated out the tanker, cargo and crew variants of MCT although they are all based on the same design.
Musk mentioned like 80,000 people per year. He didn't specify synod. So that's almost another whole order of magnitude.Really? When was that?
Millions of people needed for Mars colony, so 80k+ would just be the number moving to Mars per year
Also, you're missing a really important point: There isn't actually a market for HLV launch. Nobody* wants more than about EELV Heavy capability (except for exploration, but even then it's not required). There aren't even commercial or military processing facilities for payloads over 5 meters in diameter. There isn't a commercial launch need for BFR, and NASA doesn't even have any payloads that need an HLV. F9 and Falcon Heavy are MORE than enough for all the commercial, civil, and military payloads. I've held this viewpoint for years, I've been vocal about it, and I haven't actually changed my mind on it. The only thing that might make sense for an HLV is if you had some good rationale for flying enough payloads per year that you could take FULL advantage of a full manufacturing line (multiple shifts) /and/ full advantage of full reuse (both first and second stages). To keep a production line busy, you probably need at least 10 flights per year. To do that with multiple shifts, probably about 40 first stage cores per year (and more upper stages). To make first stage reuse make sense, you need to reuse the stage at least ten times. To make full reuse make sense, you need to reuse the first stage at least 100 times and the upper stage 10-30 times. That's roughly a thousand launches per year at a minimum. Partial reuse can get by with maybe 100 launches per year, 40 per year if some are expendable. Otherwise, it's not really the economic optimum. Unless you get to 1000 launches per year, though, you probably aren't launching enough to gain anything by using an HLV. (In fact, you'd reduce your economic efficiency since you wouldn't be reusing as much.)
https://twitter.com/elonmusk/status/273483420468932608Thanks for that.Quote from: @elonmuskMillions of people needed for Mars colony, so 80k+ would just be the number moving to Mars per year
If NASA is to be the first customer the concept is dead in thewaterspace. It would introduce decades of delay as you yourself state. It is not going to happen.
It is also unfeasible on technical grounds but no one seems willing to admit this.
How do you come to that conclusion?
The system must be viable at a lower initial volume so it can survive to ramp up over decades.
No that is not at all reasonable, they will put money generated in other areas into development but they would not RUN the whole system at a loss for decades waiting for volume to grow. That would suck up all their surplus from other activities leaving nothing for continued development and upgrading the system, you make a profitable system and then make it more profitable with your profits. This is completely counter to the entire history of careful and financially cautious development done by SpaceX up till now.
Like the old adage "Every complex program that has ever worked has evolved from a simple program that worked, not a complex program that didn't work".
Will BFR/MCT do a full mission - to Mars - refuel - back to Earth unmanned before the first crew? Elon has mentioned Droids and automation as being essential elements.
BTW I was surprised that Elon Musk mentioned they would land humans only when the return fuel is waiting for them. I would have expected to land the equipment but set it up for fuel production only after the human landing.
BTW I was surprised that Elon Musk mentioned they would land humans only when the return fuel is waiting for them. I would have expected to land the equipment but set it up for fuel production only after the human landing.
It makes perfect sense, otherwise you could literally doom/maroon the first crew. Remember that the MCT architecture (as we know it) is completely dependent on manufacturing ALL propellant for the journey home.
The first people to land are not going to be permanent colonist.
Musk is borrowing somewhat from Mars Direct (or is it Semi-Direct?) where an already-fueled ascent vehicle is fueled up on the surface. I don't see a good reason not to have a fueled up vehicle ready when they arrive.BTW I was surprised that Elon Musk mentioned they would land humans only when the return fuel is waiting for them. I would have expected to land the equipment but set it up for fuel production only after the human landing.
It makes perfect sense, otherwise you could literally doom/maroon the first crew. Remember that the MCT architecture (as we know it) is completely dependent on manufacturing ALL propellant for the journey home.
The first people to land are not going to be permanent colonist.
Not doom. They would not be colonists but they will IMO likely stay for a full synod until a replacement crew arrives. Probably a few stay on to extend their experience to the next crew. Worst case something goes wrong and they stay two synods. On the next synod they would receive spare parts to get fuel ISRU going.
BTW I was surprised that Elon Musk mentioned they would land humans only when the return fuel is waiting for them. I would have expected to land the equipment but set it up for fuel production only after the human landing.
It makes perfect sense, otherwise you could literally doom/maroon the first crew. Remember that the MCT architecture (as we know it) is completely dependent on manufacturing ALL propellant for the journey home.
The first people to land are not going to be permanent colonist.
Not doom. They would not be colonists but they will IMO likely stay for a full synod until a replacement crew arrives. Probably a few stay on to extend their experience to the next crew. Worst case something goes wrong and they stay two synods. On the next synod they would receive spare parts to get fuel ISRU going.
BTW I was surprised that Elon Musk mentioned they would land humans only when the return fuel is waiting for them. I would have expected to land the equipment but set it up for fuel production only after the human landing.
It makes perfect sense, otherwise you could literally doom/maroon the first crew. Remember that the MCT architecture (as we know it) is completely dependent on manufacturing ALL propellant for the journey home.
The first people to land are not going to be permanent colonist.
Not doom. They would not be colonists but they will IMO likely stay for a full synod until a replacement crew arrives. Probably a few stay on to extend their experience to the next crew. Worst case something goes wrong and they stay two synods. On the next synod they would receive spare parts to get fuel ISRU going.
Mostly Irrelevant, because they are not going to risk crews on the first mission and landing anyway. And the first unmanned missions need to provide a proof of concept that the ISRU equipment functions. So since they will already have hardware on the ground, making it an all-up test that produces propellant makes the most sense.
Propellant mass fractions that are on par with SSTO vehicles
combined with re-entry conditions that make lunar return look mild.
Propellant mass fractions that are on par with SSTO vehicles
Not really, there's a huge difference between what's needed for 6.5-8 km/s vs. 9.5 km/s. (I get a mass ratio of about 5.75 for 6.5 km/s, 8.6 for 8 km/s, 12.8 for 9.5 km/s).
Quotecombined with re-entry conditions that make lunar return look mild.
Yeah but they also have a better heat shield material than Apollo.
The only thing that might make sense for an HLV is if you had some good rationale for flying enough payloads per year that you could take FULL advantage of a full manufacturing line (multiple shifts) /and/ full advantage of full reuse (both first and second stages). To keep a production line busy, you probably need at least 10 flights per year. To do that with multiple shifts, probably about 40 first stage cores per year (and more upper stages). To make first stage reuse make sense, you need to reuse the stage at least ten times. To make full reuse make sense, you need to reuse the first stage at least 100 times and the upper stage 10-30 times. That's roughly a thousand launches per year at a minimum. Partial reuse can get by with maybe 100 launches per year, 40 per year if some are expendable. Otherwise, it's not really the economic optimum. Unless you get to 1000 launches per year, though, you probably aren't launching enough to gain anything by using an HLV. (In fact, you'd reduce your economic efficiency since you wouldn't be reusing as much.)
We do keep hearing high numbers from SpaceX. Fleets, scouting the globe for launch locations to move a lot of people. But I don't think we think in the right scale yet.
Also, you're missing a really important point: There isn't actually a market for HLV launch. Nobody* wants more than about EELV Heavy capability (except for exploration, but even then it's not required). There aren't even commercial or military processing facilities for payloads over 5 meters in diameter. There isn't a commercial launch need for BFR, and NASA doesn't even have any payloads that need an HLV. F9 and Falcon Heavy are MORE than enough for all the commercial, civil, and military payloads. I've held this viewpoint for years, I've been vocal about it, and I haven't actually changed my mind on it. The only thing that might make sense for an HLV is if you had some good rationale for flying enough payloads per year that you could take FULL advantage of a full manufacturing line (multiple shifts) /and/ full advantage of full reuse (both first and second stages). To keep a production line busy, you probably need at least 10 flights per year. To do that with multiple shifts, probably about 40 first stage cores per year (and more upper stages). To make first stage reuse make sense, you need to reuse the stage at least ten times. To make full reuse make sense, you need to reuse the first stage at least 100 times and the upper stage 10-30 times. That's roughly a thousand launches per year at a minimum. Partial reuse can get by with maybe 100 launches per year, 40 per year if some are expendable. Otherwise, it's not really the economic optimum. Unless you get to 1000 launches per year, though, you probably aren't launching enough to gain anything by using an HLV. (In fact, you'd reduce your economic efficiency since you wouldn't be reusing as much.)
Customers don't care how big the rocket is they care how much it costs. In order to get anywhere near the cost structure Musk desires the cost to launch the BFR would need to be well in the range where it would be a cost effective commercial launcher. While it may cost more then one sat customer would be willing to pay for we already have a NORMAL practice in the industry of launching multiple satellites simultaneously so it is not at all hard to see a viable commercial sat launch market for BFR if it achieved the full reusability which is it's main goal.
Quote... So, the Block One MCT will likely again be different from the MCT that lands the first crew on Mars June 2033 following the un-crewed MCT bringing the ISRU equipment May 2031.I have no doubt MCT will evolve, but your timeline is not the same as SpaceX's timeline. They expect crewed missions much earlier. Which makes sense, as it doesn't make sense to develop a capability and then essentially just let it languish, sucking up money while nothing is accomplished (another mistake NASA is making, though this is mostly Congress's fault).
Which is exactly why an integrated 2nd stage is a bad idea, it makes the launcher too narrow and unable to serve any other role efficiently. It is also unfeasible on technical grounds but no one seems willing to admit this.
Musk is borrowing somewhat from Mars Direct (or is it Semi-Direct?) where an already-fueled ascent vehicle is fueled up on the surface. I don't see a good reason not to have a fueled up vehicle ready when they arrive.BTW I was surprised that Elon Musk mentioned they would land humans only when the return fuel is waiting for them. I would have expected to land the equipment but set it up for fuel production only after the human landing.
It makes perfect sense, otherwise you could literally doom/maroon the first crew. Remember that the MCT architecture (as we know it) is completely dependent on manufacturing ALL propellant for the journey home.
The first people to land are not going to be permanent colonist.
Not doom. They would not be colonists but they will IMO likely stay for a full synod until a replacement crew arrives. Probably a few stay on to extend their experience to the next crew. Worst case something goes wrong and they stay two synods. On the next synod they would receive spare parts to get fuel ISRU going.
It could do a Bigelow module or large space telescope or something too, but those would be pretty rare, more one time events. Big telescopes are expensive and you only need so many big Bigelow modules in orbit.
Keep in mind that most likely Falcon won't be going anywhere even once BFR flies. SpaceX is investing if 4 pads for it. Even if they were to turn 39A into a BFR pad at some point, that's still 3 pads. It's sized good for comsats, and you aren't putting a 4mt comsat going to GTO on something the size of Saturn V or bigger. Multiple payloads are a possibility, but ArianeSpace is moving away from that, not doubling down on it with Ariane 6. Otherwise, why not make an Ariane 6 that's larger than Ariane 5 and can launch 5 or 6 sats at once? Becuase it's a bit of a challenge to get just two sats going to close enough of the same orbits that can launch together...much less more.
So I don't see comsats being launched in big clusters by BFR. That's what F9 is for and it'll do a nice job of that.
Bigger than F9 would be an FH with 3 reusable boosters.
So what -could- BFR/MCT launch besides crewed, propellant, cargo, or depot versions of MCT? Even a fully reusable BFR/MCT which requires only nominal processing between flights (unlike STS) will have a fair cost...but could be less than say a D4H, or an FH-E or A5-551 or whatever the heavy Vulcan variant is. So it could be the most affordable launcher for those big government birds.
It could also do planetary probes with an expendable kick stage.
It could do a Bigelow module or large space telescope or something too, but those would be pretty rare, more one time events. Big telescopes are expensive and you only need so many big Bigelow modules in orbit. Cargo service to a big Bigelow station is a possible routine job for BFR though.
So those are some potential markets for BFR, I'm pretty sure it's not going to take over F9R/FHR's markets.
Unfeasible? I don't know about that. How so? Because keep in mind, even if you have a dedicated reusable 2nd stage, and MCT sits on top, that 2nd stage will still almost certainly be some sort of biconic. It might get away with a more blunted nose and ballistic reentry like the reusable Falcon 9 upper stage, but it will still have to come in nose first...just like MCT. And since it's likely to come back from GTO trajectories if it's to deploy sats where most customers want them deployed, it'll be coming in faster than from just LEO.
Which an integrated MCT platform would already be designed to do. But would a dedicated S2 be designed for that? It'd only ever need to go to LEO for purposes of launching MCT, and taking propellant to a LEO depot. It'd need to have HEO return capability built into it's TPS...again, like the integrated stage already will have.
Which is what RobotBeat has hit on in a few posts. These are complementary tasks, not competing tasks. Where there is overlap anyway, why not go with it? Why have a separate 2nd stage that will have to be designed to do basically the same thing as the MCT basic platform?
And how will it be easier to launch unmanned payloads from a biconic 2nd stage that cannot carry people (in any version) than it is from a biconic 2nd stage that can carry people when configured in a certain way?
I'm not seeing the advantage...or how the later is technically unfeasible. If you can do one, you should be able to do the other. Either way, you need some sort of payload carrier that can be built in behind the nosecone cap (which would be used for docking and propellant transfer in either the integrated or non integrated concepts), or have some sort of expendable payload fairly that would mount on the nose.
The only real advantage of a dedicated 2nd stage, as best I can see, is to allow an MCT to launch unfueled with fast reaction whole-vehicle abort capability. Something that cannot be done with the integrated design. But that's immaterial for the purpose of launching unmanned payloads to space.
Lets look at all the ways they differ
Active time of flight: R2S (Reusable 2nd Stage): Hours MCT: Months
Entry Velocity: R2S: 7.7 km/s MCT: >12 km/s
Aerocapture necessary: R2S: NO MCT: YES
Landing Site: R2S: Spaceport concrete pad + support facilities MCT: Unprepared martian regolith surface
Landed Payload Mass: R2S: Self dry mass MCT: Self dry mass + 100 mT cargo (mars) 25 mT (Earth)
Payload Carrying System: R2S: Payload adapter at top of tank MCT: Internal cargo bay with doors
Payload Separation Conditions: R2S: Axial decoupling in zero-g MCT: Horizontal removal on mars surface
Single use disposable Landing systems allowed (parachutes, airbags etc): R2S: YES MCT: NO
Abort system necessary: R2S: NO MCT: YES
Take off Site: R2S: Upper atmosphere after stage separation MCT: Unprepared martian regolith surface
The differences are almost endless and I foresee a very different 2nd stage that is a fairly normal cylindrical shape about 20 m tall and 10 m diameter holding around 1200 mT of propellants and equipped with 7 Raptor engines. It would be recovered by using a petal segment heat-shield covering the engines (rather then the unstable head-first entry depicted in the old F9 reuse video). The petals would then open and act as landing legs, the heat shield material can be a single use ablative (PICAX) that is attached to the structural leg/rib. The tank sides would likely have a metallic sandwich TPS system to protect them as well. The top of the vehicle would deploy parachutes, decelerators and other disposable systems from underneath the payload adapter. Grid-fins and vernier engines for touchdown would likewise be positioned at the top. Dry mass fraction of 6% would yield a 75 mT dry mass.
The 1st and 2nd stage need only be able to get the 100mT cargo plus MCT dry weight into appropriate LEO.
Musk said, paraphrasing, that BFR is intended to get approximately 100 tonnes to LEO.
Interpretation 1: 100 tonnes including MCT dry-mass.
Interpretation 2: 100 tonnes in addition to MCT dry-mass.
An interesting all be it old study on fast transit trajectories to and FROM mars.
http://www.dtic.mil/dtic/tr/fulltext/u2/a272591.pdf
(From a couple of pages back.)The 1st and 2nd stage need only be able to get the 100mT cargo plus MCT dry weight into appropriate LEO.
Hmmm, this seems to be another point that people should clarify when they debate options.
Musk said, paraphrasing, that BFR is intended to get approximately 100 tonnes to LEO.
Interpretation 1: 100 tonnes including MCT dry-mass.
Interpretation 2: 100 tonnes in addition to MCT dry-mass.
These are vastly different requirements.
(And that's in addition to discussing whether the MCT is the second stage.)
Musk said, paraphrasing, that BFR is intended to get approximately 100 tonnes to LEO.
Interpretation 1: 100 tonnes including MCT dry-mass.
Interpretation 2: 100 tonnes in addition to MCT dry-mass.
That is not at all what was said and there is no ambiguity whatsoever. The number was 100t net payload landed on Mars.
Which does not exclude the possibility that this capability will be reached only with the second or third iteration of MCT.
How Falcon and BFR split the launch manifest is certainly something to look at but if SpaceX is operating a full portfolio of reusable rockets the trading of launches between them becomes much less of an issue because your operating a fleet at that point.
All the non-satellite launch missions for BFR I have stated already as well, and these would not actually require much modification if the vehicle is a conventional 2 stage rocket, though the Earth Escape probe would definitely need an escape stage I see something like Centaur just being put under the probe and treated as part of the payload and not something SpaceX provides.
Finally their are potential savings by 'Raptorizing' the Falcon vehicles and retiring Merlin production, 4 Raptor engines should substitute for the 9 first stage Merline engines if a small landing engine/s is used either in the center or around the periphery. So far SpaceX has saved money by only making one turbo-pump fed engines, once Raptor is available the incentive is clearly their to get rid of the old engine and propellant mix, though I am not sure what would be done in the upper stage of Falcon, possibly a low thrust variant can be made.
So effectively MCT has a superset of the R2S requirements. Anything the R2S can do the MCT can do as well, so there is no need for the R2S.
That is not what I think a reusable 2nd stage would look like, this equivalency has been your core argument from the beginning, that a reusable 2nd stage and the MCT are so equal in performance that they might as well be the same vehicle but I couldn't disagree more.
Lets look at all the ways they differ
Active time of flight: R2S (Reusable 2nd Stage): Hours MCT: Months
Entry Velocity: R2S: 7.7 km/s MCT: >12 km/s
Aerocapture necessary: R2S: NO MCT: YES
Landing Site: R2S: Spaceport concrete pad + support facilities MCT: Unprepared martian regolith surface
Landed Payload Mass: R2S: Self dry mass MCT: Self dry mass + 100 mT cargo (mars) 25 mT (Earth)
Payload Carrying System: R2S: Payload adapter at top of tank MCT: Internal cargo bay with doors
Payload Separation Conditions: R2S: Axial decoupling in zero-g MCT: Horizontal removal on mars surface
Single use disposable Landing systems allowed (parachutes, airbags etc): R2S: YES MCT: NO
Abort system necessary: R2S: NO MCT: YES
Take off Site: R2S: Upper atmosphere after stage separation MCT: Unprepared martian regolith surface
So effectively MCT has a superset of the R2S requirements. Anything the R2S can do the MCT can do as well, so there is no need for the R2S.
For GEO satellite deployment MCT could be refuelled in LEO, transfer to GEO, drop off the satellites and return. Only one MCT per year is necessary for all the commercially competed GEO satellites, plus 1-3 tanker flights.
Payload Carrying System: R2S: Payload adapter at top of tank MCT: Internal cargo bay with doors
The differences are almost endless and I foresee a very different 2nd stage that is a fairly normal cylindrical shape about 20 m tall and 10 m diameter holding around 1200 mT of propellants and equipped with 7 Raptor engines. It would be recovered by using a petal segment heat-shield covering the engines (rather then the unstable head-first entry depicted in the old F9 reuse video). The petals would then open and act as landing legs, the heat shield material can be a single use ablative (PICAX) that is attached to the structural leg/rib. The tank sides would likely have a metallic sandwich TPS system to protect them as well. The top of the vehicle would deploy parachutes, decelerators and other disposable systems from underneath the payload adapter. Grid-fins and vernier engines for touchdown would likewise be positioned at the top. Dry mass fraction of 6% would yield a 75 mT dry mass.
For GEO satellite deployment MCT could be refuelled in LEO, transfer to GEO, drop off the satellites and return. Only one MCT per year is necessary for all the commercially competed GEO satellites, plus 1-3 tanker flights.
A 2nd stage would have similar terminal velocity to Dragon, if SpaceX can do fully propulsive landings with Dragon 2, they should be able to do do propulsive for a 2nd stage. Petals which open as landing legs may even give a lower terminal velocity than Dragon 2.
Although I like the petal idea I can't quite see how they are configured during launch when they obviously cannot cover the engines.
For GEO satellite deployment MCT could be refuelled in LEO, transfer to GEO, drop off the satellites and return. Only one MCT per year is necessary for all the commercially competed GEO satellites, plus 1-3 tanker flights.
I think it better to just have it go to a GTO, and then do what's typically done today, which is drop it off there and have the payload place itself in the GEO orbit. I think the propulsive hit becomes large when entering a GEO orbit to directly deploy the sat, and then have to deorbit from there.
A GTO will bring it right back to Earth, where it can do a small deorbit burn and come right home.
Briefly butting in here. Consider gaining flight history on Raptor ahead of any BFR concept.
And consider that FH's US is limiting factor in certain mission profiles.
Only way I can factor in a non-BFR use of Raptor is as a oversized diameter US, flown on only FH. For a limited scope of missions. Which is against the SX economics as a whole I'll grant.
Things that argue for this:
* Raptor originally was a hydrolox US engine - clearly they earlier saw the need
* Raptor was scaled up in size, then scaled down in size - launch architecture clearly being "tuned" around US and Mars ascent requirements
* Environment to prove this would be in high/no atmosphere, not test stand
* You'd also want long duration in space operations, not unlike what ACES/IVF is aimed for
Quote from: Impaler link=topic=37808.msg1439383#msg1439383 So what -could- BFR/MCT launch besides crewed, propellant, cargo, or depot versions of MCT?
[/quote
Is there a chance to have a depot in a GTO orbit? Maybe an 80,000 km apogee.
Deliver satellite, adjust orbit to meet up with depot, deliver prop load, then return to Earth.
Once the depot is full, I believe it could make its way to EML with quite a low dV boost.
Cheers, Martin
For GEO satellite deployment MCT could be refuelled in LEO, transfer to GEO, drop off the satellites and return. Only one MCT per year is necessary for all the commercially competed GEO satellites, plus 1-3 tanker flights.
I think it better to just have it go to a GTO, and then do what's typically done today, which is drop it off there and have the payload place itself in the GEO orbit. I think the propulsive hit becomes large when entering a GEO orbit to directly deploy the sat, and then have to deorbit from there.
A GTO will bring it right back to Earth, where it can do a small deorbit burn and come right home.
Up to now I thought the same. But given the capabilities of a refuelled MCT GSO may well be the better option. No need to give com sats the capability to reach GSO on their own. GTO was so far the better option because the launch vehicle could be smaller and the upper stage can easily deorbit after placing the payload in GTO. None of the restrictions would be applicable for MCT, especially when refuelled but with few sats maybe even without refuelling.
Interpretation 2: 100 tonnes in addition to MCT dry-mass.
Musk mentioned ten cargo MCTs per passenger MCT, so 11 MCT's per 100 colonists.
However, orbital refuelling means multiple launches in addition to the main launch. So if you need, say, three launches of fuel per MCT-to-Mars, then you've 44 launches per 100 colonists.
If the price per colonist is $500,000, that puts the price of a single launch at $1.14m (less for the launch itself, since the total price includes all the extra orbital handling for refuelling, the ECLSS for 100day flights, etc.
Sounds like a plan to me.
When you think about how expensive BFR would have to get before it's not a viable commercial launcher, it's a lot higher than the point where it's no longer a viable Mars colony vehicle.
I think you are right Lobo. SpaceX has the same diameter S2 on Falcon 9, and the same engine optimized for vacuum. It makes sense to have the same Raptor engine second stage optimized for vacuum, same fuel as first stage, even the same landing legs. Either a clamshell heat shield for the S2 engines that doubles as interstage, or the conical one side coated heat shield on the second stage/MCT for atmospheric entry. Like you said, a dedicated S2 might through 20 more tons into orbit, but then you have three things going on. S1, S2, MCT. Cheaper would be S1, S2/MCT combo. Both build robustly and reusable for cost reusability cost effectiveness. Same with airlines. Same planes are for either passengers or cargo. Only the guppy versions for large diameter items are different.
I'm not saying a dedicated 2nd stage couldn't be made cheaper and easier than an integrated MCT basic platform stage. It's criteria aren't as harsh as for MCT.
But, that doesn't speak to your original comment about an integrated design being technically unfeasible. There may be other reasons to go with a dedicated 2nd stage, but I don't think those are reasons because the integrated design is some sort of impossible or implausible concept.
Now, if the integrated design is feasible, that just leaves the question of if it's the better way to go or not?
Mike speaks well to that here:
So effectively MCT has a superset of the R2S requirements. Anything the R2S can do the MCT can do as well, so there is no need for the R2S.
For GEO satellite deployment MCT could be refuelled in LEO, transfer to GEO, drop off the satellites and return. Only one MCT per year is necessary for all the commercially competed GEO satellites, plus 1-3 tanker flights.
That's always been my point. You have to have MCT anyway. It already can do everything a cheaper/easier S2 can do anyway. Why not just make it be the S2 then?
Will it be a little heavier than the dedicated S2? Probably. It probably won't have quote the payload capability of the integrated MCT platform. But will that be a problem? Probably not. It's only going to be an issue for two potential MCT/BFR missions.
1) Unmanned payload launching to space.
2) Tanker service to depot.
The first shouldn't be much issue, because there's already nothing that needs 100mt of throw capacity (although it would likely be even more than that, because it would be stripped down to just the tanks, engines, and OML) To my knowledge there's no current or near term future need for more capacity than D4H or FHE will have to LEO or to GTO. And such an integrated BFR would have much more capacity than those. So if you have 100mt to LEO instead of say 120mt to LEO for a dedicated lighter/cheaper S2, it doesn't hurt you from the commercial or government payload launching market.
As for a tanker, it probably wouldn't be able to get quite as much propellant to a depot per launch. But then again, you have a reusable launch system. So if it takes 5 BFR launches to fill up a depot prior to a Mars mission instead of 4, is that really a detriment?
And again, when comparing the dedicated S2 to the basic integrated MCT platform...stripped of everything but the bare shell, tanks, engines, and landing gear, I doubt the difference would be all that much.
Also, as a satellite launcher, it wouldn't really need it's big MCT solar arrays. And it would have large residuals of propellants. So a little methalox fuel cell or IVF engine could generate the short term loiter power it needs, so it wouldn't even need to have the solar arrays. A dedicated S2 would need something like that too....or a lot of batteries. Either way that would be needed.
Payload Carrying System: R2S: Payload adapter at top of tank MCT: Internal cargo bay with doors
You could either have an internal cargo bay for MCT...as there will be an empty volume between the nose and tanks, or have a payload adapter on the nose like you say for the S2. Either set up will have a TPS on the nose, so there's not much difference in have a payload adapter on one vs. the other.
The differences are almost endless and I foresee a very different 2nd stage that is a fairly normal cylindrical shape about 20 m tall and 10 m diameter holding around 1200 mT of propellants and equipped with 7 Raptor engines. It would be recovered by using a petal segment heat-shield covering the engines (rather then the unstable head-first entry depicted in the old F9 reuse video). The petals would then open and act as landing legs, the heat shield material can be a single use ablative (PICAX) that is attached to the structural leg/rib. The tank sides would likely have a metallic sandwich TPS system to protect them as well. The top of the vehicle would deploy parachutes, decelerators and other disposable systems from underneath the payload adapter. Grid-fins and vernier engines for touchdown would likewise be positioned at the top. Dry mass fraction of 6% would yield a 75 mT dry mass.
Yes, there can be differences, but it doesn't change that MCT would do everything S2 would do, and much more. So the advantage of going that way is you only need to develop one spaceship, and all of them have the same basic common platform. Where as the two spacecraft you've described appear to have different landing legs, different TPS systems, and different methods of landing...a lot of different duplicated systems that cost more to develop and support, and add extra mass to the stack when launching MCT.
I think there's a plausible argument than the economic and logistic advantage...as well as mass advantage to LEO of have the one common integrated system outweigh loosing a little mass efficiency as a tanker or satellite launcher...especially when it could still be very capable at both.
which gets to the larger point, I think as others have pointed out, MCT/BFR will be designed to do specifically one thing primarily...serve as a Mars vehicle. It'll be able to do other things of course, (which can generate money to help fund the Mars goals...definately a good thing for SpaceX) but the hardware won't be designed around those things primarily, it'll be designed around going to Mars as safely and economically and efficiently as possible.
An integrated design also has better throw capacity to LEO when launching MCT...which would be it's primary designed function. Whereas with the dedicated S2, it would have more throw capacity to LEO when launching payloads other than MCT...which isn't it's primary designed function.
So (IMHO) a dedicated S2 will probably only be part of the system if there's a Mars-centric reason for it. Like a whole vehicle fast abort option for MCT or maybe they want the most efficient tanker possible, which would be a dedicated S2-tanker. Or some other Mars related reason.
....So basically you're saying SpaceX wouldn't want their 2nd stage to be capable of cislunar flights like Vulcan's 2nd stage will be capable of?
Lets look at all the ways they differ
Active time of flight: R2S (Reusable 2nd Stage): Hours MCT: Months
Entry Velocity: R2S: 7.7 km/s MCT: >12 km/sSo basically you're saying BFR wouldn't be useful for anything beyond LEO? Interesting, as beyond-LEO is where the vast majority of the commercial and military launch market is (and commercial/military launch is basically the whole argument for why you'd want to release BFR before MCT). Anyway, SpaceX is probably going to use PICA-X, which works in both cases. A big reason why RLVs are normally considered for LEO is that they're either single stage (does not apply here) or use a low-temperature TPS (also does not apply).
Aerocapture necessary: R2S: NO MCT: YESThis is basically a subset of the above. Either could work with aerocapture.
Landing Site: R2S: Spaceport concrete pad + support facilities MCT: Unprepared martian regolith surfaceCitation needed. There's a pretty good chance SpaceX could be sending Dragons with equipment ahead of time. In either case, this would be irrelevant after the first MCT landing. You cannot say that the MCT will be landing on unprepared regolith.
Landed Payload Mass: R2S: Self dry mass MCT: Self dry mass + 100 mT cargo (mars) 25 mT (Earth)The first valid point!
Payload Carrying System: R2S: Payload adapter at top of tank MCT: Internal cargo bay with doorsYet another argument to skip straight to MCT, so they don't have to develop a payload adapter at the top of the tank.
Payload Separation Conditions: R2S: Axial decoupling in zero-g MCT: Horizontal removal on mars surfaceMeh. Horizontal removal works, we've done it before. And using MCT means you ALSO can recover payloads from orbit or potentially save expensive payloads in case of a first stage mishap. Additionally, if the payload bay is at the top, they still can do axial decoupling in zero-g just fine.
Single use disposable Landing systems allowed (parachutes, airbags etc): R2S: YES MCT: NONot from what we saw from SpaceX's Falcon 9 upper stage reuse video. Additionally, SpaceX is moving /away/ from this sort of expensive disposable system. Why would they decide to reverse course and invest millions on a technology that for them is a dead-end?
Abort system necessary: R2S: NO MCT: YESOnly needed for crewed MCT flights. There are other ways to do abort for crewed flights, like the Delta Clipper concept.
Take off Site: R2S: Upper atmosphere after stage separation MCT: Unprepared martian regolith surfaceAgain, you have NO citation for that, and there are very good arguments that MCT will be neither landing on nor taking off from unprepared Martian regolith.
The differences are almost endless and I foresee a very different 2nd stage that is a fairly normal cylindrical shape about 20 m tall and 10 m diameter holding around 1200 mT of propellants and equipped with 7 Raptor engines. It would be recovered by using a petal segment heat-shield covering the engines (rather then the unstable head-first entry depicted in the old F9 reuse video). The petals would then open and act as landing legs, the heat shield material can be a single use ablative (PICAX) that is attached to the structural leg/rib. The tank sides would likely have a metallic sandwich TPS system to protect them as well. The top of the vehicle would deploy parachutes, decelerators and other disposable systems from underneath the payload adapter. Grid-fins and vernier engines for touchdown would likewise be positioned at the top. Dry mass fraction of 6% would yield a 75 mT dry mass.So in other words, MCT doesn't fit with what your ENTIRELY made-up reusable 2nd stage would look like, even though it probably fits pretty well with what SpaceX /has/ released so far with respect to a reusable 2nd stage, so obviously they're totally different. Uh huh.
However, many viewed [the Mars Direct plan] as too technologically ambitious to be credible – citing a myriad of issues, such as hopelessly optimistic technology assumptions, and lack of adequate mass margin.
The Mars Semi-Direct architecture proposal that followed (also 1991) might be interpreted as a partial rebuttal to some of these concerns. Instead of specifying such a wildly ambitious “do it all” ERV vehicle, the function of ferrying the crew from Mars surface back to Earth was split up into two parts, to be performed by two separate vehicles: A Mars ascent vehicle (MAV), which needed only to generate enough fuel for the crew to ascend to Mars’ orbit, and the Earth return vehicle (ERV) which was pre-placed in Mars’ orbit – and which would perform the propulsive maneuver needed to send the crew back to Earth.
Did you remember to factor in the ISP differences, all thouse SSTO vehicles were going to use Hydro-Lox at 450 ISP. When you use use that kind of ISP the 9.5 km/s DeltaV requires an 8.62 ratio. So yes they are quite equivalent.
Their are g-forces, entry corridor width and landing accuracy to consider as well as heating,
if it has dropped enough to not completely consume the entire vehicle mass on it's own.
Let me clarify.
Impaler, I don't think your ideas of how to build a reusable upper stage are technically bad, I'm just not at all convinced that that is what SpaceX has in mind, based on their F9 reusable upper stage, their work with Dragon, and their various statements about MCT/BFR. A more integrated system fits very well with statements like this from Musk:
"We’re looking at our Mars transporter being around 15 million pounds of thrust"
(from here: http://forum.nasaspaceflight.com/index.php?topic=37839.msg1391925#msg1391925 )
Did you remember to factor in the ISP differences, all thouse SSTO vehicles were going to use Hydro-Lox at 450 ISP. When you use use that kind of ISP the 9.5 km/s DeltaV requires an 8.62 ratio. So yes they are quite equivalent.
Oh I think you'd use methane or propane for a pure-rocket (not Skylon) reusable SSTO in real life, liquid hydrogen just isn't worth it, especially since the greater losses on ascent (IIRC) and the huge bulky tankage largely wipe out the advantages.Their are g-forces, entry corridor width and landing accuracy to consider as well as heating,
MCT is doing propulsive landing not parachute, so it should get a landing accuracy advantage for that.Quoteif it has dropped enough to not completely consume the entire vehicle mass on it's own.
Wait, what? Mars re-entry velocities aren't nearly THAT high! The Galileo probe did an entry at over 40 km/s and even it wasn't 100% heat-shield (about 50% IIRC) and that's like x10 the kinetic energy. So how could MCT possibly have to be all heat-shield?
I'm well aware that PICA is the best ablative we currently have but neither you nor anyone else who insists it will be used can give me a heat-shield mass for a prospective vehicle and entry profile.And? Can you give me a heatshield mass estimate? I don't see your estimate. Previous times we estimated the math, we came up with something like:
And for the millionth time structural mass counts too and are of a greater concern then the thermal protection systems.For the millionth time, pressure stabilization will be used and can be used for the compressive loads. Additionally, tensile structures inside can be made out of high performance materials like Spectra (or certain types of carbon fiber), with a strength-to-weight ratio ten times that of high performance aluminum alloys as well as very high toughness.
At entry the vehicle will have nearly 50% payload mass, just SRP will require ~20% propellant fraction. That leaves something like 30% for all the structure, all the TPS, all the engines and tank mass, landing gear, systems used only in space like solar arrays and radiators that have been stowed etc etc. This is not remotely easy even when you do as I've suggested and enter from mars orbit at just 4 km/s rather then twice as fast or more coming in on direct entry.3.5km/s, actually, not 4km/s (actually, at the equator, it's more like 3.25km/s). Easy, no, but mars luckily has places with much lower terminal velocity. Valles Marineris is 5km below datum, and Hellas Basin can be 8km below datum. Terminal velocity is proportional to: e^(h/(2*10.8km)) since the scale height of Mars is 10.8km, so at -5km in Valles Marineris, you have a terminal velocity 79% of that at the datum.
...How does it make no sense when you yourself thought of an acceptable solution? :)
The last statement really makes no sense, how is MCT full of propellants at entry when you used them all to to depart Earth? Or are you in favor of SEP tugs bringing it to LMO, a solution I'm considering for Earth return? You can't bring propellants from the surface of mars...
The only real point you've got there is the requirements for landing a dry stage and landing a significant payload on Mars are different, so you'd probably need somewhat beefier legs on MCT than what would be absolutely essential for an empty upper stage. That's true, but not a good enough reason to develop a completely different vehicle
Heck, even if BFR/MCT DOES have 3 effective stages (big booster, reusable 2nd stage, and MCT), you'd still probably want to keep the 2nd stage and MCT /very/ similar in order to avoid having to pay for development of a completely different vehicle, totally different TPS, etc. (as well as the loss of safety since you aren't also testing MCT every time you launch another payload)
I need to see some better figures from you on the total mass and thrust of your BFR concept and the actual mass of this integrated 2nd stage to even get an idea what your LEO performance is and what it's likely to cost, you really can't throw around generic statements like 'lighter' when I have SPECIFIC mass estimates and DeltaV goals. You need to do more homework to demonstrate this (or show the homework from L2 as I suspect this is where it all is). The launch vehicle performance calculator http://www.silverbirdastronautics.com/LVperform.html should be the basis for comparison as I've used it for my calculations (median performance estimates only).
Wait a sec, we are both in favor of Bi-conic entry vehicles for MCT (you convinced me of that), and you can't put anything on top of a bi-conic that is where the heat-shield is, you can't have exposed attachment point for payload, they would melt and I can't see retracting them into doors as being practical either. The picture we were shown of a F9 reusable upper stage was fatally flawed when it depicted a heat shield on the top of the stage, it is both unstable as an entry vehicle and it provides no attachment, that's why I came up with the petal heat-shield/legs and bottom first entry.
MCT will necessarily have to rely on some preplaced ground infrastructure. It will be incredibly foolish to hobble the fundamental design just because of the very first mission. If it is a problem, SpaceX can use Dragon to pre-land a rover or a crane, or some one-shot modifications made to MCT to allow it to land on an unprepared surface.
I'm not sure you get to have short, stubby legs and rear engine hoverslam landings at the same time on unprepared soil. Excavation by the plume is a problem.
It's not about strength, it's about center of mass and stability; Excavation by those engines is going to fragment the landing surface erratically, and even a small slope will tip the thing over in a static case, to say nothing of the levering action of actually colliding with the surface.I'm not sure you get to have short, stubby legs and rear engine hoverslam landings at the same time on unprepared soil. Excavation by the plume is a problem.
Nobody said it would be easy. :) There will have to be a engineering trades, but I certainly believe a MCT gear must be very strong - and to save mass, short. Remember that it also needs to be capable of supporting a full propellant load before takeoff.
The legs will need to be more like Dragons legs than F9R.
I'm not sure you get to have short, stubby legs and rear engine hoverslam landings at the same time on unprepared soil. Excavation by the plume is a problem.
I believe the landing legs won't need to be like the Falcon 9 first stage legs. F9 is long and slender. MCT will be short and stubby in comparison.
But, in general, when I say that the IBMCT will mass less than an MCT + S2, and thus will have more gross LEO capability, that's simple logic. You aren't duplicating your hardware and systems, like you are with a S2R + MCT. You have one TPS, not two. [...]
So if you assume the same booster for either [...]
IF you were doing a 3 stage to LEO, that might be different. But you are talking TSTO either way [...]
No, I *strongly* disagree with this. MCT will need to be capable of landing on unprepared Martian ground. There is no question about it, if you think about it.
I do apologize if I allude to anything I can't expand on. I'm trying to stay away from doing that because it's poor form.
And remember DC-X demonstrated the Swan Dive maneuver, a transition from slightly angled down nose-first entry to vertical landing. A small aerosurface or two might help if you wanted to go at an even greater angle, but I see nothing that suggests the transition is impossible.
BTW, I guess that conceptually, I think of MCT as a less-ambitious version of DC-Y/DeltaClipper/DC-I. Less ambitious because: It wouldn't need to be SSTO (6-8km/s is all that's needed, which makes a HUGE difference vs 9.5km/s... basically it means you can afford TWICE the dry mass including payload!), it uses methane (which in spite of the Isp hit probably would make SSTO easier due to the FAR higher bulk density) instead of hydrogen, and it'd basically always operate in vacuum except for final landing at Earth, thus making an aerospike nozzle unnecessary.
I'm not sure you get to have short, stubby legs and rear engine hoverslam landings at the same time on unprepared soil. Excavation by the plume is a problem.
I believe the landing legs won't need to be like the Falcon 9 first stage legs. F9 is long and slender. MCT will be short and stubby in comparison. I believe also that the legs will not need to support MCT fully fuelled for launch. They could add supports for that purpose before tanking takes place.
But, in general, when I say that the IBMCT will mass less than an MCT + S2, and thus will have more gross LEO capability, that's simple logic. You aren't duplicating your hardware and systems, like you are with a S2R + MCT. You have one TPS, not two. [...]
C'mon, you know you can't linearly sum non-linear systems like that. The systems on two vehicles will be individually lighter (and designed for a lighter vehicle) than the systems on a larger integrated vehicle. For example, the TPS on the second stage will not need to cope with Mars direct-return velocity. Ie, 7km/s instead of 11, or just 40% of the energy and at a lower g-load. Same with all the flow-on effects of requiring larger systems, higher mass, then sturdier structures to deal with the higher mass, increasing the mass further...
So if you assume the same booster for either [...]
IF you were doing a 3 stage to LEO, that might be different. But you are talking TSTO either way [...]
The MCT has tanks and engines. Why wouldn't you take advantage of that and use it as a third stage to increase payload?
However, for me the clincher is that the if we assume SpaceX take the same incremental approach they did with the development of F9R and Dragon, then the development path goes through the lower requirements of a reusable second stage. Incrementally developing that stage will provide them with key insights in developing the MCT. For example, in theory, a larger stage will be easier to re-enter due to its lower density; but in practice, structural strength has been an issue. Which one dominates in the MCT design?
A less demanding second stage should help them learn as they go. If you can more easily solve the structural issues, you go big - Integrated stage MCT. If structural issues dominate, you split the vehicles - MCT with separate second stage. Developing directly to MCT will invariably involve making decisions early that cause problems later, since you don't know in advance which systems are going to work better than expected and which are going to be harder, more expensive, or higher maintenance.
[Example, SpaceX has apparently found for FH that increasing the performance of the Merlin engines is easier than cross-feed. And for BFR, that more engines on a single core is easier to manage than more cores; which goes against previous industry assumptions.]
The issue isn't "what is the optimum Mars vehicle that I, and my chums on L2, can design", instead it's "what is the likely lowest-cost development path for SpaceX for the whole system". That path goes through a second stage. Where it leads after that... depends on how that second stage performs.
The issue isn't "what is the optimum Mars vehicle that I, and my chums on L2, can design", instead it's "what is the likely lowest-cost development path for SpaceX for the whole system". That path goes through a second stage. Where it leads after that... depends on how that second stage performs.
My take is if the second stage can come back through the thicker earth's atmosphere, and land in the heavier gravity than Mars, then it should be easy to modify for the MCT, thus serving two purposes. Bottom half could be the same. Top half would either be empty for a payload, or the modular MCT cargo, human habitation, and solar panels for power along with the metholox production equipment.Right, we are in violent agreement. :)
...
The issue isn't "what is the optimum Mars vehicle that I, and my chums on L2, can design", instead it's "what is the likely lowest-cost development path for SpaceX for the whole system". That path goes through a second stage. Where it leads after that... depends on how that second stage performs.
In addition to your two points I would add that with the ability to land on unprepared surfaces the MCT would be usable for sub-orbital hops of perhaps several thousand miles and thus serve as a means of rapid transit between base locations and the hinterlands, effectively the same role that helicopters and bush-planes would serve on Earth and we see heavy use of such vehicle in wilderness/pioneering areas.
The issue isn't "what is the optimum Mars vehicle that I, and my chums on L2, can design", instead it's "what is the likely lowest-cost development path for SpaceX for the whole system". That path goes through a second stage. Where it leads after that... depends on how that second stage performs.
This is exactly the way that SpaceX operates - they have learned the Soviet incremental approach to goals, but with vertically integrated manufacturing.
For the same dry mass and tank size
Dry vehicle weight, to first order, scales with volume (look at the pressure vessel equation), so it is a fair comparison. Did you read my link?For the same dry mass and tank size
That's the major flaw in the work your siting, at equal volume the difference in propellant density means a huge difference in gross take off weight and gross vehicle weight would follow that relationship not remain static. And costs scale with dry vehicle weight not volume so it is simply not a fair comparison....
Quick thought that I haven't seen anyone mention.
When ISS visiting vehicles are opened up, the crew wear masks to avoid floating debris - and this is after the loading crews make efforts to avoid contamination.
For a crew that are living out of MCT on the surface, or even just making occasional visits, ISTM they will trek in a lot more debris than the CRS vehicles suffer.
How would this be mitigated? Crew wear masks and the ventilation turned to max to filter the air as quickly as possible?
Cheers, Martin
The reason they wear facemasks when opening CRS vehicles is because in zero-gee, stuff floats around and you could inhale it...
Oh, right. Once a base is set up with separate habs, you'd probably have a time when you'd clean out the MCT before launch. Swab the deck, etc. Before that, with small crews and such, you'd probably wear dust masks when you got in orbit.The reason they wear facemasks when opening CRS vehicles is because in zero-gee, stuff floats around and you could inhale it...
Yup, that was exactly what I was discussing. Human occupation of the MCT on Mars' surface will create a debris problem once it's launched back to LMO.
Cheers, Martin
...The size of boulders that would pose a problem for an MCT landing would be faaar to large to handle with a small rover. And boulders are only part of the issue, dust blasting may be more of a problem, and that certainly cannot be addressed by a small rover....
15 years ago I had a bobcat working around my house. I fail to see how you get that into the and out of the Dragon 2.
15 years ago I had a bobcat working around my house. I fail to see how you get that into the and out of the Dragon 2.
Yeah, Red Dragon is a good idea for getting some experiments to Mars without having to design a custom lander, but something like a small robotic bulldozer will need a custom lander.
...The size of boulders that would pose a problem for an MCT landing would be faaar to large to handle with a small rover. And boulders are only part of the issue, dust blasting may be more of a problem, and that certainly cannot be addressed by a small rover....
Boulders big enough to cause problems for MCT landing are easily spotted using MRO. You can just avoid them in the planning stages.
It seems like SpaceX went to some pain to eliminate hydrazine RCS from Falcon 9, replacing it with compressed nitrogen at a much poorer performance level.
What exactly would be used for high-frequency, low-latency thrust in an MCT? One of the ideas I'm tossing around is supplementing the massive methalox tanks & Raptor engines with a moderate amount of hydrazine + NTO, and using SuperDracos. This fulfills the need for RCS (which I don't think raptors are suited for), does not introduce any new engines (which SpaceX have specified: they're only working on one) and gives some thermal benefits relative to methalox alternatives.
An alternative might be one of the upcoming green propellant blends.
15 years ago I had a bobcat working around my house. I fail to see how you get that into the and out of the Dragon 2.Red Dragon would necessarily have modifications to allow egress of equipment.
MRO has a resolution of ~30cm per pixel. Multiple exposures of the same site from different angles and ESPECIALLY with the Sun at high angles (thus casting long shadows) can identify hazards....The size of boulders that would pose a problem for an MCT landing would be faaar to large to handle with a small rover. And boulders are only part of the issue, dust blasting may be more of a problem, and that certainly cannot be addressed by a small rover....
Boulders big enough to cause problems for MCT landing are easily spotted using MRO. You can just avoid them in the planning stages.
1) MRO can resolve objects of "about a meter across"...
I am quite sure they will use pressure fed methalox thrusters for RCS. The pressurized tanks can be small and get refilled from the main tanks when needed. The morpheus moon lander testbed has not only the main engine but methalox thrusters too. I would not see this as a contradiction to working on one engine only. Thrusters are not engines in that sense.
If your backing off to a hohmann transfer speed direct from mars that would indeed be ~6 km/s and comes in at a dry mass fraction of 20%, considerably more reasonable and well below the mass fraction of SSTO vehicles of any propellant, if you were to include an efficient aerocapture system like magneto-plasma I could even see this vehicle being possible all be it optimistic. But the transit time is now no good for passengers, only cargo.
MRO has a resolution of ~30cm per pixel. Multiple exposures of the same site from different angles and ESPECIALLY with the Sun at high angles (thus casting long shadows) can identify hazards....The size of boulders that would pose a problem for an MCT landing would be faaar to large to handle with a small rover. And boulders are only part of the issue, dust blasting may be more of a problem, and that certainly cannot be addressed by a small rover....
Boulders big enough to cause problems for MCT landing are easily spotted using MRO. You can just avoid them in the planning stages.
1) MRO can resolve objects of "about a meter across"...
MRO has a resolution of ~30cm per pixel. Multiple exposures of the same site from different angles and ESPECIALLY with the Sun at high angles (thus casting long shadows) can identify hazards.Boulders big enough to cause problems for MCT landing are easily spotted using MRO. You can just avoid them in the planning stages.
1) MRO can resolve objects of "about a meter across"...
Perhaps, but that kind of coverage does not exist. Yes. And even when it does, it tells you nothing about the relative strength of the surface. It could be the Martian equivalent of quicksand for all we know.
But it is still irrelevant. MCT will need to be able to land on unprepared terrain, it will be necessary to allow of off-nominal EDL and abort scenarios. So it will need a sturdy gear, and you seem reluctant for some reason to admit that.
Wooo, your landing mass is more like 200 mT, not 300 and your going to use Raptor engines for all your deceleration and bring yourself to a hover at 100 m from the surface. The touchdown engines are not slamming your into the surface they are just countering the gravity on mars and giving you fine control to maneuver around any boulders.Ahh, I knew that sounded like a lot. Thanks for spotting the error - 10 times too much mass allocated to engines. That should bring the extra mass margin down below +10%.
The thrust total your projecting is more then an entire Raptor engine would produce, if we needed or wanted that much we wouldn't bother with these vernier engines so I think you've mixed up some where (several some wheres?), for example you make each engine weight 1 mT when you were adding them to the vehicles mass.
A 200 mT lander on mars has a weight of 744 kN meaning you need just 10 Super Draco engine equivalents which means 1 mT and no significant impact to the landers dry mass. Their would be some cosine loss too but I actually expect an improved engine running on Methane to be employed improving the thrust as well as a touchdown mass of just 175 mT. The propellant for a few seconds of touchdown (20 seconds at mars gravity would be 75 m/s) is not going to be significant propellant fraction, though obviously a significant portion was burned prior to this by the Raptor engine, that is a separate calculation unaffected by our choice of touchdown systems.
On Earth were going to be lighter by 75 mT and we would be landing on a concrete pad so we can have a little higher speed at touch down and we might very briefly throttle up right at touchdown too. I've doubtful that Raptor can safely fire in atmosphere but if I'm wrong then it would certainly be used here when their is no danger of debris from the surface.
The set of eight superdracos on Dragon 2 could land 100mT on Mars with margin to spare, once velocity is reduced to near zero by the Raptor engines. Scaling these engines up and fueling them with methlox should provide the fine control needed for landing. Could be used to lift-off the surface, too, before the large, centerline engine(s) are started. This technology/approach could be useful for a reusable lander that explores undeveloped sites.
The set of eight superdracos on Dragon 2 could land 100mT on Mars with margin to spare, once velocity is reduced to near zero by the Raptor engines. Scaling these engines up and fueling them with methlox should provide the fine control needed for landing. Could be used to lift-off the surface, too, before the large, centerline engine(s) are started. This technology/approach could be useful for a reusable lander that explores undeveloped sites.
I would *really like* to employ them in an integrated single-vehicle system for liftoff, but that would also mean a hell of a lot more of them. Liftoff thrust requirements are a large multiple of landing thrust requirements.
If you've off-loaded 100mT of cargo and taken on equivalent fuel, the lift-off problem is same as landing.Could be used to lift-off the surface, too, before the large, centerline engine(s) are started.I would *really like* to employ them in an integrated single-vehicle system for liftoff, but that would also mean a hell of a lot more of them. Liftoff thrust requirements are a large multiple of landing thrust requirements.
The set of eight superdracos on Dragon 2 could land 100mT on Mars with margin to spare, once velocity is reduced to near zero by the Raptor engines. Scaling these engines up and fueling them with methlox should provide the fine control needed for landing. Could be used to lift-off the surface, too, before the large, centerline engine(s) are started. This technology/approach could be useful for a reusable lander that explores undeveloped sites.
I would *really like* to employ them in an integrated single-vehicle system for liftoff, but that would also mean a hell of a lot more of them. Liftoff thrust requirements are a large multiple of landing thrust requirements.
If you've off-loaded 100mT of cargo and taken on equivalent fuel, the lift-off problem is same as landing. Why are lift-off thrust requirements 'a large multiple' of landing? You only need to clear the ground by 100m (and maybe move laterally a bit)...
The set of eight superdracos on Dragon 2 could land 100mT on Mars with margin to spare, once velocity is reduced to near zero by the Raptor engines. Scaling these engines up and fueling them with methlox should provide the fine control needed for landing. Could be used to lift-off the surface, too, before the large, centerline engine(s) are started. This technology/approach could be useful for a reusable lander that explores undeveloped sites.
I would *really like* to employ them in an integrated single-vehicle system for liftoff, but that would also mean a hell of a lot more of them. Liftoff thrust requirements are a large multiple of landing thrust requirements.
If you've off-loaded 100mT of cargo and taken on equivalent fuel, the lift-off problem is same as landing. Why are lift-off thrust requirements 'a large multiple' of landing? You only need to clear the ground by 100m (and maybe move laterally a bit)...
Because of all that methalox! At landing the vehicle is near the penultimate dry mass. At liftoff it's at around 3.35x the dry mass (in the case of 380s Isp & LMO refueling at ~4.5km/s dV), or 6.55x the dry mass (in the case of 380s Isp & no LMO refueling with Hohmann transfer home at ~7km/s dV) or more (in fast transit cases without LMO refueling).
This translates directly into proportionately higher thrust. This higher thrust figure is achievable, but the sheer weight of the engines required adds quite a bit to the vehicle.
Admittedly, you need lower acceleration at liftoff than at landing; I need to do further math on this.
If you've off-loaded 100mT of cargo and taken on equivalent fuel, the lift-off problem is same as landing.Could be used to lift-off the surface, too, before the large, centerline engine(s) are started.I would *really like* to employ them in an integrated single-vehicle system for liftoff, but that would also mean a hell of a lot more of them. Liftoff thrust requirements are a large multiple of landing thrust requirements.
To reach TEI from Mars surface, you're looking at around 400 tonnes of fuel (depending on dry-mass). Additionally, launch needs to be at 2-3 g to reduce gravity losses. So minimum 2*9.8*500 = 9.3MN. So over 130 Super-Dracos just to reach 100m. (Hell, 25 just to hover.)
But it is still irrelevant. MCT will need to be able to land on unprepared terrain, it will be necessary to allow of off-nominal EDL and abort scenarios. So it will need a sturdy gear, and you seem reluctant for some reason to admit that.
Musk is borrowing somewhat from Mars Direct (or is it Semi-Direct?) where an already-fueled ascent vehicle is fueled up on the surface. I don't see a good reason not to have a fueled up vehicle ready when they arrive.
Perhaps, but that kind of coverage does not exist. Yes. And even when it does, it tells you nothing about the relative strength of the surface. It could be the Martian equivalent of quicksand for all we know.
But it is still irrelevant. MCT will need to be able to land on unprepared terrain, it will be necessary to allow of off-nominal EDL and abort scenarios. So it will need a sturdy gear, and you seem reluctant for some reason to admit that.
You should have watched the NASA workshop for selecting Mars landing sites. They do know a lot. They have identified landing spots with a hard surface that allow for safe landing. They do have a lot of coverage for different kinds of observation already and the teams can request more observations for each of the proposed 40 landing sites. They can do very thorough orbital survey for multiple data once they have narrowed down to few potential sites.
MCT will certainly not need to be designed for landing on any not suitable off target landing sites so don't try to include such requirements into your mass budget.
I did watch many of those workshops. And I also know that observations from the above is not as comprehensive as one might think, unless one also has ground observations to validate them. And if you have any reference of how they can judge the hardness of a surface (brittle and compressible vs hard as granite), then please link to the presentation where they show that.
If you are willing to claim that no off-nominal landing will ever happen, nor be planned for, then go ahead. I'm not part of the crowd that cries for a useless abort system, but I do think a sturdier landing gear with some extra margin is mass well spent. Extra margin to allow landing on marginal sites during an emergency/off-nominal, NOT any site on Mars.
The set of eight superdracos on Dragon 2 could land 100mT on Mars with margin to spare, once velocity is reduced to near zero by the Raptor engines. Scaling these engines up and fueling them with methlox should provide the fine control needed for landing. Could be used to lift-off the surface, too, before the large, centerline engine(s) are started. This technology/approach could be useful for a reusable lander that explores undeveloped sites.
I would *really like* to employ them in an integrated single-vehicle system for liftoff, but that would also mean a hell of a lot more of them. Liftoff thrust requirements are a large multiple of landing thrust requirements.
If you've off-loaded 100mT of cargo and taken on equivalent fuel, the lift-off problem is same as landing. Why are lift-off thrust requirements 'a large multiple' of landing? You only need to clear the ground by 100m (and maybe move laterally a bit)...
Because of all that methalox! At landing the vehicle is near the penultimate dry mass. At liftoff it's at around 3.35x the dry mass (in the case of 380s Isp & LMO refueling at ~4.5km/s dV), or 6.55x the dry mass (in the case of 380s Isp & no LMO refueling with Hohmann transfer home at ~7km/s dV) or more (in fast transit cases without LMO refueling).
This translates directly into proportionately higher thrust. This higher thrust figure is achievable, but the sheer weight of the engines required adds quite a bit to the vehicle.
Admittedly, you need lower acceleration at liftoff than at landing; I need to do further math on this.
Not true... a 100mT payload plus the vehicle dry weight (~25mT?) at landing -- at lift-off, 25mT plus 4x propellant gives same mass as at landing. (Sorry for the gross approximations.) This gets you back to LMO where you fuel-n-go for Hohmann transfer home. If vehicle is a lander, it refuels in LMO and prepares for another descent to the surface.
The 'sheer weight' of a set of 8 superdracos is less than 1mT if I recall correctly.
Ahh, I knew that sounded like a lot. Thanks for spotting the error - 10 times too much mass allocated to engines. That should bring the extra mass margin down below +10%.
If you're trying to avoid unprepared landing site excavation issues, you need space for the exhaust plume to spread out a bit. I picked ~1km for a very rough & arbitrary figure. At 100m, the exhaust plume is barely larger than the vehicle fairing diameter. Shut the main engines down at 1km AGL after doing the full job of entry & descent, and they won't destroy the landing pad. Then drop for a bit (now's a nice time to correct to vertical and unfurl the legs), & start controlled thrusting on the sideways-canted canard engines; Any terrain damage they do will be well away from the place the legs impact the ground. They still need plenty of thrust, however. If they were only capable of precisely Mars gravity acceleration, they could hover at this 1km point, but not descend (because they could never correct for that additional velocity); Pure gravity loss. To minimize gravity loss they need substantially larger thrust in order to accomplish an efficient suicide burn.
EDIT: To clear up your confusion, a supplementary set of engines towards the top of the vehicle ("Canard engines") are intended to solve the problems raised in posts like this http://forum.nasaspaceflight.com/index.php?topic=37466.msg1372150#msg1372150 . As a secondary point, they might be used if a nested MAV design turns out to be needed because of low ISRU mass payoffs.
... I expect SpaceX to sell normal commercial flights on BFR for a good long time before it is use for mars ...
... As NASA is the only conceivable customer for a first mission they need to be courted to create a mission utilizing SpaceX as the primary contractor....
Ahh, I knew that sounded like a lot. Thanks for spotting the error - 10 times too much mass allocated to engines. That should bring the extra mass margin down below +10%.
If you're trying to avoid unprepared landing site excavation issues, you need space for the exhaust plume to spread out a bit. I picked ~1km for a very rough & arbitrary figure. At 100m, the exhaust plume is barely larger than the vehicle fairing diameter. Shut the main engines down at 1km AGL after doing the full job of entry & descent, and they won't destroy the landing pad. Then drop for a bit (now's a nice time to correct to vertical and unfurl the legs), & start controlled thrusting on the sideways-canted canard engines; Any terrain damage they do will be well away from the place the legs impact the ground. They still need plenty of thrust, however. If they were only capable of precisely Mars gravity acceleration, they could hover at this 1km point, but not descend (because they could never correct for that additional velocity); Pure gravity loss. To minimize gravity loss they need substantially larger thrust in order to accomplish an efficient suicide burn.
EDIT: To clear up your confusion, a supplementary set of engines towards the top of the vehicle ("Canard engines") are intended to solve the problems raised in posts like this http://forum.nasaspaceflight.com/index.php?topic=37466.msg1372150#msg1372150 . As a secondary point, they might be used if a nested MAV design turns out to be needed because of low ISRU mass payoffs.
I think 1 km is far too high an elevation to start worrying about plumes impinging the ground. A large rocket lifting off the ground is not still bathing the launch pad in flames when it is 1 km up, rather it looks to be mostly over within 1-2 times the height of the launch tower.
Remember our goal as you point out is to avoid making craters in the ground and making dangerous ejecta which might impact the vehicle, at 100 m height their should be no danger to the vehicle even if some sand and dust are being swept up on the surface. Under the near vacuum conditions of mars the plume will also spread MUCH wider then the vehicles base, just as the plume of a rocket expands markedly as it rises, it will never resemble the 'welding torch' look of a rocket at liftoff on Earth.
... I expect SpaceX to sell normal commercial flights on BFR for a good long time before it is use for mars ...
... As NASA is the only conceivable customer for a first mission they need to be courted to create a mission utilizing SpaceX as the primary contractor....
I'm catching up on this thread so apologies for necro-quoting, but is this a widely shared opinion? Basically that by the time SpaceX is ready with BFR, they will not be ready with a mission for it?
Or did you mean that NASA is the only conceivable customer other than SpaceX itself?
I can't be 100% sure that the first BFR mission will head to Mars, but I'm pretty sure there won't be a "good long time" (years?) in which BFR is used commercially before it used for MCTs to Mars. I can see how maybe during off-season they are used for commercial purposes since why not, but they are built for a purpose, and I expect the Mars program to be pretty efficient in that things will mature by the time they are needed.
... I expect SpaceX to sell normal commercial flights on BFR for a good long time before it is use for mars ...
... As NASA is the only conceivable customer for a first mission they need to be courted to create a mission utilizing SpaceX as the primary contractor....
I'm catching up on this thread so apologies for necro-quoting, but is this a widely shared opinion? Basically that by the time SpaceX is ready with BFR, they will not be ready with a mission for it?
Or did you mean that NASA is the only conceivable customer other than SpaceX itself?
I can't be 100% sure that the first BFR mission will head to Mars, but I'm pretty sure there won't be a "good long time" (years?) in which BFR is used commercially before it used for MCTs to Mars. I can see how maybe during off-season they are used for commercial purposes since why not, but they are built for a purpose, and I expect the Mars program to be pretty efficient in that things will mature by the time they are needed.
Vehicle structural dry weight of 25mT is, in my judgement, about an order of magnitude too low for a reusable lander with integrated habitat that brings 100mT payload to the surface. Yes, there's a question about whether ISRU gear is inside or outside the '100mT useful cargo' box, but the weight of the vehicle alone is at 100mT-200mT in most other people's scenarios, I just favor raising it to 300mT-500mT with my own special sauce.
Vehicle structural dry weight of 25mT is, in my judgement, about an order of magnitude too low for a reusable lander with integrated habitat that brings 100mT payload to the surface. Yes, there's a question about whether ISRU gear is inside or outside the '100mT useful cargo' box, but the weight of the vehicle alone is at 100mT-200mT in most other people's scenarios, I just favor raising it to 300mT-500mT with my own special sauce.
The point is that a set of eight superdracos such as is installed on the (existing) Dragon 2 can land 125mT of payload on Mars, given that the Raptor engines have slowed the vehicle to near zero at a nominal distance above the ground. If the vehicle itself is somewhere between 100 and 200mT (say 150) and the delivered payload is 100mT, then double the Dragon 2 complement will do the job nicely. Doubling the complement, four quads instead of four pairs, or doubling the thrust of a pair seems to be an easy step technically.
... I expect SpaceX to sell normal commercial flights on BFR for a good long time before it is use for mars ...
... As NASA is the only conceivable customer for a first mission they need to be courted to create a mission utilizing SpaceX as the primary contractor....
I'm catching up on this thread so apologies for necro-quoting, but is this a widely shared opinion? Basically that by the time SpaceX is ready with BFR, they will not be ready with a mission for it?
Or did you mean that NASA is the only conceivable customer other than SpaceX itself?
I can't be 100% sure that the first BFR mission will head to Mars, but I'm pretty sure there won't be a "good long time" (years?) in which BFR is used commercially before it used for MCTs to Mars. I can see how maybe during off-season they are used for commercial purposes since why not, but they are built for a purpose, and I expect the Mars program to be pretty efficient in that things will mature by the time they are needed.
The business case and usage schedule is far more up in the air than than even the technical details of BFR & MCT. We don't have high-quality speculation to offer. All we know is that BFR launches will be cheaper with high launch rate than with low launch rate. On this basis and on the expectation that only a long record of successful unmanned launches proves safety of a manned launch vehicle to *my satisfaction*, I expect them to seek out whatever customers they can. It's arguably a prerequisite that they find these customers, before scaling up the Mars project; For that matter, the business viability of the Mars project as a private passenger delivery system without Apollo-grade Congressional outlays is... questionable.
... I expect SpaceX to sell normal commercial flights on BFR for a good long time before it is use for mars ...
... As NASA is the only conceivable customer for a first mission they need to be courted to create a mission utilizing SpaceX as the primary contractor....
I'm catching up on this thread so apologies for necro-quoting, but is this a widely shared opinion? Basically that by the time SpaceX is ready with BFR, they will not be ready with a mission for it?
Or did you mean that NASA is the only conceivable customer other than SpaceX itself?
I can't be 100% sure that the first BFR mission will head to Mars, but I'm pretty sure there won't be a "good long time" (years?) in which BFR is used commercially before it used for MCTs to Mars. I can see how maybe during off-season they are used for commercial purposes since why not, but they are built for a purpose, and I expect the Mars program to be pretty efficient in that things will mature by the time they are needed.
The business case and usage schedule is far more up in the air than than even the technical details of BFR & MCT. We don't have high-quality speculation to offer. All we know is that BFR launches will be cheaper with high launch rate than with low launch rate. On this basis and on the expectation that only a long record of successful unmanned launches proves safety of a manned launch vehicle to *my satisfaction*, I expect them to seek out whatever customers they can. It's arguably a prerequisite that they find these customers, before scaling up the Mars project; For that matter, the business viability of the Mars project as a private passenger delivery system without Apollo-grade Congressional outlays is... questionable.
Burninate you put your finger on the reasons the system needs to find some non-mars bound commercial business, basically cost amortization and flight history build up. I would also add that splitting development up over time, having experience with large reusable vehicles that survive high speed entry at Earth before trying it on Mars, and finally having a the launchers lift capacity fully characterized and upgraded to it's maximum potential so you know your ceiling mass for the MCT.
To clarify my prior post, I think NASA (in the course of a large Congressionaly funded mars mission) is the only conceivable first customer for a mars transport system. That transport system will be both the rocket and a mars lander and possibly other transport elements provided by SpaceX (such as SEP tugs, Propellant production equipment). Think of it like Mars-COTS, SpaceX schleps people & cargo for NASA under contract.
This is the only way SpaceX can actually make any revenue from mars, they might be able to direct their satellite launch profits into vehicle development but actually run the system they need huge payments at low volume and government is the only possible customer during that start up period even if you want to sell to private interest later, just as we are under no illusions that Dragon capsule would have been developed without the COTS program.
SpaceX will control the whole payload interface of the BFR and mars bound vehicle and design it for their goals, and these will probably not resemble any current interfaces, so NASA will design all of the cargo it wants sent to mars around the SpaceX vehicles cargo capacity both in terms of mass and volume. That is going to take time and money and training, billions of dollars and perhaps a decade.
But NASA can't actually do anything like that when it is saddled with the Orion/SLS millstone which basically sucks up all it's money. The BFR can lift this off NASA's shoulders IF it can match all the SLS capabilities, is drastically cheaper, safer AND SpaceX is manufacturing in enough states to get support or at least defuse opposition in the senate to the cancellation of these programs. SLS won't go away until BFR has a solid launch history, the political forces will always be able to justify keeping it going if SpaceX isn't superior in every metric.
So the first priority is to get SLS retired, everyone knows that SLS and BFR would be in direct competition and the universe is not big enough for them both. If SpaceX get BFR up and running quickly I could see SLS being canceled in about a decade after a few flights, probably ARM as a face-saving last mission. That opens up $3 billion a year for actual mars systems. Meanwhile SpaceX will be well into designing the mars bound vehicle/s and a it will be both politically and economically easy to slip into a parallel development course in which SpaceX finishes the mars vehicle while NASA gives out contracts for surface system development as consolation prizes to the folks who lose out when SLS/Orion ended. In another 10 years both are done and your ready to land on mars.
But we know that SpaceX, even in the fast build-up scenario, needs to send multiple unmanned payloads to Mars first.
BFR is an Earth-to-orbit vehicle, and so you're not risking a two-year delay when you first launch it. Yes, the first launch may not carry an MCT, but by the same token, it won't carry any other one-of-a-kind payload. Being reusable, it's a complete non-brainer to fly a dummy payload. If it works, you you relaunch. It it doesn't, good thing you didn't.
Actually, once you're shifted to a reusable rocket, test flights are cheap enough that you could do multiple dummy payloads before you put an MCT on top - and it's not like you wasted multiple rockets doing it.
So I think the "build up flight history" argument doesn't apply.
... I expect SpaceX to sell normal commercial flights on BFR for a good long time before it is use for mars ...
... As NASA is the only conceivable customer for a first mission they need to be courted to create a mission utilizing SpaceX as the primary contractor....
Is this a widely shared opinion?
But we know that SpaceX, even in the fast build-up scenario, needs to send multiple unmanned payloads to Mars first.
BFR is an Earth-to-orbit vehicle, and so you're not risking a two-year delay when you first launch it. Yes, the first launch may not carry an MCT, but by the same token, it won't carry any other one-of-a-kind payload. Being reusable, it's a complete non-brainer to fly a dummy payload. If it works, you you relaunch. It it doesn't, good thing you didn't.
Actually, once you're shifted to a reusable rocket, test flights are cheap enough that you could do multiple dummy payloads before you put an MCT on top - and it's not like you wasted multiple rockets doing it.
So I think the "build up flight history" argument doesn't apply.
<snip>
... I expect SpaceX to sell normal commercial flights on BFR for a good long time before it is use for mars ...
... As NASA is the only conceivable customer for a first mission they need to be courted to create a mission utilizing SpaceX as the primary contractor....
I'm catching up on this thread so apologies for necro-quoting, but is this a widely shared opinion? Basically that by the time SpaceX is ready with BFR, they will not be ready with a mission for it?
Developing BFR and MCT is a huge investment by SpaceX. I dont think they can make it without picking up money in some fashion along the way. Even the internet satellite constellation cant do that much on the time scales that are envisioned. Hell, BFR might even be used to get it to orbit.
My expectation is, that the first version of BFR will be a vehicle, purely used for commercial satellites. BFRv1 would be shorter than the one used for Mars and less capable of course. I expect it to have an early version of the Raptor engine, maybe even less engines in total. I also expect that a cargo version of MCT will ride BFR as an integrated second stage. It would have no pressurized volume, but large cargo doors to release satellites in orbit, up to GTO. It would be re-usable, test reentry and landing systems and would be a precursor to MCT in general.
I dont think SpaceX will be able to fund BFR to Mars from the get go. Not even if NASA hitches a ride. Development and production of BFR is way too expensive. They need tons of tests before cargo can be send to Mars for real. Better use BFR commercially as early as possible and refine designes on the way. Thats how SpaceX operated in the past and I assume will continue to operate in the future. Look at the development of Falcon 9 and Dragon. Its the same story, I expect history to repeate it self.
I voiced that view in before and from experience, most users here on the forum disagree with that view.
What do you mean? You can, right now, go and request new images be taken of a certain area. There's plenty of "coverage" to take multiple images of the same small area at multiple times of day.MRO has a resolution of ~30cm per pixel. Multiple exposures of the same site from different angles and ESPECIALLY with the Sun at high angles (thus casting long shadows) can identify hazards....The size of boulders that would pose a problem for an MCT landing would be faaar to large to handle with a small rover. And boulders are only part of the issue, dust blasting may be more of a problem, and that certainly cannot be addressed by a small rover....
Boulders big enough to cause problems for MCT landing are easily spotted using MRO. You can just avoid them in the planning stages.
1) MRO can resolve objects of "about a meter across"...
Perhaps, but that kind of coverage does not exist.
Yes. And even when it does, it tells you nothing about the relative strength of the surface. It could be the Martian equivalent of quicksand for all we know.Untrue, we have ground-truthing of multiple sites on Mar that allow us to calibrate the orbital instruments with on-the-ground measurements. We have day and night thermal IR, which allows us to figure out thermal inertia, etc. And there isn't such a thing as "Martian Quicksand," which requires a lot of liquid water. We can identify sand dunes, etc.
But it is still irrelevant. MCT will need to be able to land on unprepared terrain, it will be necessary to allow of off-nominal EDL and abort scenarios. So it will need a sturdy gear, and you seem reluctant for some reason to admit that.I am in favor of actual abort capability, instead of super heavy landing gear. Why do I need to "admit" something that has no basis in actual SpaceX communications, just some people's opinion on the internet? My mental model of MCT most resembles DC-Y and DC-I, which used fairly stubby landing gear and had real abort capability. (And yes, you'd need Soyuz-style cushioning thrusters on Mars in addition to a parachute. But the actual delta-v for that cushioning thrust is VERY low.)
Probably because it's another nail in the idea that MCT will have single digit dry mass percentage.Oh, give me a break. You have not come even CLOSE to establishing that single-digit-dry-mass percentage isn't possible. (Not that it'd be that bad... you can still get more than 6km/s out of it, thus within the range of my expectation for MCT of 6-8km/s.)
Lets look at what landing gear will likely mass. The primary driver is touchdown mass which would be 100 mT + vehicle dry mass + reserves from the landing propellants. I'd call that all 200 mT, and no we do not get to deduct because of mars gravity, the vehicle has kinetic energy at touchdown which is independent of gravity.That's false. Mars' lower gravity makes it easier to zero out the velocity more precisely than on Earth. So you have proportionally less kinetic energy (or, since kinetic energy is proportional to velocity squared, perhaps much, much less kinetic energy).
The landing legs on F9R are said to be 10% of the empty stage mass (and that's with use of carbon-fiber) so I take this as my basis for MCT legs structure. Around 20 mT, though I expect them to telescope inside the vehicle rather then be on the surface.No, we can do much better. For a VTVL SSTO or similar, something like 3% is often assumed for parametric purposes in "common wisdom," but we can do much better.
CO/O2 is really neat, but orbital propellant collection is very difficult, and although I certainly support its development, I have absolutely no reason to think that's what SpaceX is planning. In fact, we can be reasonably certain that Raptor will use methane (and certainly it will use some sort of hydrocarbon), and that's what MCT will use.Musk is borrowing somewhat from Mars Direct (or is it Semi-Direct?) where an already-fueled ascent vehicle is fueled up on the surface. I don't see a good reason not to have a fueled up vehicle ready when they arrive.
I see no reason why to land the fuel production equipment at all.
It can do the same job on orbit.
you have said it yourself:
http://forum.nasaspaceflight.com/index.php?topic=17984.msg620589#msg620589
But why would the passenger version of MCT use the super heavy legs of the full payload version? F9R's legs are removable and added at the launch site, so it would make perfect sense to be able to use less massive legs if you aren't landing a huge payload and that was a big mass driver.
Robotbeat, you have mentioned this 'prepared surface' in several posts. How this would be accomplished?
and you land on a prepared surface.
But why would the passenger version of MCT use the super heavy legs of the full payload version? F9R's legs are removable and added at the launch site, so it would make perfect sense to be able to use less massive legs if you aren't landing a huge payload and that was a big mass driver.
I haven't noticed this distinction before. Dramatically different masses for passenger vs cargo MCT - is that what you meant?
Earlier I'd raised this in a different way, saying that a cargo MCT would contain denser materials than a passenger MCT, and thus to have the same mass in the same volume would have "free space" that would be better shared with a small passenger group who would greatly appreciate free space. Some people responded that the cargo MCT would just be packed with extra space.
But you seem to be saying they'd actually just have a heavier version. More fuel would be part of that too of course. A higher mass for cargo MCT would impact its velocity (is it a slow trajectory?) and EDL issues/differences as well as launch differences.
Actually... perhaps you're saying the opposite. You'd make a regular cargo MCT, but there'd be a low-load light version too for passengers? I would think you wouldn't travel light in a system capable of carrying more.
I'd like to read the arguments for this... where/when was it?
What do you mean? You can, right now, go and request new images be taken of a certain area. There's plenty of "coverage" to take multiple images of the same small area at multiple times of day.MRO has a resolution of ~30cm per pixel. Multiple exposures of the same site from different angles and ESPECIALLY with the Sun at high angles (thus casting long shadows) can identify hazards....The size of boulders that would pose a problem for an MCT landing would be faaar to large to handle with a small rover. And boulders are only part of the issue, dust blasting may be more of a problem, and that certainly cannot be addressed by a small rover....
Boulders big enough to cause problems for MCT landing are easily spotted using MRO. You can just avoid them in the planning stages.
1) MRO can resolve objects of "about a meter across"...
Perhaps, but that kind of coverage does not exist.
QuoteBut it is still irrelevant. MCT will need to be able to land on unprepared terrain, it will be necessary to allow of off-nominal EDL and abort scenarios. So it will need a sturdy gear, and you seem reluctant for some reason to admit that.I am in favor of actual abort capability, instead of super heavy landing gear. Why do I need to "admit" something that has no basis in actual SpaceX communications, just some people's opinion on the internet? My mental model of MCT most resembles DC-Y and DC-I, which used fairly stubby landing gear and had real abort capability. (And yes, you'd need Soyuz-style cushioning thrusters on Mars in addition to a parachute. But the actual delta-v for that cushioning thrust is VERY low.)
....So what? MER was before MRO....
You may not recall this, but the MER scientists were *quite* surprised that the landscape was as flat and featureless as it turned out to be.
QuoteBut it is still irrelevant. MCT will need to be able to land on unprepared terrain, it will be necessary to allow of off-nominal EDL and abort scenarios. So it will need a sturdy gear, and you seem reluctant for some reason to admit that.I am in favor of actual abort capability, instead of super heavy landing gear. Why do I need to "admit" something that has no basis in actual SpaceX communications, just some people's opinion on the internet? My mental model of MCT most resembles DC-Y and DC-I, which used fairly stubby landing gear and had real abort capability. (And yes, you'd need Soyuz-style cushioning thrusters on Mars in addition to a parachute. But the actual delta-v for that cushioning thrust is VERY low.)
This is not what early MCT's will encounter. Even if your fantasyGive me a break.
of tiny ground-clearing robots delivered by Red Dragon's materialize,If you think a 10-ton bulldozer is "tiny"
the ground is not hard as diamond. It won't be even close to a prepared landing pad on Earth.Yes, there will continue to be improvements, not just clearing land.
And you continue to dodge the fact that the MCT legs also need to support a full propellant load, this will require something significantly more than an "elephant on stilts" approach.Enough with the dramatic rhetoric! I've not "dodged" anything. I haven't seen it come up to need to address it, though I have thought of it. Jeez. No, you wouldn't launch an MCT on its legs. You also need supports placed under the MCT before you fuel it up for launch, just like with DC-X (DC-X's legs could support a full load, but they were practicing for DC-Y/I).
There are several methods. As I posted earlier, some of the guys from "Swampworks" are establishing a technique that could be built with a rover. I guarantee that SpaceX is working the problem and has their own ideas of how to solve the problem.Robotbeat, you have mentioned this 'prepared surface' in several posts. How this would be accomplished?
and you land on a prepared surface.
Robotbeat, you have mentioned this 'prepared surface' in several posts. How this would be accomplished?
and you land on a prepared surface.
Robotbeat, you have mentioned this 'prepared surface' in several posts. How this would be accomplished?
and you land on a prepared surface.
NASA has some ideas so maybe SpaceX will get some advice from from them.
http://www.nasa.gov/content/landing-pads-being-designed-for-extraterrestrial-missions
It's as if you and Impaler want MCT to be a non-viable, heavy monstrosity that barely can push its own weight around. It's almost as if you're not trying to imagine how to reduce dry mass in meaningful ways that just about any aerospace undergrad would think of (let alone a team of seasoned professionals).I think it's about having "conservative" mass estimates instead of "optimistic" ones, a.k.a. the reason why Mars Direct in its original form was criticized, then revised (e.g. Semi-Direct, DRM 3.0).
Robotbeat, you have mentioned this 'prepared surface' in several posts. How this would be accomplished?
and you land on a prepared surface.
NASA has some ideas so maybe SpaceX will get some advice from from them.
http://www.nasa.gov/content/landing-pads-being-designed-for-extraterrestrial-missions
"Of all the substances we studied, ablative materials seem to work best," Metzger said.
The problem with that is "conservative" mass fractions are NOT realistic!It's as if you and Impaler want MCT to be a non-viable, heavy monstrosity that barely can push its own weight around. It's almost as if you're not trying to imagine how to reduce dry mass in meaningful ways that just about any aerospace undergrad would think of (let alone a team of seasoned professionals).I think it's about having "conservative" mass estimates instead of "optimistic" ones, a.k.a. why Mars Direct was revised (e.g. Semi-Direct, DRM 3.0).
The problem with that is "conservative" mass fractions are NOT realistic!It's as if you and Impaler want MCT to be a non-viable, heavy monstrosity that barely can push its own weight around. It's almost as if you're not trying to imagine how to reduce dry mass in meaningful ways that just about any aerospace undergrad would think of (let alone a team of seasoned professionals).I think it's about having "conservative" mass estimates instead of "optimistic" ones, a.k.a. why Mars Direct was revised (e.g. Semi-Direct, DRM 3.0).
If NASA, for instance, were doing initial design of F9 v1.1 full thrust and Falcon Heavy, no way would they contemplate giving the boosters mass fractions of 25 and 30, respectively.
Heck, when NASA was trading ULA's Centaur for ESAS, they literally sandbagged the dry mass of Centaur 50% more than it actually weighs today! How they managed to do that, I do not know.
But "conservative" mass fractions are the least likely for a company that both prides itself in world-class mass-fractions and that can't afford building something much larger and with extra stages than necessary. SpaceX doesn't need to spread work amongst several Centers, there's no reason they'll want to expand the number of stages beyond what they think they can achieve. "Conservative" is not realistic when you don't have infinite money.
QuoteThis is not what early MCT's will encounter. Even if your fantasyGive me a break.Quoteof tiny ground-clearing robots delivered by Red Dragon's materialize,If you think a 10-ton bulldozer is "tiny"
It's as if you and Impaler want MCT to be a non-viable, heavy monstrosity that barely can push its own weight around. It's almost as if you're not trying to imagine how to reduce dry mass in meaningful ways that just about any aerospace undergrad would think of (let alone a team of seasoned professionals).
Construction equipment taken to Mars will probably be made from aircraft aluminum or titanium to reduce weight.
"A rocket is typically tested only once with a dummy payload with subsequent launches being with paid customers who have satellites."
But we know that SpaceX, even in the fast build-up scenario, needs to send multiple unmanned payloads to Mars first.
BFR is an Earth-to-orbit vehicle, and so you're not risking a two-year delay when you first launch it. Yes, the first launch may not carry an MCT, but by the same token, it won't carry any other one-of-a-kind payload. Being reusable, it's a complete non-brainer to fly a dummy payload. If it works, you you relaunch. It it doesn't, good thing you didn't.
Actually, once you're shifted to a reusable rocket, test flights are cheap enough that you could do multiple dummy payloads before you put an MCT on top - and it's not like you wasted multiple rockets doing it.
So I think the "build up flight history" argument doesn't apply.
It will be cheaper per kg to LEO, but it is by no means going to be cheap enough to build an adequate flight history (~12 launches) with nothing but dummy payloads all at SpaceX's own expense. A rocket is typically tested only once with a dummy payload with subsequent launches being with paid customers who have satellites. I expect prices of ~200 million per launch even with recovery (about $1000 per kg, Musk's optimistic goal) so it would cost Billions to do these launches without customers.
I expect that 1st stage recovered will be attempted on every single BFR flight with very likely full success from the start. If first stage recovery works on the dummy flight and is declared a solved problem then the price point will likely be set such that SpaceX is fully covering the cost of the 2nd stage in case it is lost as I expect 2nd stage recover to require a long campaign of attempts with lots of failures and redesigns as we have seen with F9, the customer will not care any more about the success or failure of these recovery attempts any more then they care about 1st stage recovery attempts now.
By the time you have your 12 flight history your close to getting the 2nd stage to recover reliably and can drop the price to perhaps 100 million and try to get more volume.
Something closer to a utility truck rather than an extreme sports car with razor thin margins.
Did not Musk say somewhere that there would be approximately 10 - 100 ton cargo flights for every one 100 passenger flight? Cargo being necessary for living areas, food growing units, power stations, fuel manufacturing equipment, mining equipment, long term food storage, etc. To me it would be safer and easier just to send 10 people with each MCT. All vehicles would be the same. All would have cargo and people to work, but not be overloaded with large expensive human habitat area on the vehicles. At least until a fairly large viable colony is built for massive transfers of people.
Something closer to a utility truck rather than an extreme sports car with razor thin margins.
Maybe that's the problem. You're saying a utility truck while Musk, the person who's actually going to decide how the MCT is built, is saying a bus type vehicle (and maybe even a cruise ship vehicle later) that has access to landing pads. You're just on the wrong thread.
I was imagining more of a pickup, but, you know, one with an 8-foot bed *and* a crew cab. Because those are nice.Something closer to a utility truck rather than an extreme sports car with razor thin margins.
Maybe that's the problem. You're saying a utility truck while Musk, the person who's actually going to decide how the MCT is built, is saying a bus type vehicle (and maybe even a cruise ship vehicle later) that has access to landing pads. You're just on the wrong thread.
Ah, the lost art of intentionally mis-reading a post. ::)
CO/O2 is really neat, but orbital propellant collection is very difficult, and although I certainly support its development, I have absolutely no reason to think that's what SpaceX is planning. In fact, we can be reasonably certain that Raptor will use methane (and certainly it will use some sort of hydrocarbon), and that's what MCT will use.Musk is borrowing somewhat from Mars Direct (or is it Semi-Direct?) where an already-fueled ascent vehicle is fueled up on the surface. I don't see a good reason not to have a fueled up vehicle ready when they arrive.
I see no reason why to land the fuel production equipment at all.
It can do the same job on orbit.
you have said it yourself:
http://forum.nasaspaceflight.com/index.php?topic=17984.msg620589#msg620589
Where is the 12 flight requirement coming from?
And suppose there is some requirement. You're implying customer flights, and revenue-generating ones at that, are cheaper than SpaceX flight?
For a reusable rocket, test flights are the cheapest option.Quote"A rocket is typically tested only once with a dummy payload with subsequent launches being with paid customers who have satellites."
But the whole point was that this is a reusable rocket. Your "typically" applies to expendable rockets that operate in a field where there are worthy second-rate payloads that can both generate revenue and be of lesser value if lost.
None of that applies here.
If you're building a reusable rocket, then you can test it like you test any aircraft. Send it up (you can test the first stage independently of the rest of the rocket), certify it, then start using it.
You can't use "typically" across a paradigm shift.But we know that SpaceX, even in the fast build-up scenario, needs to send multiple unmanned payloads to Mars first.
BFR is an Earth-to-orbit vehicle, and so you're not risking a two-year delay when you first launch it. Yes, the first launch may not carry an MCT, but by the same token, it won't carry any other one-of-a-kind payload. Being reusable, it's a complete non-brainer to fly a dummy payload. If it works, you you relaunch. It it doesn't, good thing you didn't.
Actually, once you're shifted to a reusable rocket, test flights are cheap enough that you could do multiple dummy payloads before you put an MCT on top - and it's not like you wasted multiple rockets doing it.
So I think the "build up flight history" argument doesn't apply.
It will be cheaper per kg to LEO, but it is by no means going to be cheap enough to build an adequate flight history (~12 launches) with nothing but dummy payloads all at SpaceX's own expense. A rocket is typically tested only once with a dummy payload with subsequent launches being with paid customers who have satellites. I expect prices of ~200 million per launch even with recovery (about $1000 per kg, Musk's optimistic goal) so it would cost Billions to do these launches without customers.
I expect that 1st stage recovered will be attempted on every single BFR flight with very likely full success from the start. If first stage recovery works on the dummy flight and is declared a solved problem then the price point will likely be set such that SpaceX is fully covering the cost of the 2nd stage in case it is lost as I expect 2nd stage recover to require a long campaign of attempts with lots of failures and redesigns as we have seen with F9, the customer will not care any more about the success or failure of these recovery attempts any more then they care about 1st stage recovery attempts now.
By the time you have your 12 flight history your close to getting the 2nd stage to recover reliably and can drop the price to perhaps 100 million and try to get more volume.
You're implying customer flights, and revenue-generating ones at that, are cheaper than SpaceX flight?
For a reusable rocket, test flights are the cheapest option.
Methane and O2 can be produced on orbit provided a supply of H2 from the surface.
that way you get 11 times the fuel on orbit then what you launch.
Methane and O2 can be produced on orbit provided a supply of H2 from the surface.
that way you get 11 times the fuel on orbit then what you launch.
Errr, where does the carbon and oxygen come from?
1. Fuel will be needed to maintain a low orbit, constant adjustments. It would be easier and quicker just to land, refuel, then launch back.1. fuel will be there. that's the point. some of it will be used to maintain altitude and some will be gained.
2. Mars' atmosphere isn't that thick. An orbit low enough to scoop enough atmosphere is probably going to be too low, like to stay in orbit maybe as low as 20 or 30 miles. Might as well land and relaunch.
You should not be so easy to dismiss an idea. check out that thread and see that a lot of work was put into atmospheric scooping starting in the 60's or so.1. Fuel will be needed to maintain a low orbit, constant adjustments. It would be easier and quicker just to land, refuel, then launch back.1. fuel will be there. that's the point. some of it will be used to maintain altitude and some will be gained.
2. Mars' atmosphere isn't that thick. An orbit low enough to scoop enough atmosphere is probably going to be too low, like to stay in orbit maybe as low as 20 or 30 miles. Might as well land and relaunch.
2. than have a bigger scoop or an eliptical orbit.
Can we leave the crackpot ideas off this forum thread? The idea that you can scoop enough enough atmosphere to generate enough propellant to overcome the friction and *also* generate extra propellant is just two degrees short of a perpetual motion machine. It has been studied many times and also immediately rejected as many times.
There are ways to make it work. You need to use a tether (with existing materials, not exotics), thus reducing your relative velocity while also allowing your draggy power source to not slow you down. Practicality is still questionable, but it's an interesting idea.Can we leave the crackpot ideas off this forum thread? The idea that you can scoop enough enough atmosphere to generate enough propellant to overcome the friction and *also* generate extra propellant is just two degrees short of a perpetual motion machine. It has been studied many times and also immediately rejected as many times.
Using SEP, assuming highest performance panels and highest possible thrust drive, the raw numbers just barely work. Although we're talking a scoop in the hundred of metres length range. TRL is maybe 2, if we're generous. It doesn't work with chemical engines, therefore you can't use the scooped fuel to power the SEP. So in practice, TRL is zero or negative. So yeah, it doesn't belong here.
-snip-
From Hop_David's spreadsheet (http://forum.nasaspaceflight.com/index.php?topic=37536.msg1371984#msg1371984), I offer five cases:
Straight Hohmann with propulsive capture - 259 days
A straight-up Hohmann burn from a {1AUx1AU heliocentric, 300km x 300km altitude geocentric} orbit to a {1.524AUx1.524AU heliocentric, 3697x23459km semimajor axis aereocentric} coplanar orbit using a {1AUx1.5240001AU heliocentric} transfer costs 510m/s plus change on Earth departure, over and above the 3220m/s escape velocity burn. Then to propulsively capture into a highly elliptical Mars orbit it costs 1021m/s. Conservative aerobraking can reduce that ellipse down to LMO for another 668m/s that we don't need to pay, where standard EDL penalties apply (I'll assume them to be 2km/s for purposes of this discussion).
Total: 3220 + 510 + 1021 + 2000 = 6751m/s
Wet to drymass ratio at 380s Isp: 6.12 to 1
The spreadsheet author's suggested non-Hohmann trajectory with propulsive capture - 102 days
The author would reduce perihelion in order to balance the perigee and periaerion burn into something sensible, assuming that propulsive capture is necessary. He uses non-prograde burns or a suboptimal burn time in Earth orbit, which is highly inefficient, to balance out the Earth and Mars sides for minimal total dV. From a {1AUx1AU heliocentric, 300km x 300km altitude geocentric} orbit to a {1.524AUx1.524AU heliocentric, 3697x23459km semimajor axis aereocentric} coplanar orbit using a {0.87AUx1.8AU heliocentric} transfer orbit costs 3146m/s plus change on Earth departure, over and above the 3220m/s escape velocity burn. Then to propulsively capture into a highly elliptical Mars orbit it costs 4668m/s. Conservative aerobraking can reduce that ellipse down to LMO for another 668m/s that we don't need to pay, where standard EDL penalties apply (I'll assume them to be 2km/s for purposes of this discussion).
Total: 3220 + 3146 + 4668 + 2000 = 13034m/s
Wet to drymass ratio at 380s Isp: 33.1 to 1
A 100 day semi-Hohmann transit given perfect free aerocapture - 100 days
From a {1AUx1AU heliocentric, 300km x 300km altitude geocentric} orbit to a {1.524AUx1.524AU heliocentric, 3697x23459km semimajor axis aereocentric} coplanar orbit using a {1AUx3.31AU heliocentric} transfer orbit costs 2241m/s plus change on Earth departure, over and above the 3220m/s escape velocity burn. Then to propulsively capture into a highly elliptical Mars orbit it costs 9276m/s, but we're not going to do that: Instead, magnetoshell or some other aerocapture technology is going to do that all in one go; Then it's going to, in the same step, further reduce the elliptical orbit down to LMO for another 668m/s that we don't need to pay, then go directly into EDL, where standard EDL penalties apply (I'll assume them to be 2km/s for purposes of this discussion).
Total: 3220 + 2241 + 2000 = 7461m/s
Wet to drymass ratio at 380s Isp: 7.41 to 1
A 100 day semi-Hohmann transit given propulsive capture - 100 days
Examining the previous proposition without the non-prograde burns that the spreadsheet author makes. From a {1AUx1AU heliocentric, 300km x 300km altitude geocentric} orbit to a {1.524AUx1.524AU heliocentric, 3697x23459km semimajor axis aereocentric} coplanar orbit using a {1AUx3.31AU heliocentric} transfer orbit costs 2241m/s plus change on Earth departure, over and above the 3220m/s escape velocity burn. Then to propulsively capture into a highly elliptical Mars orbit it costs 9276m/s. Conservative aerobraking can reduce that ellipse down to LMO for another 668m/s that we don't need to pay, where standard EDL penalties apply (I'll assume them to be 2km/s for purposes of this discussion).
Total: 3220 + 2241 + 9276 + 2000 = 16737m/s
Wet to drymass ratio at 380s Isp: 89.3 to 1
A reasonable 180 day near-Hohmann transfer with mild aerocapture - 180 days
From a {1AUx1AU heliocentric, 300km x 300km altitude geocentric} orbit to a {1.524AUx1.524AU heliocentric, 3697x23459km semimajor axis aereocentric} coplanar orbit using a {1AUx1.652AU heliocentric} transfer orbit costs 655m/s plus change on Earth departure, over and above the 3220m/s escape velocity burn. Then to propulsively capture into a highly elliptical Mars orbit it costs 2443m/s, but we're not going to do that: Instead, magnetoshell or some other aerocapture technology is going to do that. Conservative aerobraking can reduce that ellipse down to LMO for another 668m/s that we don't need to pay, where standard EDL penalties apply (I'll assume them to be 2km/s for purposes of this discussion).
Total: 3220 + 655 + 2000 = 5875m/s
Wet to drymass ratio at 380s Isp: 4.84 to 1
They're reversible, but you're forgetting the Oberth effect: because on the way from Earth to Mars, you can dump your exhaust in a deeper gravity well than Mars to Earth, it takes less delta-v.Would anyone like to do the work of reversing that spreadsheet so we could look at more refined estimates of return delta V? I might try it at some point, but not today.
Orbits are time reversible so don't we just need to look at the 'Vinf at mars' and calculate the DeltaV needed to archive that escape velocity from mars surface, then we would (if we pointed ourselves in the right direction) be headed back down the equivalent half of the outbound orbit and we should reach Earth in the specified time and with the specified Vinf so we know what we need to do to aerocapture at Earth as well.
It would be nice to have this done for us on the spreadsheet though.
I was just referring to the heliocentric portion of the transit being reversible, we would obviously need to take mars's smaller gravity well into account, that's why I said the Vinf at mars is needs to be looked at. Here let me take some of Burnate's example and reverse them into mars-Earth transits and show how I would do it.
Straight Hohmann with propulsive capture - 259 days
Vinf at mars is 2.652818827 km/s that is our velocity at this point in the heliocentric orbit before we enter the mars gravity well, naturally we accelerate when falling into that gravity well but we would just come out again with the same speed aka were hyperbolic. To get back to Earth we need that same heliocentric speed (pointed back in now) when leaving mars to return to Earth in 259 days. To determine this we just take the escape velocity at mars surface 5 km/s, square it along with the needed Vinf add them together and get the square root which is 5.660 km/s.
Likewise we can use the previously 'escape' burn from Earth to know how much deceleration we need to shed to go into that high orbit at Earth after which we will aerobrake down to LEO, I'll assume that we can be refueled by tanker at this point in order to do our Earth landing as that will avoid having to blast off from mars with propellant we wont need until now in LEO where we know their is a continuous refueling process in place.
Earth return reversal
Total: 5660 + 510 = 6170m/s
Wet to drymass ratio at 380s Isp: 5.24 to 1
The spreadsheet author's suggested non-Hohmann trajectory with propulsive capture - 102 days
Earth return reversal
Total 9241 + 3145 = 12386 m/s
Wet to drymass ratio at 380s Isp: 27.82 to 1
A 100 day semi-Hohmann transit given perfect free aerocapture - 100 days
Earth return reversal
Total 13816 m/s
Wet to drymass ratio at 380s Isp: 40.85 to 1
This can be improved over simply reversing the outbound orbit which was {1AUx3.31AU heliocentric}, I've found {.76AUx1.67AU heliocentric} with the same travel time which can decrease the mars escape burn substantially but at the cost of raising Earth side capture needs.
Earth return reversal
Total 8792 m/s
Wet to drymass ratio at 380s Isp: 10.6 to 1
A 100 day semi-Hohmann transit given propulsive capture - 100 days
Earth return reversal
Total 13816 + 2241 = 16057 m/s
Wet to drymass ratio at 380s Isp: 74.57 to 1
Using my alternative trajectory
Total 8792 + 4597 m/s = 13389
Wet to drymass ratio at 380s Isp: 36.42 to 1
A reasonable 180 day near-Hohmann transfer with mild aerocapture - 180 days
Earth return reversal
Total 7047 m/s
Wet to drymass ratio at 380s Isp: 6.63 to 1
Conclusions: There seems to be no viable fast (100 day) returns from mars, unlike on the outbound even with perfect free aerocapture doesn't allow it at any kind of propellant fraction and DeltaV which would be believable. The problem is that were starting from mars surface unlike LEO in the the outbound assumptions, that's putting us around 3600 m/s behind where we would be. If we had propellant depots in mars orbit fast return would be a reasonable number if we also had free aerocapture as the wet/dry ratio would be 4:1, which would require something around 300 mT of propellant in mars orbit a huge amount.
Given these, I'm going to ask for consensus from both sides of the debate:
Is it possible to conclude at this point that the *only way* to achieve 1 mission per synod per vehicle (given obvious assumptions like methalox + Raptor) is using high-velocity aerocapture as well as refueling from an already-prepared Mars surface fuel depot during a short (days-weeks) surface stay?
Given these, I'm going to ask for consensus from both sides of the debate:
Is it possible to conclude at this point that the *only way* to achieve 1 mission per synod per vehicle (given obvious assumptions like methalox + Raptor) is using high-velocity aerocapture as well as refueling from an already-prepared Mars surface fuel depot during a short (days-weeks) surface stay?
I know no other way to achieve 1 mission per synod per vehicle than using opposition class missions. This requires refuelling from a Mars surface depot.
The NASA Tracjectory Browser (http://trajbrowser.arc.nasa.gov/traj_browser.php?NEAs=on&NECs=on&chk_maxMag=on&maxMag=25&chk_maxOCC=on&maxOCC=4&chk_target_list=on&target_list=Mars&mission_class=roundtrip&mission_type=rendezvous&LD1=2025&LD2=2040&maxDT=1.9&DTunit=yrs&maxDV=15.0&min=DV&wdw_width=-1&submit=Search#a_load_results) can give useful results. Earth re-entry speeds may be pretty high.
Why not fuel depots at either end as well as on the surface?
Given these, I'm going to ask for consensus from both sides of the debate:
Is it possible to conclude at this point that the *only way* to achieve 1 mission per synod per vehicle (given obvious assumptions like methalox + Raptor) is using high-velocity aerocapture as well as refueling from an already-prepared Mars surface fuel depot during a short (days-weeks) surface stay?
I know no other way to achieve 1 mission per synod per vehicle than using opposition class missions. This requires refuelling from a Mars surface depot.
The NASA Tracjectory Browser (http://trajbrowser.arc.nasa.gov/traj_browser.php?NEAs=on&NECs=on&chk_maxMag=on&maxMag=25&chk_maxOCC=on&maxOCC=4&chk_target_list=on&target_list=Mars&mission_class=roundtrip&mission_type=rendezvous&LD1=2025&LD2=2040&maxDT=1.9&DTunit=yrs&maxDV=15.0&min=DV&wdw_width=-1&submit=Search#a_load_results) can give useful results. Earth re-entry speeds may be pretty high.
One other possibility involves LMO refuelling and cargo transfer to dedicated surface to LMO craft.
Given these, I'm going to ask for consensus from both sides of the debate:
Is it possible to conclude at this point that the *only way* to achieve 1 mission per synod per vehicle (given obvious assumptions like methalox + Raptor) is using high-velocity aerocapture as well as refueling from an already-prepared Mars surface fuel depot during a short (days-weeks) surface stay?
I know no other way to achieve 1 mission per synod per vehicle than using opposition class missions. This requires refuelling from a Mars surface depot.
The NASA Tracjectory Browser (http://trajbrowser.arc.nasa.gov/traj_browser.php?NEAs=on&NECs=on&chk_maxMag=on&maxMag=25&chk_maxOCC=on&maxOCC=4&chk_target_list=on&target_list=Mars&mission_class=roundtrip&mission_type=rendezvous&LD1=2025&LD2=2040&maxDT=1.9&DTunit=yrs&maxDV=15.0&min=DV&wdw_width=-1&submit=Search#a_load_results) can give useful results. Earth re-entry speeds may be pretty high.
One other possibility involves LMO refuelling and cargo transfer to dedicated surface to LMO craft.
If you think this is a legit alternative, could you expand on it sufficient to understand what you're proposing? It doesn't seem to mitigate the need for extreme aerocapture or a Mars surface depot or a short stay, and it sounds fairly difficult on top of that.
Still waiting to hear acknowledgements or challenges from several others on my point.
My solution would be to make a vehicle that splits in two, a large drive section and a simple freight section that is just a frame with cargo containers attached all over it's surface and propellant tanks. Upon arrive at the destination the vehicles splits apart and changes out it's freight section and picks up a new one. At Earth a freight section full of cargo and propellants is picked up, at mars this freight section is empty of propellant and is dropped off and a section full of return propellant is picked up. With 3 freight sections per drive section you have one being loaded, one being unloaded and one in transit at all times, and the freight section acts as the propellant depot and the turn-around operation is simplified to a single docking and un-docking for the drive section.
My solution would be to make a vehicle that splits in two, a large drive section and a simple freight section that is just a frame with cargo containers attached all over it's surface and propellant tanks. Upon arrive at the destination the vehicles splits apart and changes out it's freight section and picks up a new one. At Earth a freight section full of cargo and propellants is picked up, at mars this freight section is empty of propellant and is dropped off and a section full of return propellant is picked up. With 3 freight sections per drive section you have one being loaded, one being unloaded and one in transit at all times, and the freight section acts as the propellant depot and the turn-around operation is simplified to a single docking and un-docking for the drive section.
Something like this? ;D
Impaler's concept always made sense to me, but with VTVL.
I envisioned MCT as a fuel-and-propulsion central core, and then cargo/habitat all around it.
It lands, sheds the payload, and can fly back much lighter.
You automatically get the benefit of a large cross section for Mars EDL, and a smaller one when returning to Earth.
Basically, Mars needs enclosed volumes for habitation and storage. It seems such a shame to haul back an empty cargo hold.
This configuration also solves the "how to get the payload to the ground" question. It's already at ground level when you land.
Get back to earth, attach new "saddle bags", and lift off again.
IMO there won't be in-orbit refueling around Mars, and so it is important to minimize the empty mass of MCT on the flight back, so the same engine and tanks will give you more dV.
My solution would be to make a vehicle that splits in two, a large drive section and a simple freight section that is just a frame with cargo containers attached all over it's surface and propellant tanks. Upon arrive at the destination the vehicles splits apart and changes out it's freight section and picks up a new one. At Earth a freight section full of cargo and propellants is picked up, at mars this freight section is empty of propellant and is dropped off and a section full of return propellant is picked up. With 3 freight sections per drive section you have one being loaded, one being unloaded and one in transit at all times, and the freight section acts as the propellant depot and the turn-around operation is simplified to a single docking and un-docking for the drive section.
Something like this? ;D
No nothing like that, it would look like this 8)
(http://vignette3.wikia.nocookie.net/memoryalpha/images/1/1b/USS_Enterprise-D_saucer_separation.jpg/revision/latest?cb=20120205044747&path-prefix=en)
You see I'm not proposing a landing craft, I'm proposing a pure spacecraft (did I not mention it was SEP propulsion based).
>
Not to pick nits, but Enterprise saucer section could land. The entire Intrepid Class (ex: Voyager) could land.
The BFR first stage won't even reach Earth orbit, dose that mean it's not part of the Mars Colonial Transport SYSTEM? I have already clearly described a landing craft (which you railed against most vehemently).But.... "Land the whole thing", right?
Since a lot more freight would have to be hauled to Mars than people. In space could be done by SEP tugs on a continuous bases. LEO to LMO via one or more SEP tugs. A specialty lander for use on Mars could indeed take freight from LMO to the surface. The lander could have ISRU equipment/solar panels to manufacture it's fuel from Martian atmosphere when not in use or there could be a separate fuel farm nearby. Once refueled, it could go pick up another load.
Humans could travel at a much faster rate with the MCT.
This plan would require the BFR.
It would require a reusable second stage.
It would require a fleet of large SEP tugs.
A cargo carrier that can be transferred from the second stage to the SEP tug.
A re-usable lander at Mars.
A fuel farm at Mars or a lander large enough to have it's on ISRU equipment.
An MCT that could be refueled in LEO and fly to Mars. It might not need to land on Mars, just transfer the human habitation module to the lander. MCT would fly back to earth. Lander would take people to Mars surface. The habitation module could be or would be about the same size as a cargo module.
This plan might be cheaper to operate overall, but would require a lot of building and development of specialty components. It would also require a lot of in space dockings and transfers. However, these specialty components could be built by other companies or countries to have a stake in the colonialization process.
On the other hand, if everything Musk wants to build is big, then It might be more simple to go directly from Earth to Mars will a couple of refueling stops for both humans and cargo an have contributors supply fuel or Martian surface cargo to have a stake in colonialization.
The lander could have ISRU equipment/solar panels to manufacture it's fuel from Martian atmosphere when not in use or there could be a separate fuel farm nearby. Once refueled, it could go pick up another load.
Cargo or freight on the other hand might travel a different way.
I'm one of those who in this thread adheres to Musk quotes like "land the whole thing" for the purpose of coherent speculation. However I'll bet that most here agree that SX's concepts for MCT have likely evolved considerably from the few sometimes off the cuff statements by Elon, many several years old. I'll state further that I believe it probable the MCT that actually flies will again have notable differences from the MCT concept that Musk reveals later this year, if he even meets that schedule.
Impaler, Lars-J, spacenut, etc.,
I wonder if the following might be an 'evolutionary' path that SpaceX might pursue: Have an MCT that initially is used for all phases of the mission (launched to Earth orbit via BFR, is refueled there, launches to Mars, lands, drops off cargo and some/most crew, is refueled via ISRU, launches and does a direct return to Earth, lands on Earth). This MCT would not ever transport 100 passengers, much less 100 passengers plus their cargo. Instead it would be used for all expeditionary missions - those used to perform discovery (find easily accessible water sources), setup habitation and other infrastructure (IRSU, comms, modest chemical plants), prepare launch/landing sites, etc. This would proceed for several synods. Crew would be on the order of 5-15 people. Some staying, some rotating out.
Meanwhile, the SEP interplanetary transporter is being assembled in LEO (or other staging area). When this is complete, and a sufficient number of expeditionary missions have transpired, the function of the MCT vehicles transforms: At Earth, they are used to shuttle colonists and their cargo from the surface to the SEP transporter. At Mars, they are used for the converse - they launch empty from the Mars surface, dock with the SEP transporter, offload passengers and cargo, and return to Mars surface. Each SEP transporter will in fact transport 100 (or more) passengers, plus their cargo.
I think this minimizes the number of vehicles developed, stays consistent with SpaceX's reusability ethos, and is mostly consistent with all of the statements thus far ("land the whole thing", "100 passengers at a time", "we're looking at everything, including SEP", etc.)
Impaler, I know you doubt the technical feasibility of an MCT craft that could do all of the phases listed in the first paragraph, but if it only has to accommodate a max crew of, say, 12, does that change the outlook?
Impaler, Lars-J, spacenut, etc.,
I wonder if the following might be an 'evolutionary' path that SpaceX might pursue: Have an MCT that initially is used for all phases of the mission (launched to Earth orbit via BFR, is refueled there, launches to Mars, lands, drops off cargo and some/most crew, is refueled via ISRU, launches and does a direct return to Earth, lands on Earth). This MCT would not ever transport 100 passengers, much less 100 passengers plus their cargo. Instead it would be used for all expeditionary missions - those used to perform discovery (find easily accessible water sources), setup habitation and other infrastructure (IRSU, comms, modest chemical plants), prepare launch/landing sites, etc. This would proceed for several synods. Crew would be on the order of 5-15 people. Some staying, some rotating out.
Meanwhile, the SEP interplanetary transporter is being assembled in LEO (or other staging area). When this is complete, and a sufficient number of expeditionary missions have transpired, the function of the MCT vehicles transforms: At Earth, they are used to shuttle colonists and their cargo from the surface to the SEP transporter. At Mars, they are used for the converse - they launch empty from the Mars surface, dock with the SEP transporter, offload passengers and cargo, and return to Mars surface. Each SEP transporter will in fact transport 100 (or more) passengers, plus their cargo.
I think this minimizes the number of vehicles developed, stays consistent with SpaceX's reusability ethos, and is mostly consistent with all of the statements thus far ("land the whole thing", "100 passengers at a time", "we're looking at everything, including SEP", etc.)
Impaler, I know you doubt the technical feasibility of an MCT craft that could do all of the phases listed in the first paragraph, but if it only has to accommodate a max crew of, say, 12, does that change the outlook?
That is very nearly exactly what I have been saying, only with the MCT being slightly less capable.
I think the MCT will be dependent on SEP tug assistance to do initial crew missions at an acceptable speed (for crew health/GCR issues), but may be able to do cargo missions on it's own.
We have argued a lot about achievable DeltaV values, let me lay down some numbers that express the DeltaV achieved from full propellant loads with different amounts of cargo, 100 mT (the outbound cargo goal), 25 mT (the return cargo goal which would be some kind of habitat), and 0 mT (presumably the return from a cargo mission).
100 mT Cargo + 300 mT propellant + 75 mT dry mass -> 3.7 km/s
25 mT Cargo + 300 mT propellant + 75 mT dry mass -> 5.1 km/s
0 mT Cargo + 300 mT propellant + 75 mT dry mass -> 6 km/s
These numbers are a lot lower then other people want to see but I think they are realistic and if intelligently combined an evolving set of missions can be created that serve both cargo and passenger missions. The empty 6 km/s allows direct Earth return on a hohmann transfer of an empty cargo mission, if the 100 mT cargo is ISRU equipment which deploys to the surface, pumps propellants into the MCT and then remains behind on launch, then a propellant farm is built up without abandoning any vehicles and the return capability is well validated before humans are sent.
The 4.9 km/s assent with 25 mT cargo is enough to reach mars orbit AND have enough propellant for another decent with FULL cargo (800 m/s propulsion), this allows the vehicle to be a reusable tanker to mars orbit depositing 25 mT per trip of any desired mass, and to act as a rapid reusable downward cargo hauler AT THE SAME TIME.
Also it means that the MCT fully fueled in mars orbit can do rapid transit back to Earth (150 day) though this would require 12 tanker-up/cargo-down flights to get the necessary propellants in orbit (with the first flight staying in orbit to act as a depot), but this is consistent with the expected 10:1 crew/cargo ratio. The very first exploration missions wouldn't have so many MCT's or propellants available and will refuel via SEP tugs sent to mars with propellants. Only once a high rate of propellant production is in place could this many MCT's be refueled and relaunched.
This will also allow the unloading of cargo from an in-space-only freight vessel while simultaneously refilling that vessel with propellant for Earth return. These large freight vehicles would be a later addition to the system and would greatly increase the cargo delivery rate and it's efficiency, but are not necessary for the initial deployment as cargo can be sent in the MCT directly all be it with poor amortization.
The 3.7 km/s when fully loaded allows the vehicle to make a fast transit to mars (150 day) from a high orbit (EML1 followed by lunar and Earth slingshot burns) while retaining enough propellants to land at mars. It would be placed in high orbit and sent propellants by a SEP tug in the 1 MW power range, if we were sending cargo by slow hohmann transfer then only 1/6th of the full propellant load (~50 mT) is needed at EML1 which may be low enough to just have it carried in the MCT from LEO eliminating propellant transfer for such cargo missions.
It also makes the fully loaded MCT ideal for making the fast (3-5 day) transfer from LEO to EML1 to quickly ferry passengers when the numbers are great enough to make use of Dragon or other small taxi craft inconvenient. Being such a short transit the accommodations can be far more cramped then would be possible for the full transit and I expect some from of transit-hab will be employed. The MCT in question will likely be docked to and towed by the transit vehicle which would transfer the necessary landing propellants and passengers back to the MCT which would land and disembark the passengers at mars using the same short term accommodations.
It is all an extremely elegant arrangement, at each cargo loading point the vehicle has sufficient DeltaV for the critical mission legs it would actually face at that point while dove-tailing well with force-multiplier vehicles that would be added later like SEP tugs.
Burninate: Terminal velocity at Mars is much, much lower than 1.5km/s for any reasonable ballistic coefficient and for any of the low altitude landing sites that are likely to be used. For Dragon-like ballistic coefficient of 300kg/m^2, you have about 350m/s terminal velocity. And terminal velocity is proportional to the square root of ballistic coefficient, so even if you think I'm wrong about ballistic coefficient, it won't make much difference.
That is from high speed direct entry and low L:D ratio, I am proposing just the opposite.That's like a high speed direct entry straight into the planet and zero L/D ratio, and probably a relatively high altitude.
100 mT Cargo + 300 mT propellant + 75 mT dry mass -> 3.7 km/s
25 mT Cargo + 300 mT propellant + 75 mT dry mass -> 5.1 km/s
0 mT Cargo + 300 mT propellant + 75 mT dry mass -> 6 km/s
SEP is not a difficult technology and I propose HALL thrusters which SpaceX is going to produce for it's own satellite swarm, the Solar panels are likewise a well established technology, the only challenge is scale
I estimated all tanks and propellant lines at 5% of the propellant mass for a total of 15 mT,
and 6 mT of engines.
their are lots of other parasitic masses involved in such a vehicle, it is not just a big rocket stage with single digit dry mass fraction.
I really can't see SEP being used. [....]
- the LV is already supposed to be too big for existing infrastructure, so the marginal cost of making it even larger may be relatively small
I really can't see SEP being used. [....]
- the LV is already supposed to be too big for existing infrastructure, so the marginal cost of making it even larger may be relatively small
I'm not seeing your reasoning here. I would think that transferring much of the TMI and TEI requirements onto a separate SEP vehicle would allow the MCT (demoted to a lander) to be smaller, not larger. Ditto its Earth launcher.
MCT would be a heck of a lot better propellant taxi and tanker if it had good mass fraction.
My objection is in creating/development an extra vehicle not the base technology itself needing development.
It's well withi [EDIT: well within their capabilities, certainly.]
That seems high. IIRC FH side booster is more like a mass ratio of 30 (at least that's what's been quoted on this forum) and that includes engines, etc.
Assuming a TWR of 100:1 or better that seems reasonable
Certainly. It needs a heat shield, it needs landing legs and a cargo container, it needs better communication than a rocket stage so it can talk to Earth from Mars.
But I don't think those will force them to limit the delta-v drastically. They'll just end up with a huge vehicle.
(Passenger version needs a lot more stuff but that will come out of the cargo-version max capacity...)
The DC-X was specifically not designed to demonstrate the most critical issue for any practical SSTO: structural mass fraction
First off you are the one repeatedly trying to argue from historical vehicles...Because if you're trying to prove something is possible, a single example from history is sufficient. If you're trying to prove something isn't possible, historical analogy isn't useful.
First off you are the one repeatedly trying to argue from historical vehicles, pointing out various metrics of performance and claiming they are achievable all in a single vehicle without consideration of how they conflict.Okay, show that you cannot achieve a rocket that does a 10:1 mass fraction with available materials and using chemical propulsion.
Now with that line of logic thoroughly refuted your now abdicating any connection with real world materials or systems, and claim I need to show that the laws of physics alone invalidate your claims?
This is preposterous and insulting because you know very well that their is no first principle of physics that limits a vehicles DeltaV, only the speed of light is denied to us by the laws of physics.
You can always propose exponentially smaller and smaller dry mass fractions and higher and higher T:W ratio engines made from any theoretically possible material and any theoretically possible propellant we can imagine that will make any DeltaV achievable. An anti-matter rocket made of neutronium is certainly allowed by the laws of physics but has no place in this this thread....
Where do you get this belief that the burned of proof is on me?...Because you claim it's not possible. Not just unfeasible, but not possible.
Where do you get this belief that the burned of proof is on me? You are the one claiming cutting edge performance on par with what a SSTO vehicle needs. YOU need to prove that envelope pushing is possible and can be done at reasonable costs, I am STILL waiting for your numbers that you hinted at several posts ago.No I'm not. I don't claim to think I can predict MCT, just establish some bounds about what COULD be done.
You are basically admitting that your logical process consists of pushing every performance margin to to braking point and claiming THAT will be the MCT design
.while I am intentionally try to find a conservative design that would not require as much 'cleverness' (all though far from none).......that line of thinking would never have produced Falcon 9 v1.1 or Full Thrust, which is necessary to do their partial reusability.
Where do you get this belief that the burned of proof is on me? You are the one claiming cutting edge performance on par with what a SSTO vehicle needs. YOU need to prove that envelope pushing is possible and can be done at reasonable costs, I am STILL waiting for your numbers that you hinted at several posts ago.No I'm not. I don't claim to think I can predict MCT, just establish some bounds about what COULD be done.
You are basically admitting that your logical process consists of pushing every performance margin to to braking point and claiming THAT will be the MCT designQuote.while I am intentionally try to find a conservative design that would not require as much 'cleverness' (all though far from none).......that line of thinking would never have produced Falcon 9 v1.1 or Full Thrust, which is necessary to do their partial reusability.
I said it'd take time. I may get to it this weekend.
If you were going to try to use earth based requirements to help fund your MCT... could a tanker upper stage be made that ALSO launches a single satellite (up to FH size)?
I think he was referring to the launch vehicle, aka make an even large BFR to launch an enormous MCT which has the propellant capacity to do a direct departure to mars. In other words he thinks doubling the size of the launch rocket is worth it to avoid making a SEP tub because you know 'fewer vehicles' logic. ::)
Total cost is the metric we care about, if it be 1 vehicle or 10.
Keep in mind that tank is 100 mT larger and lacks any insulation.
Finally I expect their to be multiple smaller tanks through the vehicle to accommodate good weight distribution and avoid the cargo-holds large volume, this will cut down on the efficiency of the tank.
If you were going to try to use earth based requirements to help fund your MCT... could a tanker upper stage be made that ALSO launches a single satellite (up to FH size)?
Two reasons at least why it does not make sense. Minor, it would need development of a small fairing, which is still big and expensive. Such a fairing would otherwise not be needed. Major, as you already mentioned, the orbits will not match. Fuel will go to near equatorial LEO or to the easiest LEO reachable from the launch site. The payload will need a different inclination and transfer orbit. Changing inclination in LEO is HARD.
...
In any event this type of lunar mission, even without crew would be an excellent shake-down mission for the MCT as it would duplicate many of the aspects of a Mars mission.
Have you consider using unmanned MCT tankers in LEO and LLO to top off the MCT lunar lander's prop tanks? In case of LLO for both descent & ascent from the Lunar surface.
...
In any event this type of lunar mission, even without crew would be an excellent shake-down mission for the MCT as it would duplicate many of the aspects of a Mars mission.
Have you consider using unmanned MCT tankers in LEO and LLO to top off the MCT lunar lander's prop tanks? In case of LLO for both descent & ascent from the Lunar surface.
Have you consider using unmanned MCT tankers in LEO and LLO to top off the MCT lunar lander's prop tanks? In case of LLO for both descent & ascent from the Lunar surface.
I have thought of a fully refueled tanker and a fully fueled MCT going to TLI together. The tanker then transfers its fuel to the MCT and does a direct return to earth while the MCT goes to land and relaunch from the moon. It's the most fuel efficient mission profile. The tricky part is rendezvous and refuelling on the way to the moon.
I have thought of a fully refueled tanker and a fully fueled MCT going to TLI together. The tanker then transfers its fuel to the MCT and does a direct return to earth while the MCT goes to land and relaunch from the moon. It's the most fuel efficient mission profile. The tricky part is rendezvous and refuelling on the way to the moon.
Obviously you dual launched the MCT and the tanker from a couple of nearby launch facilities at the same time. ;)
Unless you think there will only be one launch pad for the BFR. :o
...
In any event this type of lunar mission, even without crew would be an excellent shake-down mission for the MCT as it would duplicate many of the aspects of a Mars mission.
Have you consider using unmanned MCT tankers in LEO and LLO to top off the MCT lunar lander's prop tanks? In case of LLO for both descent & ascent from the Lunar surface.
No that would be very inefficient. The MCT has a 75 mt dry mass which would be moved all the way to LLO to do a job that could easily be done by a mere tank with a ~5% dry mass fraction.
I would put that tank on a SEP tug which is how both the propellants and MCT will be moved from LEO to the staging point for the lunar landing.
If we need more propellant then I would just move the staging point closer to the moon as I described, this easily brings the cargo delivery mass to over the 100 mt design goal and the MCT will almost have a volume limit just like every vehicle which means we can't just arbitrarily increase useful cargo even if we have the DeltaV to push it on paper it still needs to fit into the vehicle.
Speculation about 'MCT Tanker' is in my opinion misguided, the vehicle would be a terrible tanker due to it's dry mass which is highly specialized for other functions. The LEO tanker will be a stretched 2nd stage which will transfer propellant to SEP tugs which will move propellants beyond LEO.
I have thought of a fully refueled tanker and a fully fueled MCT going to TLI together. The tanker then transfers its fuel to the MCT and does a direct return to earth while the MCT goes to land and relaunch from the moon. It's the most fuel efficient mission profile. The tricky part is rendezvous and refuelling on the way to the moon.
Obviously you dual launched the MCT and the tanker from a couple of nearby launch facilities at the same time. ;)
Unless you think there will only be one launch pad for the BFR. :o
They can take a month or two to launch and refuel both in LEO, just the way they would assemble a Mars fleet to launch into a launch window. What's the problem?
They can take a month or two to launch and refuel both in LEO, just the way they would assemble a Mars fleet to launch into a launch window. What's the problem?
Was thinking of a LEO prop depot for the refueling. So if you dual launch, you minimized the time needed to complete the mission.
You way is also workable.
Since you will need the tankers in whatever form for the Methane & Lox propellants for Mars missions anyway. Using them even inefficiently might be a better trade than developing a separate SEP inspace vehicle with a different propellant (Xenon) that is not readily available in large quantities.
IMO the MCT tanker will not be too different from the MCT. Otherwise you will be developing yet another vehicle that can be reuse. Presuming you version of the MCT tanker is not expendable.
It is cheaper to developed a single general purpose vehicle than 3 different specialized vehicles in parallel.
[snip]
IMO the MCT tanker will not be too different from the MCT. Otherwise you will be developing yet another vehicle that can be reuse. Presuming you version of the MCT tanker is not expendable.
It is cheaper to developed a single general purpose vehicle than 3 different specialized vehicles in parallel.
Tankers are needed but they are hardly a 3rd vehicle, they are just stretched 2nd stages with a nose cone covering a Xenon/Krypton tank and a hose port. This can launch and deliver 100 mt of propellants of varying combinations, either all Metho-Lox from residuals or SEP propellants.
The SEP tugs would have an integrated Xenon/Krypton tank for it's own use or around 50 mt and a Metho-Lox tank for offloading to the MCT, around 150 mt would be sufficient. These tanks should mass 5 and 2.5 mt respectively and the rest of the vehicle would only come out to around 10 mt meaning it can be launched mostly fueled on the BFR.
A vehicle this size is going to be a lot simpler to design and produce then the MCT itself which goes through incredible stresses and flight regimes. The SEP is basically made of off the shelf satellite parts, Solar panels and Hall thrusters both of which are going to be mass produced by SpaceX in it's Satellite Swarm.
Two tankers refuel 1 SEP with MethoLox and Xenon to make a delivery of Metho-Lox to EML1. One tanker refuels two SEPs with just Xenon to move two MCT's to EML1. That's 3 fuel launches and 2 MCT so a 2:5 ratio for cargo delivery. Nearly doubling the LEO to Mars efficiency at the cost of a small SEP tug that should be a fraction of the cost of the vehicles that are already necessary.
A single-stage BFR is not likely to be a good cargo or liquid ferry to LEO, much less to GTO. Payload mass fractions on a single stage are very small. You can expect a *reusable* single stage rocket-based launch vehicle to be extraordinarily difficult to do with any payload at only 360-380s Isp.
essentially a reusable second stage with a long-loiter package and refueling capabilityfor less money then just making exactly THAT from a clean sheet design with a common diameter and tank fabrication process shared with the first stage.
Are you on L2, Burninate?
Direct quote from Elon implying no second stage, just booster and mars spacecraft from an article in GQ:
"Well, there's two parts of it—there's a booster rocket and there's a spaceship. So the booster rocket's just to get it out of Earth's gravity because Earth has quite a deep gravity well and thick atmosphere, but the spaceship can go from Mars to Earth without any booster, because Mars's gravity is weaker and the atmosphere's thinner, so it's got enough capability to get all the way back here by itself. It needs a helping hand out of Earth's gravity well. So, technically, it would be the BFR and the BFS." As in "Big frakking Spaceship."
http://www.gq.com/story/elon-musk-mars-spacex-tesla-interview?utm_source=10370
A BFR can be conceived with lots less than 30 1st stage engines.
As a point of reference, I’ve taken these Musk’s statements as a given even though I bet they’ll be modified at his supposed late 2015 (likely spring 2016) briefing.
100mT land on Mars cargo
Land the whole thing and re-use it; i.e. return it to Earth
“there's a booster rocket and there's a spaceship. So the booster rocket's just to get it out of Earth's gravity because Earth has a deep gravity well and thick atmosphere, but the spaceship can go from Mars to Earth without any booster”
Raptor thrust over 230mT; use a lot of them.
380 seconds ISP vacuum
My BFR is 12.5m diameter, making a relatively squat first stage under 30m tall. Propellant tanks alone 23.5m. Delivers ~2.8Km/sec after gravity losses. The wider diameter leaves room for growth for later larger versions, allowing for lengthening the tanks in both stages, adding engines and lengthening the cargo hold. Assumed a heavy dry weight of 90mT for stage 2 MCT.
No worries about pad towers for 100m high rockets. BFR with MCT 2nd stage is shorter than the 70m F9 but 10 times more massive. BFR’s multi Raptor engine driven 12.5m diameter is a nice size to fit living quarters and various colonial heavy equipment.
Don’t need 15 million LBS thrust, but with mass ratios ~4.3% need a bit over 12 million LBS (54 million Newtons) thrust, 24 engines arranged in rings of 16 and 8 with room for a future center engine or 2.
Given the huge delta V requirements for both Mars departure from refueling at LEO AND later functioning as a SSTO taking off from Mars’ surface and return to Earth, I put the Km/sec budget into the 2nd stage. Stage one goes low & slow, under 3 Km/sec, boosting the heavy 2nd stage before return to launch site, RTLS. The dry weight plus fuel of the returning 1st stage exceeds Raptor thrust so any 2 of the 8 engine inner ring engines throttled down provide landing thrust.
Total BFR mass 4450mT or 9.8 million LBS. LEO mass fraction 4.3%.
Stage One:
12.5m diameter with 23.5m length propellant tanks
3280mT 7.2 million LBS 1st stage fueled mass
230mT thrust engines 506K LBS
24 engines 56 million Newtons 12.2 million LBS Thrust; T/W 1.24
Rings of 16 engines, and 8 engines
Avg ISP from sea level to vacuum 325
After 1 Km/sec additional Delta V reserved for RTLS Rocket Equation gives 3.4 Km/sec but nets under 3 Km/sec Delta V after gravity losses
Stage Two The MCT:
Dry Mass 90mT; 100mT is cargo
12.5m diameter with 7.5 m length for propellant tanks.
Cargo 8.5m length ; 1040 m3 volume
(Plenty of space for expanded fuel tanks in a simple modified tanker version.)
1165mT fueled mass (2.6 million LBS)
380 seconds ISP vac; Rvac engines assumed 14% higher thrust as with F9 FT
6 Rvac engines 3.5 million LBS Thrust
6.75Km/sec Delta V capability, via Rocket eq.
8.4Km/sec Mars liftoff with only 25mT return cargo
I also modeled another small version to see just how small a BFR could meet Elon’s goals. I optimistically lowered the MCT 2nd stage dry weight from 90mT to 75mT. This minimal, 12m diameter optimistic version needs only 19 engines with rings of 12, 6 and two center engines. The MCT stage 2 has only 5 engines, reducing the cost of building this BFR/MCT from 30 Raptors to 24. Engines are the big cost driver for mass producing BFRs.
From another thread, here is a recent quote by Musk in an interview:I'd just read that article and came here.
"So, technically, it would be the BFR and the BFS." As in "Big frakking Spaceship.""
Just to confirm: you have the BFR contributing only 1km/s delta V before separation?
philw1776: Your engine count is off because you forgot to account for lower thrust at Sea-level, 230 mt is the VAC performance goal for Raptor. I'm estimating that Sea-level thrust is ~77% of Vac which raises the engine count to ~30 to get very nearly identical total thrust as your estimate.
Q: Has the Raptor engine changed in its target thrust since the last number we have officially heard of 1.55Mlbf SL thrust?
A: Thrust to weight is optimizing for a surprisingly low thrust level, even when accounting for the added mass of plumbing and structure for many engines. Looks like a little over 230 metric tons (~500 klbf) of thrust per engine, but we will have a lot of them :)
philw1776: Your engine count is off because you forgot to account for lower thrust at Sea-level, 230 mt is the VAC performance goal for Raptor. I'm estimating that Sea-level thrust is ~77% of Vac which raises the engine count to ~30 to get very nearly identical total thrust as your estimate.
A question. Do we know that the 230mt quoted (I assume in the reddit Q&A Elon did) was for Vac? Or has there been any more information on this elsewhere? I'm asking because Elon was answering a question that was about SL thrust.
The quote I remember.QuoteQ: Has the Raptor engine changed in its target thrust since the last number we have officially heard of 1.55Mlbf SL thrust?
A: Thrust to weight is optimizing for a surprisingly low thrust level, even when accounting for the added mass of plumbing and structure for many engines. Looks like a little over 230 metric tons (~500 klbf) of thrust per engine, but we will have a lot of them :)
It would be amazing to one day see hardware with a giant 'SpaceX BFR' Decal on the side. I wonder if the 'name' will survive that long... Of it gets renamed something else... 'Tiny' maybe.
I understand that usually it is assumed that the first stage will give the upper stage/MCT a horizontal speed of ~3km/s, similar to what the Falcon first stage does. For RTLS of the first stage that needs to be reversed, cutting into payload.
I do wonder if another approach could be effective. MCT will have a very large delta-v budget. It needs it to perform its functions towards Mars. Could the first stage go up almost straight similar to the BO New Shepard, but maybe up to 150km peak altitude, eating all the gravity and air resistance losses and use the second stage for the task of building up orbital speed? On the way down with its large diameter it may not need a reentry burn or only a very small one. Reuse fuel would be mainly only the small amount of landing fuel.
Did anyone of those who did thorough analysis ever consider such a scenario or am I way off?
The vast majority of the energy in a rocket launch is expended reaching orbital velocity, not orbital height. Getting to orbital velocity requires a large amount of horizontal velocity. Getting that horizontal velocity takes about the same amount of dV, no matter what altitude you are at. If you started at LEO altitude, and started accelerating horizontally, you'd still need to obtain 17,100 mph of horizontal velocity. Without the horizontal velocity, you're suborbital.
The vast majority of the energy in a rocket launch is expended reaching orbital velocity, not orbital height. Getting to orbital velocity requires a large amount of horizontal velocity. Getting that horizontal velocity takes about the same amount of dV, no matter what altitude you are at. If you started at LEO altitude, and started accelerating horizontally, you'd still need to obtain 17,100 mph of horizontal velocity. Without the horizontal velocity, you're suborbital.
I know all of this. I even explicitly stated it in my post. So what's your point?
Usually, I understand, a first stage does a lot of delta-v towards orbital speed. The Falcon 9 does much less, to facilitate RTLS and leaves more of the buildup of orbital speed to the second stage. That is not the optimum approach for expendable vehicles. It is a better approach for reusable vehicles.
My suggestion was to carry the idea to the extreme. Let the first stage eat all the gravity loss and drag loss and leave all or almost all of the buildup of orbital speed to the second stage. MCT needs the big delta-v budget anyway to get to Mars from LEO and land.
"Big Falcon Spaceship"
I know all of this. I even explicitly stated it in my post. So what's your point?
Usually, I understand, a first stage does a lot of delta-v towards orbital speed. The Falcon 9 does much less, to facilitate RTLS and leaves more of the buildup of orbital speed to the second stage. That is not the optimum approach for expendable vehicles. It is a better approach for reusable vehicles.
My suggestion was to carry the idea to the extreme. Let the first stage eat all the gravity loss and drag loss and leave all or almost all of the buildup of orbital speed to the second stage. MCT needs the big delta-v budget anyway to get to Mars from LEO and land.
. Let the first stage eat all the gravity loss and drag loss and leave all or almost all of the buildup of orbital speed to the second stage. MCT needs the big delta-v budget anyway to get to Mars from LEO and land.
. Let the first stage eat all the gravity loss and drag loss and leave all or almost all of the buildup of orbital speed to the second stage. MCT needs the big delta-v budget anyway to get to Mars from LEO and land.
?? gravity is pretty much the same at 160km up as it is on the surface. Gravity losses starting from 0 or 160km up are very similar. It's all about the orbital velocity.
The point I tried to make is that this is a lot of delta-v needed. MCT might be able to do it, if its designed to stage from LEO, but in that case it might be almost empty when it reaches orbital speed. You will then need to transfer more fuel to it so that it can start the mission to Mars.
At the same time, the "slingshot to orbital height S1" might make re-usability easier, but you would still have to design it so that it could carry other payloads than MCT (for example, a tanker, or a hub, or mars infrastructure, or commercial payloads etc). If S1 is designed to launch more things than strictly MCT, then it would not be easy (I think) to merge both capabilities in the same structure.
. Let the first stage eat all the gravity loss and drag loss and leave all or almost all of the buildup of orbital speed to the second stage. MCT needs the big delta-v budget anyway to get to Mars from LEO and land.
?? gravity is pretty much the same at 160km up as it is on the surface. Gravity losses starting from 0 or 160km up are very similar. It's all about the orbital velocity.
As I have already stated twice, I know that it is all about orbital velocity. I also know that gravity is almost as strong at 160km altitude. But I was not talking about gravaity, but about gravity losses. The difference is the vector at which the engines fire. However the first stage would fire vertical or almost vertical while the second stage could fire horizontal.
I will just wait for the first flights. I expect that MCT will stage at less than 3km/s, but we will see.
The one design criteria most overlooked is minimum operating costs. This is probably the most significant item for the design and delta -V requirements that the BFR must meet. These operating costs are for sending an MCT to Mars meaning less tanker flights means much lower costs. An increase of the BFR's deliverable delta V by 750m/s increases the tanker prop load from 100mt to 200mt. This nearly halves the total cost for an MCT to Mars by almost 50% from 10 flights to only 5 flights. This is without any change to the design of the BFS itself. The BFS is fairly easily sized because the three design flight uses: as second stage, as EDS, and as Mars SSTO all are very close to the same propellant loads requirements.
. Let the first stage eat all the gravity loss and drag loss and leave all or almost all of the buildup of orbital speed to the second stage. MCT needs the big delta-v budget anyway to get to Mars from LEO and land.
?? gravity is pretty much the same at 160km up as it is on the surface. Gravity losses starting from 0 or 160km up are very similar. It's all about the orbital velocity.
Is it still inefficient if you take in to account the 1st stage needs to RTLS, and therefor needs some sort of burn to negate it's horizontal DV? In fact it needs to impart twice the horizontal DV to rtls as it used to get to separation. It's carrying less fuel so is a lot lighter however, so maybe its not that big of an ask.
As I have already stated twice, I know that it is all about orbital velocity. I also know that gravity is almost as strong at 160km altitude. But I was not talking about gravaity, but about gravity losses. The difference is the vector at which the engines fire. However the first stage would fire vertical or almost vertical while the second stage could fire horizontal.
I will just wait for the first flights. I expect that MCT will stage at less than 3km/s, but we will see.
"less than 3km/h" is much more than almost zero horizontal velocity.
Making the first stage launch vertically and just eat the gravity and atmospheric losses would mean that
1) second stage would have to do all the work for the 7.5km/h horizontal delta-v
2) The trajectory of the rocket would be very inefficient. Doing the vertical and horizontal part of the acceleration totally separately means much more total delta-v is needed, pythagoras is our friend here. Just to get to 200km altitude 2km/s delta-v is needed for the first stage. And this is ignoring the gravity losses. With 45 degrees burn direction, 2.8km/s delta-v gives 2kms/s vertical AND 2km/s horizontal delta-v. The gravity turns real rockets are doing are even much better.
Please, try this in KSP. You will see that how inefficient your trajectory idea is.
Making the first stage launch vertically and just eat the gravity and atmospheric losses would mean that
...
The one design criteria most overlooked is minimum operating costs. This is probably the most significant item for the design and delta -V requirements that the BFR must meet. These operating costs are for sending an MCT to Mars meaning less tanker flights means much lower costs. An increase of the BFR's deliverable delta V by 750m/s increases the tanker prop load from 100mt to 200mt. This nearly halves the total cost for an MCT to Mars by almost 50% from 10 flights to only 5 flights. This is without any change to the design of the BFS itself. The BFS is fairly easily sized because the three design flight uses: as second stage, as EDS, and as Mars SSTO all are very close to the same propellant loads requirements.
I'm certainly not convinced that SpaceX will use an obscenity when they finally introduce the BFR. I suspect they will continue to (at least occasionally) use Falcon in public, if they give any expansion for BFR at all.I don't think it'll ever be an official name, but given Musk's personality I'd say it was definitely that at first. There'll be a rename, unless "Big Falcon" sticks.
Direct quote from Elon implying no second stage, just booster and mars spacecraft from an article in GQ
I beleive you are misinterpreting him simplifying things for the magazine readers. He seems to just be saying there will be a independant spacecraft and rocket. Just like there is a Dragon spacecraft and Falcon "booster" rocket. Thats just my take on it though. I could be wrong and perhaps they have found a reliable supplier of unobtainium and dilithium crystals from which to build this single stage booster rocket out of.
1) Given that the dV needed to pull a SSTE from Mars surface is about 8 km/s, I'd say you're set in the dV department.. Let the first stage eat all the gravity loss and drag loss and leave all or almost all of the buildup of orbital speed to the second stage. MCT needs the big delta-v budget anyway to get to Mars from LEO and land.
?? gravity is pretty much the same at 160km up as it is on the surface. Gravity losses starting from 0 or 160km up are very similar. It's all about the orbital velocity.
As I have already stated twice, I know that it is all about orbital velocity. I also know that gravity is almost as strong at 160km altitude. But I was not talking about gravaity, but about gravity losses. The difference is the vector at which the engines fire. However the first stage would fire vertical or almost vertical while the second stage could fire horizontal.
I will just wait for the first flights. I expect that MCT will stage at less than 3km/s, but we will see.
"less than 3km/h" is much more than almost zero horizontal velocity.
Making the first stage launch vertically and just eat the gravity and atmospheric losses would mean that
1) second stage would have to do all the work for the 7.5km/h horizontal delta-v
2) The trajectory of the rocket would be very inefficient. Doing the vertical and horizontal part of the acceleration totally separately means much more total delta-v is needed, pythagoras is our friend here. Just to get to 200km altitude 2km/s delta-v is needed for the first stage. And this is ignoring the gravity losses. With 45 degrees burn direction, 2.8km/s delta-v gives 2kms/s vertical AND 2km/s horizontal delta-v. The gravity turns real rockets are doing are even much better.
Please, try this in KSP. You will see that how inefficient your trajectory idea is.
Direct quote from Elon implying no second stage, just booster and mars spacecraft from an article in GQ
I beleive you are misinterpreting him simplifying things for the magazine readers. He seems to just be saying there will be a independant spacecraft and rocket. Just like there is a Dragon spacecraft and Falcon "booster" rocket. Thats just my take on it though. I could be wrong and perhaps they have found a reliable supplier of unobtainium and dilithium crystals from which to build this single stage booster rocket out of.
The range of LEO capability given by SpaceX way back for the MCT was 180-250mt. As we have discussed tremendous about what the BFS dry weight is possible 80mt is sort of a average or consensus value + 100mt of payload, making the 180mt to LEO the absolute minimum that the system must meet. But what if the performance was closer to the other end 250mt. That means that about 70mt of extra propellant is delivered per flight.The one design criteria most overlooked is minimum operating costs. This is probably the most significant item for the design and delta -V requirements that the BFR must meet. These operating costs are for sending an MCT to Mars meaning less tanker flights means much lower costs. An increase of the BFR's deliverable delta V by 750m/s increases the tanker prop load from 100mt to 200mt. This nearly halves the total cost for an MCT to Mars by almost 50% from 10 flights to only 5 flights. This is without any change to the design of the BFS itself. The BFS is fairly easily sized because the three design flight uses: as second stage, as EDS, and as Mars SSTO all are very close to the same propellant loads requirements.
Could not agree more.
I was in a high-tech industry engineering products where low production cost was paramount. We did the same cost driver analysis on every single aspect that Musk does.
I see the 1st stage as going low & slow under 3 Km/sec while the delta V is in the MCT stage two which makes it utilitarian for all Mars purposes. Large propellant tanks make 'mods" for a tanker minimal.
I also think those speculating that the BFR might be larger than minimal models may be onto something as they reduce # of flights to refuel the MCT in LEO or wherever. It's a complex system analysis whether to make a minimum parts cost re-useable BFR vs making one a bit larger (a bit more expensive per unit) that reduces # of flights to fuel up an MCT in orbit for a Mars journey.
Direct quote from Elon implying no second stage, just booster and mars spacecraft from an article in GQ
I beleive you are misinterpreting him simplifying things for the magazine readers. He seems to just be saying there will be a independant spacecraft and rocket. Just like there is a Dragon spacecraft and Falcon "booster" rocket. Thats just my take on it though. I could be wrong and perhaps they have found a reliable supplier of unobtainium and dilithium crystals from which to build this single stage booster rocket out of.
Direct quote from Elon implying no second stage, just booster and mars spacecraft from an article in GQ
I beleive you are misinterpreting him simplifying things for the magazine readers. He seems to just be saying there will be a independant spacecraft and rocket. Just like there is a Dragon spacecraft and Falcon "booster" rocket. Thats just my take on it though. I could be wrong and perhaps they have found a reliable supplier of unobtainium and dilithium crystals from which to build this single stage booster rocket out of.
Their is a large group of people who are simply dead-set on this 'super-direct' architecture and always have been and will interpret everything as confirmation of that architecture, no amount of technical arguments on my part about it's in-feasibility, enormous cost and risk have dissuaded them.
From another thread, here is a recent quote by Musk in an interview:Let me just say, in a totally mature manner (regarding BFS--what we used to call MCT--being the integrated second stage):Direct quote from Elon implying no second stage, just booster and mars spacecraft from an article in GQ:
"Well, there's two parts of it—there's a booster rocket and there's a spaceship. So the booster rocket's just to get it out of Earth's gravity because Earth has quite a deep gravity well and thick atmosphere, but the spaceship can go from Mars to Earth without any booster, because Mars's gravity is weaker and the atmosphere's thinner, so it's got enough capability to get all the way back here by itself. It needs a helping hand out of Earth's gravity well. So, technically, it would be the BFR and the BFS." As in "Big frakking Spaceship."
http://www.gq.com/story/elon-musk-mars-spacex-tesla-interview?utm_source=10370
So... I guess we are back to basics, what many of us have arguing, it seems? This thread has been devalued (and made less interesting) in the last few months by people wanting to discuss their own architectures (yes, you know who you are), so hopefully this can narrow down the discussion again.
The range of LEO capability given by SpaceX way back for the MCT was 180-250mt. As we have discussed tremendous about what the BFS dry weight is possible 80mt is sort of a average or consensus value + 100mt of payload, making the 180mt to LEO the absolute minimum that the system must meet. But what if the performance was closer to the other end 250mt. That means that about 70mt of extra propellant is delivered per flight.The one design criteria most overlooked is minimum operating costs. This is probably the most significant item for the design and delta -V requirements that the BFR must meet. These operating costs are for sending an MCT to Mars meaning less tanker flights means much lower costs. An increase of the BFR's deliverable delta V by 750m/s increases the tanker prop load from 100mt to 200mt. This nearly halves the total cost for an MCT to Mars by almost 50% from 10 flights to only 5 flights. This is without any change to the design of the BFS itself. The BFS is fairly easily sized because the three design flight uses: as second stage, as EDS, and as Mars SSTO all are very close to the same propellant loads requirements.
Could not agree more.
I was in a high-tech industry engineering products where low production cost was paramount. We did the same cost driver analysis on every single aspect that Musk does.
I see the 1st stage as going low & slow under 3 Km/sec while the delta V is in the MCT stage two which makes it utilitarian for all Mars purposes. Large propellant tanks make 'mods" for a tanker minimal.
I also think those speculating that the BFR might be larger than minimal models may be onto something as they reduce # of flights to refuel the MCT in LEO or wherever. It's a complex system analysis whether to make a minimum parts cost re-useable BFR vs making one a bit larger (a bit more expensive per unit) that reduces # of flights to fuel up an MCT in orbit for a Mars journey.
The whole reason for the larger BFR and less flights would be the same regardless of whether the BFR is reusable or expendable. If the minimum design BFR takes 10 flights to accomplish sending a single BFS to Mars but a bigger BFR with the exact same design BFS that takes only 5 flights although the BFR costs 50% more per flight still gives a reduction to per mission of 75% of the 10 flights configuration. There are other advantages to requiring half the flights and that is pad availability. In order to support sending 4 BFS's to Mars the 10 flights minimal BFR configuration would require 40 launches in the 780 day period (a launch every 19.5 days). In order to support sending 4 BFS's to Mars the 5 flights large BFR configuration would require 20 launches in the 780 day period (a launch every 39 days). For the 10 cargo to 1 crew ratio of missions making the possibility of 20 cargo and 2 crew missions in a synod (780 day period) that would require
a) 10 flight config-> a launch every 3.5 days
b) 5 flight config-> a launch every 7 days
As mission counts increase the number of launches becomes a greater cost factor than any other consideration.
Could the first stage go up almost straight similar to the BO New Shepard, but maybe up to 150km peak altitude, eating all the gravity and air resistance losses and use the second stage for the task of building up orbital speed? On the way down with its large diameter it may not need a reentry burn or only a very small one. Reuse fuel would be mainly only the small amount of landing fuel.
While the entry speed may be technically lower than a 3km/s horizontal entry, the rate of atmospheric density increase will be extremely sharp. That induces stresses on the stage in addition to mere re-entry heating.
(Travelling 100km through the first ten kilometres of atmosphere, then 100km through the second 10km... vs travelling 30km through the first 30km.... See what I mean. 200km of deceleration before you reach 30km altitude, vs just 30km deceleration and you're already deep in the atmosphere at the 20km mark.)
Just re-read parts of the GQ article. Here is a bit that is just below the previous quote that I think belongs in this thread -
"Musk has previously said that he would publicly present some specifics of his Mars-colonization plans later this year, though he tells me that it may now be early next year. "Before we announce it, I want to make sure that we're not gonna make really big changes to it," he says. "Um, yeah. I think it's gonna seem pretty crazy, no matter what."
Just because it's so far beyond what people would imagine?
He laughs. "It's really big." And laughs again. "It's really big. There's not been any architecture like this described that I'm aware of."
That's from December 2015.
SpX Mars suit worn by Elon Musk?
BFS in the background?
https://www.instagram.com/elonmusk/?hl=en
Is this a Martian rescue fleet arriving after a Earth comet slam?
SpX Mars suit worn by Elon Musk?
BFS in the background?
https://www.instagram.com/elonmusk/?hl=en
Is this a Martian rescue fleet arriving after a Earth comet slam?
Nice catch! That must being showing Musk on Earth though with his helmet off, unless he figures he'll be able to terraform Mars before he gets there.
Also, the vehicle in the background has aerodynamic features, so it's not just ballistic like the Falcon 9 1st stage.
Of course all this assumes that the drawing reveals some real details, but it could be pure fantasy...
Doesn't look like the SpaceX suit design to me. Doesn't particularly look like Elon, either, for that matter.
The only thing that is familiar is the Dragon-like canted thruster arrangement.
It's that quote that I keep thinking of when people start saying what they think the architecture will be, big capsules etc. I don't think anyone has yet described an architecture that would fit with Musk's statement.
1) Um... why?
It's that quote that I keep thinking of when people start saying what they think the architecture will be, big capsules etc. I don't think anyone has yet described an architecture that would fit with Musk's statement.
I'm just going to suggest an outrageous one. The MCT is an Earth SSTO vehicle sized near the weight limit of LC39A that can deliver 236mTons to LEO. With a BFR under it it reaches orbit high on fuel. Fuel capacity is significantly higher than required for the trip to Mars, so one BFR flight can fuel up more than one MCT. Early flights can de-risk ISRU by leaving an MCT in Mars orbit with fuel for TEI and landing with enough fuel to get back to orbit.
I think that's probably big enough.
SpX Mars suit worn by Elon Musk?
BFS in the background?
https://www.instagram.com/elonmusk/?hl=en
Is this a Martian rescue fleet arriving after a Earth comet slam?
1) Um... why?
It's that quote that I keep thinking of when people start saying what they think the architecture will be, big capsules etc. I don't think anyone has yet described an architecture that would fit with Musk's statement.
I'm just going to suggest an outrageous one. The MCT is an Earth SSTO vehicle sized near the weight limit of LC39A that can deliver 236mTons to LEO. With a BFR under it it reaches orbit high on fuel. Fuel capacity is significantly higher than required for the trip to Mars, so one BFR flight can fuel up more than one MCT. Early flights can de-risk ISRU by leaving an MCT in Mars orbit with fuel for TEI and landing with enough fuel to get back to orbit.
I think that's probably big enough.
2) Deliver 236 tons payload?
2A) Assuming 'yes', and you mean 236 tons of bricks to orbit on top of the rocket:EDIT: Actually, the Shuttle tank's structural mass fraction should be higher (worse) than the Falcon tank according to first principles; This is an apples to oranges comparison because the Shuttle External Tank had no engines. Tank mass fraction is supposed to scale inversely with propellant density; The density of methalox is lower than the density of RP-1-LOX. Perhaps I should rerun this with 7% or 8%.
The structural mass fraction of the F9 first stage is currently estimated here at 6.1% (http://spaceflight101.com/spacerockets/falcon-9-v1-1-f9r/). The structural mass fraction of the Shuttle External Tank (SLWT edition) is estimated here at 3.5% (https://en.wikipedia.org/wiki/Space_Shuttle_external_tank). Let's split the difference and call the estimated rocket structural mass fraction 5%.
Assuming 5% structural mass fraction on the rocket:
The burn mass ratio for a liftoff burn of 9.2km/s at an average Isp of 370s, will be about 12.6:1 wet to dry.
Plug that into here (http://www.quantumg.net/rocketeq.html) with a 9200m/s LEO, and you need 7143mt launch mass to reach orbit with 236mt of payload atop an empty rocket with calculated mass 329mt. Add about 25% for liftoff thrust margin, and you need 87.5MN liftoff thrust. Raptor thrust is ~= 2250kN; That makes for 39 Raptors.
LC-39A is said to be designed to 12.5Mlbf == 55.6MN.
2B) Assuming 'no', and you meant 236t total mass to orbit:
Plugging the same parameters in (with a little iteration to work the linked calculator backwards) but with an m1 (total mass to orbit) of 236t, you get 98.6mt payload mass atop a 137.4mt empty rocket to orbit, using a launch mass of 2984mt. With the same coefficient, that's 36.6MN launch thrust, which is within the scope of LC-39A. That makes for 17 Raptors... but below 100mt to orbit.
I think these wildly huge BFR/MCT estimates are way, way off. The thing has to be affordable to build in quantity and has to be able to launch without evacuating the surrounding populace and rebuilding puny steel & concrete launch pads.
I think these wildly huge BFR/MCT estimates are way, way off. The thing has to be affordable to build in quantity and has to be able to launch without evacuating the surrounding populace and rebuilding puny steel & concrete launch pads.
Yeah, we keep pingponging between "But the math and some educated guesses say it has to be at least this big, even optimistically" and "But Musk said it would be $500k/ticket. Hundreds of thousands of passengers. To do that it has to be a quarter that mass and twice the speed! Build it out of unobtainium!".
I think there may be some middle ground in having one vehicle with multiple *configurations* for different purposes.
The primary bottomline variable is delta-V capability. I showed above that a 9.2km/s stage for MCT is extremely problematic, perhaps impossible depending on structural mass fraction. There are several steps (starting with ISRU and working on up) where refueling can drop the dV capability needed of the vehicle by splitting the longest leg of the mission, very substantially.
It's not that an 88-Raptor MCT is definitely impossible, it's just impractical & unnecessary; There are easier ways, lower-hanging fruit. It doesn't make *sense* to avoid propellant depots, to avoid ISRU, to avoid LEO cargo loading.
I see a GLOW of around 4,600 mt, 31 Raptor engines on the booster and a MCT/BFS of only 75 mt dry. A vastly more achievable size.
To get a lot smaller than 34,000 MT requires more smaller vehicles, rendezvous, etc. per flight.
It also likely increases total dV by increasing the number of different intermediate parking/rendezvous orbits.
In terms of affordable building in mass, WW II Liberty ships, produced cheaply in mass (2,700 built in only a few years), had deadweight over 10,000 MT. An MCT stack 34,000 MT fueled would be less than 5,000 MT structural construction. It seems like as hard as it is one giant stack is probably the best hope for $500k/person.
Umbrella, do you even read other posts, or is this a post-only thread for you?(I think they were responding to philw1776 asking about where and how this "behemoth" would be launched)
My smallest BFR/MCT is 12.5m in diameter, has 25 engines and 4630mT GLOW. Delivers 180mT to LEO at a mass fraction of 3.9% or 26:1 wet to dry mass ratio. Goes low & slow. Propellant reserve for abort to Earth landing. Built for rugged quick turn RTLS and re-launch. The 1st stage is the easier vehicle for SpaceX to design. Meets the claimed requirement of 100mT cargo to LEO.
My larger preferred BFR/MCT is 15m diameter, has 28 engines and 5100mT GLOW. Delivers 185mT to LEO at a mass fraction of 3.6% or 28:1 wet to dry mass ratio. Goes low & slow. Built for rugged quick turn RTLS and re-launch. I think Musk goes 15m diameter to allow for future engine growth and provide margin when a 85mT dry weight MCT turns out to be too optimistic with complications from things like TPS and complex extra Mars landing engines located "higher" along the structure.
The 80-85mT dry weight upper stage is "near SSTO" with 7.7Km/sec capability fully fueled. Allows for "fast" transits to Mars. Less refueling for slower cargo flights to Mars.
When you calculate the mass of propellant for a 15m vehicle you realize that the "BFR" is short and stout, the opposite of the F9 family.
Anything larger than 12m will probably have to be made on the coast. A three 8m core heavy version could be make almost anywhere with an 8m core single version replacing Falcon heavy. Three core heavy for MCT.
My smallest BFR/MCT is 12.5m in diameter, has 25 engines and 4630mT GLOW. Delivers 180mT to LEO at a mass fraction of 3.9% or 26:1 wet to dry mass ratio. Goes low & slow. Propellant reserve for abort to Earth landing. Built for rugged quick turn RTLS and re-launch. The 1st stage is the easier vehicle for SpaceX to design. Meets the claimed requirement of 100mT cargo to LEO.
My larger preferred BFR/MCT is 15m diameter, has 28 engines and 5100mT GLOW. Delivers 185mT to LEO at a mass fraction of 3.6% or 28:1 wet to dry mass ratio. Goes low & slow. Built for rugged quick turn RTLS and re-launch. I think Musk goes 15m diameter to allow for future engine growth and provide margin when a 85mT dry weight MCT turns out to be too optimistic with complications from things like TPS and complex extra Mars landing engines located "higher" along the structure.
The 80-85mT dry weight upper stage is "near SSTO" with 7.7Km/sec capability fully fueled. Allows for "fast" transits to Mars. Less refueling for slower cargo flights to Mars.
When you calculate the mass of propellant for a 15m vehicle you realize that the "BFR" is short and stout, the opposite of the F9 family.
I think that's a little optimistic on mass ratios, but I do expect that a 15m diameter vehicle is going to have a very low ballistic coeefficient, with good lift it can spend a lot of time in the upper atmosphere for a low g and thermal load entry, giving a it a very good mass fraftion for an entry vehicle. But I also expect a design that doesn't depend on a never before achieved mass ratio.
I see a GLOW of around 4,600 mt, 31 Raptor engines on the booster and a MCT/BFS of only 75 mt dry. A vastly more achievable size.
Which assumes 75mt dry mass for a reusable 100mt Mars lander is realistic...
I can see maybe overcoming the problems with a super massive launch vehicle on Earth but the ISRU requirements get out of control. Can you really make several thousand mTons of propellant on Mars for each MCT? What's the footprint for a solar array to get that done in 2 years?
It's more realistic then most other proposals which presume the lander has >1000 mt of propellant tanks at the same mass. I targeting just 300 mt for tank capacity resulting in a much smaller vehicle and a dry mass fraction of 20%.
Mass ratios are roughly equivalent to today's. Probably among the most pessimistic/conservative cited here.
Not accurate.I see a GLOW of around 4,600 mt, 31 Raptor engines on the booster and a MCT/BFS of only 75 mt dry. A vastly more achievable size.
Which assumes 75mt dry mass for a reusable 100mt Mars lander is realistic...
It's more realistic then most other proposals which presume the lander has >1000 mt of propellant tanks at the same mass. I targeting just 300 mt for tank capacity resulting in a much smaller vehicle and a dry mass fraction of 20%.
...
Not accurate.I see a GLOW of around 4,600 mt, 31 Raptor engines on the booster and a MCT/BFS of only 75 mt dry. A vastly more achievable size.
Which assumes 75mt dry mass for a reusable 100mt Mars lander is realistic...
It's more realistic then most other proposals which presume the lander has >1000 mt of propellant tanks at the same mass. I targeting just 300 mt for tank capacity resulting in a much smaller vehicle and a dry mass fraction of 20%.
...
It's more realistic then most other proposals which presume the lander has >1000 mt of propellant tanks at the same mass. I targeting just 300 mt for tank capacity resulting in a much smaller vehicle and a dry mass fraction of 20%.
Does that include your self-deployable, disposable, wire landing/launch pad?
I see a GLOW of around 4,600 mt, 31 Raptor engines on the booster and a MCT/BFS of only 75 mt dry. A vastly more achievable size.
Which assumes 75mt dry mass for a reusable 100mt Mars lander is realistic...
It's more realistic then most other proposals which presume the lander has >1000 mt of propellant tanks at the same mass. I targeting just 300 mt for tank capacity resulting in a much smaller vehicle and a dry mass fraction of 20%.
I see a GLOW of around 4,600 mt, 31 Raptor engines on the booster and a MCT/BFS of only 75 mt dry. A vastly more achievable size.
Which assumes 75mt dry mass for a reusable 100mt Mars lander is realistic...
I see a GLOW of around 4,600 mt, 31 Raptor engines on the booster and a MCT/BFS of only 75 mt dry. A vastly more achievable size.
Which assumes 75mt dry mass for a reusable 100mt Mars lander is realistic...
It's more realistic then most other proposals which presume the lander has >1000 mt of propellant tanks at the same mass. I targeting just 300 mt for tank capacity resulting in a much smaller vehicle and a dry mass fraction of 20%.
Can you find any previous estimates of the mass ratio for dry mass to cargo mass for a (presumably non-reusable) Mars lander, or alternately if your architecture lands back at Earth, for a reusable Earth lander?
I was extremely critical of this aspect in the past on the basis that a habitat which needs to hold 100 passengers for 1000 days with 1000 days of food, survive reentry, and produce its own propellant on the other end is probably going to be several times that dry mass; But I'm increasingly moving towards backing off this conjunction of requirements, and assuming they will not occur simultaneously.
Even so; Something like the Bigelow BA-2100 is supposed to be heavier than this. On what basis do you propose 75 tons instead of 9 tons or 400 tons?
It is just for launch and would be directly under the engines with a total area of ~100 m^2, even thick 6 gauge steel wire at a 1 inch spacing would be only 1,220 kg, add in stakes and simple rams that push the wire onto the ground and pin it and your still looking at a tiny fraction of the cargo capacity which is where it is accounted for.
I see a GLOW of around 4,600 mt, 31 Raptor engines on the booster and a MCT/BFS of only 75 mt dry. A vastly more achievable size.
Which assumes 75mt dry mass for a reusable 100mt Mars lander is realistic...
It's more realistic then most other proposals which presume the lander has >1000 mt of propellant tanks at the same mass. I targeting just 300 mt for tank capacity resulting in a much smaller vehicle and a dry mass fraction of 20%.
Can you find any previous estimates of the mass ratio for dry mass to cargo mass for a (presumably non-reusable) Mars lander, or alternately if your architecture lands back at Earth, for a reusable Earth lander?
I was extremely critical of this aspect in the past on the basis that a habitat which needs to hold 100 passengers for 1000 days with 1000 days of food, survive reentry, and produce its own propellant on the other end is probably going to be several times that dry mass; But I'm increasingly moving towards backing off this conjunction of requirements, and assuming they will not occur simultaneously.
Even so; Something like the Bigelow BA-2100 is supposed to be heavier than this. On what basis do you propose 75 tons instead of 9 tons or 400 tons?
Yes the dry mass to cargo mass is optimistic as nothing so far has had a greater then 1:1 ratio. The cargo mass though is just a little under half ~46% the mars atmospheric entry mass though as propellants would be ~18% and the lander dry mass would be 36% of the total.
Of the requirements you site I see almost none of these requirements being simultaneous.
I see a transit habitat being used for the 100 count passengers and their life-support needs with the landing vehicle holding them for only a few days at a time in something like air-plane densities. I see re-entry being from much lower orbital speeds at Earth and Mars rather then direct entry. The propellant production will be done with systems that take up most or all of a single landers cargo capacity, the system will be deployed to the surface and remain their, none of it will be integrated to the vehicle but the vehicle that landed it will remain connected and drink-up the first batch of propellant and use that to slowly return to Earth completely empty of cargo. Crew return will be in two legs with the lander just returning to mars orbit and then using SEP to return to Earth orbit.
I see a GLOW of around 4,600 mt, 31 Raptor engines on the booster and a MCT/BFS of only 75 mt dry. A vastly more achievable size.
Which assumes 75mt dry mass for a reusable 100mt Mars lander is realistic...
It's more realistic then most other proposals which presume the lander has >1000 mt of propellant tanks at the same mass. I targeting just 300 mt for tank capacity resulting in a much smaller vehicle and a dry mass fraction of 20%.
Can you find any previous estimates of the mass ratio for dry mass to cargo mass for a (presumably non-reusable) Mars lander, or alternately if your architecture lands back at Earth, for a reusable Earth lander?
I see a GLOW of around 4,600 mt, 31 Raptor engines on the booster and a MCT/BFS of only 75 mt dry. A vastly more achievable size.
Which assumes 75mt dry mass for a reusable 100mt Mars lander is realistic...
I really expect it to be less - F9 has very good mass ratios IIRC, and MCT will be more like a stage (with cargo bay) than a capsule. (I would expect the passenger accommodations to count against the cargo-version 100mt payload.)
It is just for launch and would be directly under the engines with a total area of ~100 m^2, even thick 6 gauge steel wire at a 1 inch spacing would be only 1,220 kg, add in stakes and simple rams that push the wire onto the ground and pin it and your still looking at a tiny fraction of the cargo capacity which is where it is accounted for.
So not only do you have to worry about rocks, you would also have the possibility of long stakes flying up at your spacecraft. Way to many variables in ground composition and installation among other issues to be a viable option. Permanent launch pads will have to be built and by default will become landing pads.
I see a GLOW of around 4,600 mt, 31 Raptor engines on the booster and a MCT/BFS of only 75 mt dry. A vastly more achievable size.
Which assumes 75mt dry mass for a reusable 100mt Mars lander is realistic...
It's more realistic then most other proposals which presume the lander has >1000 mt of propellant tanks at the same mass. I targeting just 300 mt for tank capacity resulting in a much smaller vehicle and a dry mass fraction of 20%.
Can you find any previous estimates of the mass ratio for dry mass to cargo mass for a (presumably non-reusable) Mars lander, or alternately if your architecture lands back at Earth, for a reusable Earth lander?
There are many previous estimates. Only with HIAD you get to 0.75 or better, from what I've seen. For a non-reusable lander.
I see a GLOW of around 4,600 mt, 31 Raptor engines on the booster and a MCT/BFS of only 75 mt dry. A vastly more achievable size.
Which assumes 75mt dry mass for a reusable 100mt Mars lander is realistic...
It's more realistic then most other proposals which presume the lander has >1000 mt of propellant tanks at the same mass. I targeting just 300 mt for tank capacity resulting in a much smaller vehicle and a dry mass fraction of 20%.
Can you find any previous estimates of the mass ratio for dry mass to cargo mass for a (presumably non-reusable) Mars lander, or alternately if your architecture lands back at Earth, for a reusable Earth lander?
There are many previous estimates. Only with HIAD you get to 0.75 or better, from what I've seen. For a non-reusable lander.
I'm looking at a kind of mechanical flap at the base of the vehicle and telescoping or umbrella-like system that performs as a HIAD while being fully retractable and reusable to achieve most deceleration down to sub-sonic followed by brief landing burn.
You just seem to be fishing for negatives at this point, stakes driven into the ground are not going to come out in the time scale of a 2g launch. Ground composition could be an issue and this may mean that their will be a lower maximum boulder size at an acceptable landing site then otherwise.
Not sure if this video has been posted yet..but a schematic of a 100MT lander is illustrated at 46 min mark.
Interesting vid but warning lots of physics!! :)
Cheers
Jb
https://youtu.be/GQueObsIRfI
Like TVG's Michelle-B?
http://www.thespacereview.com/article/409/1
What would be wicked pissah would be a well conceived poll where we could select our BFR/MCT parameters before Elon makes his BFR/MCT announcement* and see who is at least in the ballpark.
* OK, I'm optimistically assuming we'll still be alive at that time.
More difficult: but I would like to see an enumeration of all the basic configurations that are popular.. though I accept the most straightforward design is probably correct. All those wacky ideas just give me a "magnificent men in their flying machines" moment.What would be wicked pissah would be a well conceived poll where we could select our BFR/MCT parameters before Elon makes his BFR/MCT announcement* and see who is at least in the ballpark.
* OK, I'm optimistically assuming we'll still be alive at that time.
I have been thinking about a poll just for number of engines on the first stage. I would choose 36.
Enjoy, Matthew
You just seem to be fishing for negatives at this point, stakes driven into the ground are not going to come out in the time scale of a 2g launch. Ground composition could be an issue and this may mean that their will be a lower maximum boulder size at an acceptable landing site then otherwise.
A less than optimal launch, loose soil, or a manufacturing defect in your unproven wire mesh could lead to a very bad day. You still haven't explained exactly how this pad gets installed. Are you going to depend on a machine that can break down or crew members that might make a mistake? How to you handle quality control when your trying build a new pad for every MCT launch?
As long as your permanent mesh landing/launching pads can hold up to the weight of the MCT and can work for multiple landing and launches, I'm perfectly fine with that.
Not sure if this video has been posted yet..but a schematic of a 100MT lander is illustrated at 46 min mark.
Interesting vid but warning lots of physics!! :)
Cheers
Jb
https://youtu.be/GQueObsIRfI
You'd be lying flat on your back
Though would 'anti-pedalling' to slow down undo all the spin you imparted?
You'd be lying flat on your back
Remember that laying flat is used as a crude proxy for weightlessness in research on Earth, precisely because it causes many of the issues experienced in space even though you are under 1g. Hence, your bike method will achieving no more than peddling on a stationary bike in weightlessness.
Better to simply have a ring-track around the widest inside diameter of the hab section. And use centripetal force to hold you down as you jog/run. A la Skylab & 2001.Though would 'anti-pedalling' to slow down undo all the spin you imparted?
The total energy of the ship and rider (or runner in my case) are conserved, unless mass is thrown overboard. So there'll only be induced spin during the exercise itself; but zero net spin, regardless of which direction(s) you ride.
If you don't want the ship to spin at all, you would probably use a flywheel large enough to store the momentum during the exercise sessions, (reversed when they stop.)
I like the idea of a bike track that the rider lays relatively "flat" to the outer circumference. Quick calculations for a 40 ft inner diameter track says a rider needs to be going 25 feet per second (12 rpm) inside the track for 1g, and 17 feet per second (8.4 rpm) for .5g. Those velocities (11.5 to 17 mph) are easily achieved by a bicycle, but not too easy for a person to sustainably run and get any benefit of daily "g" doses.
I'd minimize flywheel demands by having two tracks that different riders can go opposite directions to cancel out the angular momentum issues. A computer could direct each occupant to step up the pace, or back off on the effort to keep things in balance. Make it a game and take the boredom of out of going in circles. Concurrent riders can match each other in weight. Heck, there's even no reason you couldn't couple the forces produced by the rider to a counterweight rotating in the opposite direction that would cancel his own momentum. It would just add resistance, which is the whole point of exercise.
For artificial gravity, I would simply go with two MCTs linked by a long cable, spinning about each other. The longer the cable, the better.
As to the bike idea - if you were going to use flywheels for rotational control/pitch control anyway on the trip to Mars (no idea how it scales), then just connect the bike to that. Saves you an extra flywheel.
Position relative to CoG would influence tumble.
Why this design over something with a smaller overall diameter and spinning faster with a larger tube diameter, why so few spokes? Why not just go with the more common proposal of docking at the hub? It just looks like a pretty design with no engineering offered as to why that design is an improvement over countless others.Coriolis Nausea (http://en.wikipedia.org/wiki/Artificial_gravity#Rotation) afflicts very small diameter designs. Reasonable rotation rates are up to 2-4rpm; Above 7rpm there seem to be intractable issues in the short term.
At 480m radius, you can get 1G at 1.365rpm - a guaranteed comfortable experience apparently
At 120m radius, you can get 1G at 2.73rpm - Some side effects, but everyone can acclimate to the condition eventually
At 60m radius, you can get 1G at 5.46rpm - Extremely uncomfortable, some people can eventually acclimate, some can't
For artificial gravity, I would simply go with two MCTs linked by a long cable, spinning about each other. The longer the cable, the better.
Not attempting to be rude, but wasn't this idea suggested a long time ago and debunked? It adds numerous mission complications (and contingency risks) and the absence of zero-G increases volume limitations to a spacecraft that is already volume limited.
As to the bike idea - if you were going to use flywheels for rotational control/pitch control anyway on the trip to Mars (no idea how it scales), then just connect the bike to that. Saves you an extra flywheel.
and the absence of zero-G increases volume limitations to a spacecraft that is already volume limited.
I'm pretty sure we've gone through this discussion before. Attempts to simulate gravity on a Mars-bound spacecraft would take up mass and space that could be put to better uses. Adjusting to Mars' lower gravity will not take long. The crew does not need to rush out and do handsprings. If they need to move around right after landing for an emergency, then they will do just as Soyuz passengers returning from the ISS have done after an off-target landing: they will cope.
I like the idea of a bike track that the rider lays relatively "flat" to the outer circumference. Quick calculations [...] for 1gYou'd be lying flat on your backRemember that laying flat is used as a crude proxy for weightlessness in research on Earth, precisely because it causes many of the issues experienced in space even though you are under 1g. Hence, your bike method will achieving no more than peddling on a stationary bike in weightlessness.
Not sure if this video has been posted yet..but a schematic of a 100MT lander is illustrated at 46 min mark.
Interesting vid but warning lots of physics!! :)
Cheers
Jb
https://youtu.be/GQueObsIRfI
I agree that the only viable mars emergency system is to have engine-out capability ideally from the moment of lift-off. Likewise you need lots of thrust to even think about separating from a booster even if it is not a fast separation, both these requirements would lead us to a MCT which has substantially higher thrust to weight ratio then a rocket would typically have.
My own concept is for MCT to mass just 200 mt at launch due to being nearly empty of propellants, 4 Raptor engines give it an acceleration of 4.7 g's. In the event of separation from the booster (which might need assistance from pressure-fed or solids to account for spin-up time of Raptor) a RTLS landing might be possible in the early phases of launch but a down range water landing is most likely. I agree that this would likely result in a scrapping of the vehicle (just as we scrap an airplane that has had a water landing) but it will be well worth it if it saves lives. I see the vehicle simply doing a landing burn to soft land in water (as early F9 tests did) and floating due to it's tanks, passengers will wait for rescue inside the vehicle rather then jumping into life-rafts.
On mars surface with 25 mt cargo and with full propellant tanks GLOW would be 400 mt and the vehicle would have 2.3 g's acceleration which is reduced by mars gravity to just under 2 g's upward at launch, if an engine is lost a diagonal engine must also be shut down to maintain balance and the upward acceleration is only .75 g upward at take off, still very fast by conventional launch standards and more then sufficient to reach orbit with minimal gravity losses.
7 engines at Mars surface just about fit into a 15m diameter. With that many, engine failure where you (hypothetically) have to shut off the opposite thruster is at most a 28% drop, rather than a 50% drop. I like the notion of abort to water during Earth ascent.
Launching with MCT nearly empty would be required to achieve high-G abort, but this would mean the BFR would nearly need to reach orbit. That eliminates or makes very expensive the possibility of a 1-stage BFR.
I think we're probably going to see crew & bulk cargo launch separately from the MCT structure.
My own concept is for MCT to mass just 200 mt at launch due to being nearly empty of propellants, 4 Raptor engines give it an acceleration of 4.7 g's. In the event of separation from the booster (which might need assistance from pressure-fed or solids to account for spin-up time of Raptor) a RTLS landing might be possible in the early phases of launch but a down range water landing is most likely. I agree that this would likely result in a scrapping of the vehicle (just as we scrap an airplane that has had a water landing) but it will be well worth it if it saves lives. I see the vehicle simply doing a landing burn to soft land in water (as early F9 tests did) and floating due to it's tanks, passengers will wait for rescue inside the vehicle rather then jumping into life-rafts.
On mars surface with 25 mt cargo and with full propellant tanks GLOW would be 400 mt and the vehicle would have 2.3 g's acceleration which is reduced by mars gravity to just under 2 g's upward at launch, if an engine is lost a diagonal engine must also be shut down to maintain balance and the upward acceleration is only .75 g upward at take off, still very fast by conventional launch standards and more then sufficient to reach orbit with minimal gravity losses.
So the MCT is the 3rd stage or am I mis-understanding?
So the MCT is the 3rd stage or am I mis-understanding?
Basically yes though the first 2 stages are a full orbital launch system capable of putting any payload into LEO. Despite other folks desire to interpret Musk statements as unconditional support for their positions it is nothing of the sort, Musk simply said the mars spacecraft needs a booster on Earth, he did not specify the stages in said booster.
The impracticality of a 1 stage booster that is dependent on being paired with a fully interplanetary spacecraft of gigantic size to do any launch at all is so high that the idea should be discounted until explicitly confirmed by Musk, and even then I would be highly doubtful such an architecture would successfully get out of development.
So the MCT is the 3rd stage or am I mis-understanding?
Basically yes though the first 2 stages are a full orbital launch system capable of putting any payload into LEO. Despite other folks desire to interpret Musk statements as unconditional support for their positions it is nothing of the sort, Musk simply said the mars spacecraft needs a booster on Earth, he did not specify the stages in said booster.
The impracticality of a 1 stage booster that is dependent on being paired with a fully interplanetary spacecraft of gigantic size to do any launch at all is so high that the idea should be discounted until explicitly confirmed by Musk, and even then I would be highly doubtful such an architecture would successfully get out of development.
I think you are making a straw man there. There could be one booster paired with different reusable upper stages. An general purpose one for most LEO/GTO payloads (developed first), and a more specialized Mars variant. That's how I see it.
Then you are deliberately misreading people and interjecting your own ideas about how it must be done.
MCT may be very stripped down when in cargo mode, essentially a reusable second stage with a long-loiter package and refueling capability. Don't know why you'd insist on yet another stage.
I am at a loss here. I cannot understand where the idea of a second stage comes from.
The abilities of a second stage are a subset of what MCT needs to do going from LEO to Mars surface and, after refuelling, back to earth. So why duplicate that existing capability with a separate second stage?
I'm willing to guess that SX will not make a S2 and put the spaceship on top of it. It will actually use the spaceship to finish the orbit, and then maybe re-fuel it there to get to mars (or an even higher staging point).
Remember, the spaceship we are talking about should be capable of some serious dv if it can reach mars from LEO/HEO and land (or launch from mars and return). You are not going to need another stage on the rocket to get it to LEO, unless in-orbit refueling is out of the question. A re-usable S1 is enough, dilithium crystals would be overkill at this point.. ;)
Moreover, a second stage that reaches orbit is either very expensive (expendable), or very expensive and complex (re-usable). If BFR stays as an exclusive MCT booster (as has been hinted by SpaceX in the past), then its more safe to assume that there won't be a second stage.
In any case, we will have to see how this unfolds. We may get some information by Elon on the architecture by Q1 2016.
Misreading? I'll provide quotes if you like, but it is indisputable that many people have adamantly rejected the idea of a 2 stage BFR and insist on an exclusive MCT as the only 2nd stage.
Geez, not this again. Read what I wrote closer. (Hint: I did not state that you misread *all*) Some have advocated that, but I certainly have not.
I have always advocated a BFR upper stage with a cargo bay, so good luck putting an MCT in or on top of that. An MCT would be an evolved version of such a stage. There is no reason to launch a Mars capable vehicle for most BFR missions, which will be LEO, GTO, and propellant deliveries.
But why do I bother. Will it be Impaler vs the world for another few weeks here? If so, let me know when the dust has settled.
A human rated MCT needs a launch abort system.
The LAS will probably eject the crew compartment only, not the big engines and fuel tanks. So
A mct\bfs will be built of two seperate parts - the propultion module and the utility module.
The propultion module is practically the BFR's 2nd stage and the utility module will be mission specific.
That's how I see it. Does it makes sence?
Maybe it's just me but I read this as Elon saying that the BFR/MCT (now called BFS) is a 2 part entity which I interpret as that stage 2 IS the BFS. A booster rocket (stage one with lots of engines) and a spaceship (stage 2).
Maybe it's just me but I read this as Elon saying that the BFR/MCT (now called BFS) is a 2 part entity which I interpret as that stage 2 IS the BFS. A booster rocket (stage one with lots of engines) and a spaceship (stage 2).Well, still,
"Well, there's two parts of it—there's a booster rocket and there's a spaceship. So the booster rocket's just to get it out of Earth's gravity because Earth has quite a deep gravity well and thick atmosphere, but the spaceship can go from Mars to Earth without any booster, because Mars's gravity is weaker and the atmosphere's thinner, so it's got enough capability to get all the way back here by itself. It needs a helping hand out of Earth's gravity well. So, technically, it would be the BFR and the BFS." As in "Big frakking Spaceship."
http://www.gq.com/story/elon-musk-mars-spacex-tesla-interview?utm_source=10370
This also pretty much says that the spaceship part needs to have the delta V capability for Mars surface to Earth (debatable Earth surface or LEO, maybe either). That means more than small couple hundred mT propellant tanks.
Whether this is the best approach or the final architecture remains to be seen but it would be surprising to see a radical departure from this in the short term assuming there's a spring 2016 BFR/MCT information release.
It does not say whether the BFS is monolitic or modular.
It doesn't mention LEO refuel before mars injection and we know that from previos qoutes.
It does not say 'direct mars surface to earth' only 'from Mars to Earth without any booster' , so refuel in mars orbit is also an option.
Maybe it's just me but I read this as Elon saying that the BFR/MCT (now called BFS) is a 2 part entity which I interpret as that stage 2 IS the BFS. A booster rocket (stage one with lots of engines) and a spaceship (stage 2).
"Well, there's two parts of it—there's a booster rocket and there's a spaceship. but the spaceship can go from Mars to Earth without any booster, because Mars's gravity is weaker and the atmosphere's thinner, so it's got enough capability to get all the way back here by itself. It needs a helping hand out of Earth's gravity well. So, technically, it would be the BFR and the BFS." As in "Big frakking Spaceship."
Point is that the only thing definitively ruled out by Musk statement as a single-stage-to-mars vehicle. Your interpreting every statement by Musk as total confirmation of your speculation but it's far from that.
Point is that the only thing definitively ruled out by Musk statement as a single-stage-to-mars vehicle. Your interpreting every statement by Musk as total confirmation of your speculation but it's far from that.
No, it is the other way around. I base my speculations on statements by SpaceX, you don't. Your concepts may be completely valid, I don't deny that. They are in conflict with SpaceX however.
Point is that the only thing definitively ruled out by Musk statement as a single-stage-to-mars vehicle. Your interpreting every statement by Musk as total confirmation of your speculation but it's far from that.
No, it is the other way around. I base my speculations on statements by SpaceX, you don't. Your concepts may be completely valid, I don't deny that. They are in conflict with SpaceX however.
Sorry guckyfan, ISTM that you take even the slightest off-the-cuff statements and hyperbolic tweets that Musk makes as always being absolute literal truth, with no margin for error. IMO you take some things far too literally and as absolutisms.
How does the MCT escape module land on Mars w/o killing everyone as it slams into the regolith? Chutes ain't gonna git 'er done there.
Just make your upper stage like an airliner, either it gets there or you have no more worries at tax time, ever.
How does the MCT escape module land on Mars w/o killing everyone as it slams into the regolith? Chutes ain't gonna git 'er done there.......they will if you add a retro-rocket to them, like this:
Robotbeat, you have mentioned this 'prepared surface' in several posts. How this would be accomplished?
and you land on a prepared surface.
NASA has some ideas so maybe SpaceX will get some advice from from them.
http://www.nasa.gov/content/landing-pads-being-designed-for-extraterrestrial-missions
Exactly, Dr. Phil is the guy I was thinking of.
Here's some quotes from the article:
"Robotic landers would go to a location on Mars and excavate a site, clearing rocks, leveling and grading an area and then stabilizing the regolith to withstand impact forces of the rocket plume," Mueller said. "Another option is to excavate down to bedrock to give a firm foundation. Fabric or other geo-textile material could also be used to stabilize the soil and ensure there is a good landing site."
...
""We've tested several types of materials and it seems that basalt regolith mixed with polymer binders hold up well," Metzger said."
Sulfer concrete might be a good solution too.
http://www.citylab.com/tech/2016/01/mars-build-house-concrete-sulfur-study/423288/?utm_source=yahoo
I wonder if it could be used for monolithic domes?
https://en.wikipedia.org/wiki/Monolithic_dome
...they will if you add a retro-rocket to them, like this...
How does the MCT escape module land on Mars w/o killing everyone as it slams into the regolith? Chutes ain't gonna git 'er done there.......they will if you add a retro-rocket to them, like this:
https://www.youtube.com/watch?v=4uGfOppQD_g
This is my idea of MCT, focusing on survival on all critical parts of Mars Journey, since there will multiple MCT on the way and they will not probably use 100 people before base is build and they could help each others.That reminds me of the idea which was to fly a F9 reusable 2nd stage together with a dragon 2 as one spaceship.
This is my idea of MCT, focusing on survival on all critical parts of Mars Journey, since there will multiple MCT on the way and they will not probably use 100 people before base is build and they could help each others.That reminds me of the idea which was to fly a F9 reusable 2nd stage together with a dragon 2 as one spaceship.
How does the crew transfer between the escape pod and the hab module through the heat shield?
Is the heat shield necessary for an aerobreak?
Smaller objects have lower terminal velocity and can reach terminal velocity quick enough. And yeah, you need more rocket impulse, but rockets scale well, as long as you're below like 100m/s....they will if you add a retro-rocket to them, like this...
That chute will help some after the heat shield reaches terminal velocity, but in an atmosphere that's 0.0059 X the density of Earth's atmosphere at the surface, you're still going to have a lot more velocity to overcome with retrorockets than that pallet has. Your gravity losses will be less, but you still have the inertia to kill off.
Can it really not do Mars entry engines first? Rocket engines already deal with pretty intense heat...
Can it really not do Mars entry engines first? Rocket engines already deal with pretty intense heat...
If it was feasible to do so, why did we ever invent ablative heatshields? Rocket engines deal with heat either ablatively or with regenerative cooling - heating up the propellant - in which case you need to expend the propellant steadily.
The guidance from the SSTO guys is that for Mars, direct retro burns basically overwhelm any drag benefit of the rest of the spacecraft; You're left with a mass-intensive propulsive descent, as if Mars had no atmosphere.
Randy, Speaking of the SSTO designs of the past, has it occured to anyone that the current design for the Dragon V2 could easily be the same basic design, externally, for the MCT?
If SpaceX were to use a asymetrical aerospike system for each of the quads, the design would be very similar to many of the SSTO designs of the 1960's and 70's. I meantion this idea as what most people know is that exhaust effeciency at various altitudes changes as you get higher, resulting in the need for a radically different exhaust bell than you'd have at sea level. (Part of the reason we stage rockets).
Aerospikes don't suffer this issue and asymetric aerospikes, using mostly Raptor components, should prove most effecient for Landings on both earth and Mars and launches from Mars.
From what I've heard, it seems that the Raptor engines are primarily for the BFR first stage, while the same plumbing could be used for the MCT. This should also give the advantage of less debris being kicked up under the MCT, as the exhaust would be spread and off to the sides, rather than concentrated underneith the craft, and having to make holes in the TPS.
Making holes in the TPS adds mass for the hing and closing systems, (which would also have to be shielded against teh exhaust of engines cutting through the TPS) plus it complicates the landing sequence more than needed.
I have further thoughts on the MCT design that I'll post later.
I'm pointing out that the cooling mechanism on a regeneratively cooled engine, the thing that keeps it from melting during normal operation, is cold propellant flowing through it rapidly on the way to combustion. You can't do that without having the engine turned on. Having the engines directly in the flow poses perhaps the worst aerothermodynamic problem: a parachute-shaped facing surface with sharp frontal edges.Can it really not do Mars entry engines first? Rocket engines already deal with pretty intense heat...
If it was feasible to do so, why did we ever invent ablative heatshields? Rocket engines deal with heat either ablatively or with regenerative cooling - heating up the propellant - in which case you need to expend the propellant steadily.
The guidance from the SSTO guys is that for Mars, direct retro burns basically overwhelm any drag benefit of the rest of the spacecraft; You're left with a mass-intensive propulsive descent, as if Mars had no atmosphere.
That's because the exhaust creates a virtual "aerospike" which reduces the drag on the vehicle, but has nothing to do with the question. He didn't ask about using the engines but if they could stand the reentry temperatures and therefore have an "engines-forward" entry.
Engineers working on the SERV concept originally estimated that they would need TPS doors to cover the engine nozzles during reentry but subsequent testing showed that especially in SERVs case the overall base diameter was sufficient to reduce the heating load so that no TPS doors or special insulation was required. That's Earth though and Mars is a bit different but you should be able to, if the base is wide enough, to enter engines first as long as they are designed properly and integrated into the base of the vehicle. At worst you can probably cold-flow some propellant through the engine system during the period of highest heating, dumping it through the nozzle and overboard.
The benefit is once you're down far enough you only have to add in the other propellant to the engine and ignite it for the powered landing.
Randy
Can it really not do Mars entry engines first? Rocket engines already deal with pretty intense heat...
If it was feasible to do so, why did we ever invent ablative heatshields?
Rocket engines deal with heat either ablatively or with regenerative cooling
- heating up the propellant - in which case you need to expend the propellant steadily.
I'm pointing out that the cooling mechanism on a regeneratively cooled engine, the thing that keeps it from melting during normal operation, is cold propellant flowing through it rapidly on the way to combustion. You can't do that without having the engine turned on. Having the engines directly in the flow poses perhaps the worst aerothermodynamic problem: a parachute-shaped facing surface with sharp frontal edges.Can it really not do Mars entry engines first? Rocket engines already deal with pretty intense heat...
If it was feasible to do so, why did we ever invent ablative heatshields? Rocket engines deal with heat either ablatively or with regenerative cooling - heating up the propellant - in which case you need to expend the propellant steadily.
The guidance from the SSTO guys is that for Mars, direct retro burns basically overwhelm any drag benefit of the rest of the spacecraft; You're left with a mass-intensive propulsive descent, as if Mars had no atmosphere.
That's because the exhaust creates a virtual "aerospike" which reduces the drag on the vehicle, but has nothing to do with the question. He didn't ask about using the engines but if they could stand the reentry temperatures and therefore have an "engines-forward" entry.
No, SpaceX have been sidestepping reentry heating by starting off slow (I'd guess ~2-4km/s at separation instead of ~8km/s from LEO to ~12-14km/s Mars-Earth Transfer) and then spending reentry reducing speed propulsively behind the exhaust plume down to some velocity at which heating is not an issue.I'm pointing out that the cooling mechanism on a regeneratively cooled engine, the thing that keeps it from melting during normal operation, is cold propellant flowing through it rapidly on the way to combustion. You can't do that without having the engine turned on. Having the engines directly in the flow poses perhaps the worst aerothermodynamic problem: a parachute-shaped facing surface with sharp frontal edges.Can it really not do Mars entry engines first? Rocket engines already deal with pretty intense heat...
If it was feasible to do so, why did we ever invent ablative heatshields? Rocket engines deal with heat either ablatively or with regenerative cooling - heating up the propellant - in which case you need to expend the propellant steadily.
The guidance from the SSTO guys is that for Mars, direct retro burns basically overwhelm any drag benefit of the rest of the spacecraft; You're left with a mass-intensive propulsive descent, as if Mars had no atmosphere.
That's because the exhaust creates a virtual "aerospike" which reduces the drag on the vehicle, but has nothing to do with the question. He didn't ask about using the engines but if they could stand the reentry temperatures and therefore have an "engines-forward" entry.
I know it's probably been mentioned many, many times now but it's something which SpaceX have already accomplished, several times now?
https://www.nasa.gov/press/2014/october/new-commercial-rocket-descent-data-may-help-nasa-with-future-mars-landings/
The Space Shuttle also had hatches in it's heat shield for landing gear, I believe. Since the Space Shuttle has flown 133 successful times, I think heat shields like this are very well proven.This is my idea of MCT, focusing on survival on all critical parts of Mars Journey, since there will multiple MCT on the way and they will not probably use 100 people before base is build and they could help each others.That reminds me of the idea which was to fly a F9 reusable 2nd stage together with a dragon 2 as one spaceship.
How does the crew transfer between the escape pod and the hab module through the heat shield?
Is the heat shield necessary for an aerobreak?
Gemini-B would have had a hatch in the heat shield.
I'm a bit confused guys...
When did SpaceX say that the MCT was going to be two sections, a main ship and an escape module?
If you were to have an abort system on MCT, it'd have to work for Mars ascent (as well as Earth), and probably even terminal landing as well. No, this is not impossible. Hard, but not impossible.
Or just not have an abort system.
On the issue of an abort capsule I agree with guckyfan and Robotbeat, it is impractical and offers very little use at Mars while on Earth a whole vehicle abort is more practical.That's not what I think. I think you can do abort on Earth AND Mars. And "whole vehicle" abort is likely not appropriate.
Shotwell mentioned about BFR a few months ago at the South Summit 2015 (Oct 7-9), in Madrid, " [Falcon Heavy] This one is about 4M pounds of thrust, and the mock... the vehicle that takes us to Mars will be three or four times that size"
https://youtu.be/omBF1P2VhRI?t=10m46s
(original video, mostly Spanish-language conference proceeding, but Shotwell's voice still appears beneath a title graphic for the first ten minutes, though not her face. The video I linked above seems to have been created a while after this one was promoted, and does a proper cut to her presentation alone)
I also vaguely remember her mentioning offhand that they were developing a 180-210t to LEO superheavy launcher. I've been trying to find the interview, but can't turn anything up.
15m wide is pretty reasonable, but the main utility of that is to prevent it from being anywhere near 300m tall. I don't expect it to be anywhere near the fineness ratio of F9. A cylindrical rocket doesn't have infinite room to scale up; The base grows as n^2 while mass grows as n^3, and eventually the base's space to fit rocket motors limits further growth. Additionally, we have raised issues with eg the launch facilities here.Shotwell mentioned about BFR a few months ago at the South Summit 2015 (Oct 7-9), in Madrid, " [Falcon Heavy] This one is about 4M pounds of thrust, and the mock... the vehicle that takes us to Mars will be three or four times that size"
https://youtu.be/omBF1P2VhRI?t=10m46s
(original video, mostly Spanish-language conference proceeding, but Shotwell's voice still appears beneath a title graphic for the first ten minutes, though not her face. The video I linked above seems to have been created a while after this one was promoted, and does a proper cut to her presentation alone)
I also vaguely remember her mentioning offhand that they were developing a 180-210t to LEO superheavy launcher. I've been trying to find the interview, but can't turn anything up.
Above Quote from the Update thread
3 to 4 times FH thrust fits well with the 12.7-14.7 million pounds thrust BFR models I've built that conform to prior SX statements.
It contradicts the ludicrous 15m wide 300+m tall !!! behemoths posted on Reddit which would require over 400 Raptors to lift off if the posters had simply calculated the propellant volume and subsequent mass.
15m wide is pretty reasonable, but the main utility of that is to prevent it from being anywhere near 300m tall. I don't expect it to be anywhere near the fineness ratio of F9. A cylindrical rocket doesn't have infinite room to scale up; The base grows as n^2 while mass grows as n^3, and eventually the base's space to fit rocket motors limits further growth. Additionally, we have raised issues with eg the launch facilities here.
Above Quote from the Update thread
3 to 4 times FH thrust fits well with the 12.7-14.7 million pounds thrust BFR models I've built that conform to prior SX statements.
It contradicts the ludicrous 15m wide 300+m tall !!! behemoths posted on Reddit which would require over 400 Raptors to lift off if the posters had simply calculated the propellant volume and subsequent mass.
I suspect you can't build it much smaller than 12m and still have realistic height.
If you try to build it larger than a number somewhere in the 15m-18m range, then you run into issues where bridges need to be rebuilt to get the parts in place, and propellant tanks lose their cylindrical character.
With ~30 engines you probably don't need gimbaling, or at least not on most of the engines. Some variable thrust and a few engines along the edge that gimbal on just one axis should be sufficient and is the Russian way of doing it.
With ~30 engines you probably don't need gimbaling, or at least not on most of the engines. Some variable thrust and a few engines along the edge that gimbal on just one axis should be sufficient and is the Russian way of doing it.
That worked sooo great for the N1. :) I know that was failure of testing, but using variable thrust for control is a truly *terrible* idea. That's when you need more power, not less of it.
A few central engines may be fixed, but most should fully gimbal for general control authority and engine out capability. Having all the engines be identical (including their gimbal setups) also simplifies mass production and testing procedures, as SpaceX illustrate whenever their F9 rocket launches.
(it seems to be a VERY common misconception that some F9 engines are fixed, they are NOT)
With that said, I don't have any good reason to think non-gimballing engines are going to be a thing on BFR. With MCT, on the other hand, there may be value in fixing the combustion chambers as well as the massive engine bells rigidly to the spacecraft; Those gimbals are a source of potential failure with a few years of vacuum between uses, and whether the odds of that are greater than the odds of some other steering system failing, is a bit more of an open question than with BFR.
~30 engines (+/-50%) sounds about right to me. Around the same as Falcon Heavy, but in a single core (which makes engine-out management a little simpler).
15m, too.
...by the way, we can know for a fact that if it's only 15m in diameter, it is going to be limited to around, say, 50 engines of about 500klbf each if they are squished together as close as they can go and have an exit pressure of around near sea level and a nice high chamber pressure of about 2000-3000psi. You cannot physically fit in more engines. For a lift-off T/W of 1.1 or more, that puts a hard upper bound of around 10,000 tons. So if 15m is the diameter (it probably is), then the lift-off mass cannot be more than 10,000 tons. And that leaves ZERO room for gimballing.
A more sane packing arrangement would leave them at around 6000 tons or less.
Re the gimbal / variable-thrust steering discussion for the BFR (booster, not MCT). Given ~30 Raptors what would be the approximate weight saving of variable-thrust?
If you were to have an abort system on MCT, it'd have to work for Mars ascent (as well as Earth), and probably even terminal landing as well. No, this is not impossible. Hard, but not impossible.
Or just not have an abort system.
I still think the BFR/MCT diameter will not exceed 10m. There are many reasons for this. I also think this version would be more economical in the long run. A 3 core heavy for MCT launch. Say 5.5m-8m for a single core for launching deep space probes, filling a fuel depot and other cash making activities. The single core would be able to launch 80-100 tons. A 3 core heavy version 250 tons. A fully reusable 80 ton launcher might eliminate Falcon Heavy. Say it had 9 engines at 550k lbs thrust or slightly more for a 5 million lb thrust reusable rocket. This rocket can put up a lot of stuff to make money from the Air Force, NASA or others to help pay for MCT. A single core could also be adapted to have two Falcon 9 cores attached for boosting slightly over 100 tons to LEO and could still launch from the cape.
I still think the BFR/MCT diameter will not exceed 10m. There are many reasons for this. I also think this version would be more economical in the long run. A 3 core heavy for MCT launch. Say 5.5m-8m for a single core for launching deep space probes, filling a fuel depot and other cash making activities. The single core would be able to launch 80-100 tons. A 3 core heavy version 250 tons. A fully reusable 80 ton launcher might eliminate Falcon Heavy. Say it had 9 engines at 550k lbs thrust or slightly more for a 5 million lb thrust reusable rocket. This rocket can put up a lot of stuff to make money from the Air Force, NASA or others to help pay for MCT. A single core could also be adapted to have two Falcon 9 cores attached for boosting slightly over 100 tons to LEO and could still launch from the cape.
This size also would greatly expand locations for manufacturing and shipping via cheap barge and launch locations. Otherwise if too wide it would have to be built at a shipyard and transported via ocean to various launch locations. River and intercoastal barge widths in America are 36', or 10m maybe 11m.
This may make MCT cylindrical which might mean horizontal landing. A try at using the vacuum Raptor in a reusable upper stage would be a good test for MCT cylinder design.
I still think the BFR/MCT diameter will not exceed 10m. There are many reasons for this. I also think this version would be more economical in the long run. A 3 core heavy for MCT launch. Say 5.5m-8m for a single core for launching deep space probes, filling a fuel depot and other cash making activities. The single core would be able to launch 80-100 tons. A 3 core heavy version 250 tons.
With ~30 engines you probably don't need gimbaling, or at least not on most of the engines. Some variable thrust and a few engines along the edge that gimbal on just one axis should be sufficient and is the Russian way of doing it.
That worked sooo great for the N1. :) I know that was failure of testing, but using variable thrust for control is a truly *terrible* idea. That's when you need more power, not less of it.
A few central engines may be fixed, but most should fully gimbal for general control authority and engine out capability. Having all the engines be identical (including their gimbal setups) also simplifies mass production and testing procedures, as SpaceX illustrate whenever their F9 rocket launches.
(it seems to be a VERY common misconception that some F9 engines are fixed, they are NOT)
Lars you seem to have mentally dropped half of the proposal in your responses so that you can rail against throttling, I said gimbaled engines on the outer perimeter of the vehicle would be used and a vehicle so equipped would only need to vary thrust if these gimbaled engines are insufficient, just as ANY rocket with multiple engines can and would vary thrust beyond what it's gimbaling provides.I liked the idea of gimbaling but I have wondered about the effect of variations in thrust between different engines. They are still actually at full thrust so if variation is 1%, then if engines are gimballed fully outwards you still have sideways 'noise' in your signal equivalent to 1% full thrust, whereas a smaller downwards pointing engine would equivalently only have variation of 1% of its current thrust, and this variation almost entirely straight down in one dimension instead of sideways.
After a couple synods of crewed missions (assuming the number of BFSes each synod grows), there will be a sizable presence on Mars, and it would soon be cheap to have a "launch on need" BFS available (that would be used for a cargo flight back if not used for rescue during that synod, so doesn't use any resources other than space near the pad).If you were to have an abort system on MCT, it'd have to work for Mars ascent (as well as Earth), and probably even terminal landing as well. No, this is not impossible. Hard, but not impossible.
Or just not have an abort system.
"Work" would be subjective. ejecting and landing on mars with no supplies would leave you just as dead.
With ~30 engines you probably don't need gimbaling, or at least not on most of the engines. Some variable thrust and a few engines along the edge that gimbal on just one axis should be sufficient and is the Russian way of doing it.
That worked sooo great for the N1. :) I know that was failure of testing, but using variable thrust for control is a truly *terrible* idea. That's when you need more power, not less of it.
A few central engines may be fixed, but most should fully gimbal for general control authority and engine out capability. Having all the engines be identical (including their gimbal setups) also simplifies mass production and testing procedures, as SpaceX illustrate whenever their F9 rocket launches.
(it seems to be a VERY common misconception that some F9 engines are fixed, they are NOT)
Lars you seem to have mentally dropped half of the proposal in your responses so that you can rail against throttling, I said gimbaled engines on the outer perimeter of the vehicle would be used and a vehicle so equipped would only need to vary thrust if these gimbaled engines are insufficient, just as ANY rocket with multiple engines can and would vary thrust beyond what it's gimbaling provides.
Your N-1 comparison is completely disingenuous by your own admission, the failure were not the result of insufficient control authority provided by fixed engines, the Russians knew this and continued to use fixed engines for decades because they are mechanically simpler. And when engines do gimbal they generally do so on only one axis outward to provide maximum leverage again simplifying all the mechanics. All engines being provided with 2 axis gimbaling is just an American excess not worth the cost and complexity much like Hydrogen fuel.
A 30 engine booster will have around a dozen engines on the out perimeter that can all gimbal on one axis. Only the center engine would need 2 axis gimbaling for landing. That leaves some ~18 engines that can be fixed and packed closer together for maximum thrust density. This is very similar to the configuration the Soyuz uses 4 fixed nozzles and 4 small single axis vernier nozzles, and as we all know the Soyuz is the most successful Russian vehicle and one SpaceX would be wise to copy.
With that said, I don't have any good reason to think non-gimballing engines are going to be a thing on BFR. With MCT, on the other hand, there may be value in fixing the combustion chambers as well as the massive engine bells rigidly to the spacecraft; Those gimbals are a source of potential failure with a few years of vacuum between uses, and whether the odds of that are greater than the odds of some other steering system failing, is a bit more of an open question than with BFR.
I strongly disagree with that. I consider fixed engines more likely on a BFR booster than a MCT. First, the fewer engines you have, the more critical it is that they all gimbal. Otherwise you really have no engine out capability at all, a *long* way from home. Yes, your gimbal actuators need to be VERY reliable, but I think they need to be there to provider multiple levels of redundancy.
Another possibility would be an unmanned land-on-need BFS (or other vehicle) kept in Mars orbit with reserve supplies; after an abort you could drop it next to the survivors to give them additional supplies, shelter, and/or surface mobility. It could also bail out the main settlement if something bad happened to their storage on the surface."Work" would be subjective. ejecting and landing on mars with no supplies would leave you just as dead.After a couple synods of crewed missions (assuming the number of BFSes each synod grows), there will be a sizable presence on Mars, and it would soon be cheap to have a "launch on need" BFS available (that would be used for a cargo flight back if not used for rescue during that synod, so doesn't use any resources other than space near the pad).
It's been suggested before that a small number of the very first BFS's (formerly MCT) landers on Mars will not return to Earth. It shouldn't be hard to modify one or two into a point-to-point planetary hopper. It's a useful tool to have, and serves as a rescue vehicle.
It's been suggested before that a small number of the very first BFS's (formerly MCT) landers on Mars will not return to Earth. It shouldn't be hard to modify one or two into a point-to-point planetary hopper. It's a useful tool to have, and serves as a rescue vehicle.
If they can do such hops, they can probably return to earth as well.
Depends on the size of the hops Paul was going for, and... mathematically that doesn't seem logical. What's your reasoning, Guckyfan? Whilst I agree with you all BFSs will RTE, the margins to hop and the margins for Earth return are considerably far apart from each other.
It's been suggested before that a small number of the very first BFS's (formerly MCT) landers on Mars will not return to Earth. It shouldn't be hard to modify one or two into a point-to-point planetary hopper. It's a useful tool to have, and serves as a rescue vehicle.
If they can do such hops, they can probably return to earth as well.
Depends on the size of the hops Paul was going for, and... mathematically that doesn't seem logical. What's your reasoning, Guckyfan? Whilst I agree with you all BFSs will RTE, the margins to hop and the margins for Earth return are considerably far apart from each other.
My reasoning. BFS are capable of return to earth when functioning. They need to be functioning to make the hops.
Depends on the size of the hops Paul was going for, and... mathematically that doesn't seem logical. What's your reasoning, Guckyfan? Whilst I agree with you all BFSs will RTE, the margins to hop and the margins for Earth return are considerably far apart from each other.
My reasoning. BFS are capable of return to earth when functioning. They need to be functioning to make the hops.
Being functioning and being supplied by an active ISRU facility is a wholly separate thing. Without propellant it doesn't matter.
Or ISRU methane production is built into the early BFS's, or maybe all of them for redundancy. How large would the device be?
About a Megawatt of power is needed to produce enough propellant for a MCT in a year. 40 tons unless you hook up to base-side infrastructure. If you have nuclear, you need huge radiators (or ground infrastructure) and shielding of some kind. Solar of 1MW requires even larger deployment on the ground. Electrolysis and Sabatier are a little smaller but still substantial.
It simply isn't a good idea to keep it on the MCTs.
Or ISRU methane production is built into the early BFS's, or maybe all of them for redundancy. How large would the device be?
It's been suggested before that a small number of the very first BFS's (formerly MCT) landers on Mars will not return to Earth. It shouldn't be hard to modify one or two into a point-to-point planetary hopper. It's a useful tool to have, and serves as a rescue vehicle.
A launch-on-need hopper can be a one-way trip. And if it has 7km/s of delta-V, it can go anywhere on Mars.
A launch-on-need hopper can be a one-way trip. And if it has 7km/s of delta-V, it can go anywhere on Mars.
Huh? what use is it if it can't return to where it started from?
A launch-on-need hopper can be a one-way trip. And if it has 7km/s of delta-V, it can go anywhere on Mars.
Huh? what use is it if it can't return to where it started from?
It can carry enough supplies and solar power units to keep the stranded crew alive until a ground retrieve mission can reach them. If need be additional resupply flights.
If it can only get as far as the stranded people it can't make subsequent flights. You are using up hoppers on a one way trip. So whether it is a custom designed hopper or an MCT that was left behind it is an emergency solution as much as a left behind MCT (and eventually several) will be the evac solution for a settlement if something makes it non viable early on.
Or ISRU methane production is built into the early BFS's, or maybe all of them for redundancy. How large would the device be?
Bingo. There will be a lot of cargo MCTs to develop a colony, very few will likely carry ISRU equipment.
...however every kilo of propellant used in that endeavour would represent 3 - 4 kilos of ISRU propellant and at say 200t of propellant per tanker flight, you are adding engine cycles on your reusable MCT's fairly quickly.That's where the incentive for Phobos\Deimos\asteroid mining and atmospheric scooping lay.
If it can only get as far as the stranded people it can't make subsequent flights. You are using up hoppers on a one way trip. So whether it is a custom designed hopper or an MCT that was left behind it is an emergency solution as much as a left behind MCT (and eventually several) will be the evac solution for a settlement if something makes it non viable early on.
Isn't this highly dependent on how far the people are away from base and how the hopper is built? I would assume a custom hopper could weigh quit a bit less than an MCT. Even if it could only make a one-way trip, why couldn't a slow moving wheeled tanker just come along later and refuel it? Sure a orbital depot might make sense for exploration, but not IMO required for a rescue op.
So just go into very low Orbit and deorbit. That'd take less than ~5km/s delta-v.I was using 4.1km/s for surface to LMO
btw, are you relying on Earth figures, or are you recalculating them for Mars? The lower gravity makes a huge difference here.
...BTW, this is all off-topic. I just brought up the idea to counter the false (but oft-repeated) claim that abort would be useless for MCT because there'd be no way to get to them.
care to show your work?So just go into very low Orbit and deorbit. That'd take less than ~5km/s delta-v.I was using 4.1km/s for surface to LMO
btw, are you relying on Earth figures, or are you recalculating them for Mars? The lower gravity makes a huge difference here.
...BTW, this is all off-topic. I just brought up the idea to counter the false (but oft-repeated) claim that abort would be useless for MCT because there'd be no way to get to them.
yes I was calculating for Mars for the point to point distances
The concept of "hoppers" for Mars is not really practical the way many people are throwing it around here. IF you expect the hopper to go point to point on Mars and return to the origin point without refuelling the range of a hopper that has enough ΔV to make it to low Mars orbit is less than 200km. The same hopper could go nearly 1000km if it can refuel at its destination, but that is not useful for exploration because you can't just go anywhere. For serious exploration by rocket powered craft what is needed is a craft that can descend from orbit fully fuelled to any location on Mars and, after landing have enough fuel to make it back to orbit. If that can be achieved then there is a point to using that sort of craft to travel on Mars, otherwise the limitations on such craft is such that wheeled vehicles will be far more practical. Even so to make it work you would need fuel depots in Mars orbit.
A fully fuelled MCT (ΔV = 7.5km/s) on the surface of Mars could make a one way trip to any point on Mars most likely, but a two way trip would be limited to something around 700km.
Note MCT left at Mars for their use at Mars would make excellent support craft for a Mars Orbital Station/Depot and if a few dedicated smaller craft optimized for orbit-surface-orbit on a single load of propellant existed that depot could support exploration of any point on Mars surface, however every kilo of propellant used in that endeavour would represent 3 - 4 kilos of ISRU propellant and at say 200t of propellant per tanker flight, you are adding engine cycles on your reusable MCT's fairly quickly.
Still as pointed out above the real limit early on will be the ISRU capacity.
The BFS is a stage in and of itself. If it's intact enough to land, you'd likely be better off aborting to orbit.
BTW, MCT is the whole system, BFR and BFS together. What you call the MCT is what Musk calls the BFS.
care to show your work?So just go into very low Orbit and deorbit. That'd take less than ~5km/s delta-v.I was using 4.1km/s for surface to LMO
btw, are you relying on Earth figures, or are you recalculating them for Mars? The lower gravity makes a huge difference here.
...BTW, this is all off-topic. I just brought up the idea to counter the false (but oft-repeated) claim that abort would be useless for MCT because there'd be no way to get to them.
yes I was calculating for Mars for the point to point distances
It stands to reason that if a hopper over 700km exceeds orbital delta-V, then an aborted crew would not be more than ~700km away.
That close, and a land-train of pressurized rovers would likely be able to meet them within a couple days of travel (or even 8-10 hours with a slightly prepared path, within EVA+contingency time). If the aborted crew aborted along with a rebreather system, they should be able to hang out in their spacesuits that long.
There would likely be only one orbital track for return from Earth, so you could possibly prepare a route before hand.
No evidence of a depot in LMO. Vastly more likely to launch directly to Earth or possibly visit a depot in high orbit like MSL-2.Obviously there is no evidence of a depot in orbit around Mars yet, but there are no people, hoppers or vehicles returning to Earth. There is evidence that depots in LMO are a practical solution to a number of issues of the logistics of settling and exploring Mars. If you read the various threads on this forum there is as much evidence of depots as of other systems discussed here.
BFS should be able to get to orbit in a jiffy. They would only abort if they couldn't make it to orbit (even with secondary thrusters). It doesn't take 10,000km to get to orbit in a BFS with high thrust.
If you're that close, thrusters could get you to orbit. You just said that more than 700km would take more than orbital velocity, a claim I still have seen no calculations for that include lift.No evidence of a depot in LMO. Vastly more likely to launch directly to Earth or possibly visit a depot in high orbit like MSL-2.Obviously there is no evidence of a depot in orbit around Mars yet, but there are no people, hoppers or vehicles returning to Earth. There is evidence that depots in LMO are a practical solution to a number of issues of the logistics of settling and exploring Mars. If you read the various threads on this forum there is as much evidence of depots as of other systems discussed here.BFS should be able to get to orbit in a jiffy. They would only abort if they couldn't make it to orbit (even with secondary thrusters). It doesn't take 10,000km to get to orbit in a BFS with high thrust.
HUH again, if we are talking about a failure that causes an abort the failure could take place anywhere between launch and when the final orbit is achieved. So if that failure takes place at 95% of orbital velocity just 100km down range from the launch site, momentum would take them halfway around Mars.
If you're that close, thrusters could get you to orbit. You just said that more than 700km would take more than orbital velocity, a claim I still have seen no calculations for.
The only reason the crew would be back on the surface is if the vehicle had a really bad day and the crew had to bail out, perhaps in a pod or ejection seats. Probably a pod, since it'd allow survival from Near Mars orbital velocity.
Then parachute down, with solid retrorockets to keep parachute size reasonable. The whole abort system could be relatively low mass. 2kg for the chute, 5-10kg for the rockets per passenger, plus whatever the suit would weigh and the pod to protect against the relatively modest Mars reentry (much more modest than. Typical Mars probe which comes in hyperbolically).
What delta-V are you assuming is needed for landing?If you're that close, thrusters could get you to orbit. You just said that more than 700km would take more than orbital velocity, a claim I still have seen no calculations for.
It is above! You can also do the math yourself see the link I quoted above along with my calculations:
https://www.rand.org/content/dam/rand/pubs/research_memoranda/2008/RM3752.pdf (https://www.rand.org/content/dam/rand/pubs/research_memoranda/2008/RM3752.pdf)
I said that it would take the same ΔV as going to orbit to launch and land 700km down range. You only need half orbital velocity to launch and crash 700km down range.
Equal to take off, particularly on hops that are less than 2km/s launch velocity since they are also really high angleWhat delta-V are you assuming is needed for landing?If you're that close, thrusters could get you to orbit. You just said that more than 700km would take more than orbital velocity, a claim I still have seen no calculations for.
It is above! You can also do the math yourself see the link I quoted above along with my calculations:
https://www.rand.org/content/dam/rand/pubs/research_memoranda/2008/RM3752.pdf (https://www.rand.org/content/dam/rand/pubs/research_memoranda/2008/RM3752.pdf)
I said that it would take the same ΔV as going to orbit to launch and land 700km down range. You only need half orbital velocity to launch and crash 700km down range.
Ok Robobeat you modified your post to say calculations with lift, that was not there when I replied. Lift is pretty irrelevant at lower speeds especially since to get efficient ballistic distances from your launch velocity you have to launch at an angle greater than 40 degrees (which means as you approach the surface (and lift can only play a role in the last 10 km or so) you are coming in at the same angle. So lift is not going to increase your range significantly.Boost-skip is one strategy to significantly increase point to point range. Lift is a big component of that.
Poor assumption. Lift changes the optimum angle significantly, can allow you to reach near terminal velocity.Equal to take off, particularly on hops that are less than 2km/s launch velocity since they are also really high angleWhat delta-V are you assuming is needed for landing?If you're that close, thrusters could get you to orbit. You just said that more than 700km would take more than orbital velocity, a claim I still have seen no calculations for.
It is above! You can also do the math yourself see the link I quoted above along with my calculations:
https://www.rand.org/content/dam/rand/pubs/research_memoranda/2008/RM3752.pdf (https://www.rand.org/content/dam/rand/pubs/research_memoranda/2008/RM3752.pdf)
I said that it would take the same ΔV as going to orbit to launch and land 700km down range. You only need half orbital velocity to launch and crash 700km down range.
Are you assuming the crew would have enough supplies
Impaler for the shorter distances aero lift and braking will not make any significant difference and have much less impact than gravity losses. Once you go above 2km/s in launch speed then it is possible that you can either use one of or both lift and drag to somewhat reduce the amount landing ΔV required. But you will still need a significant proportion of your launch ΔV to land even at the antipode.
Nadreck, an abort scenario has to be realistic. If the BFS is able to be used as its own abort vehicle, it needs to be able to land. That means at least some engines are functional. So anything over half orbital velocity and your mission profile would be abort-to-orbit, then land at the base in the normal way once your orbit aligns with the base again.
During the first couple of minutes, you're close enough and slow enough to RTLS. After 5-6 minutes you default to abort-to-orbit. So there's a narrow window where you can't make orbit, but can land, but are too far from the base to RTLS. That's not going to be 10,000km.
(If they crash, they die.)
With a light payload, decent lift and ballistic coefficient, You shouldn't need more than 1km/s to land from orbit. It takes 4.5km/s or less to get to low orbit. 7km/s would be enough delta-V to land anywhere on the planet, maybe except Olympus Mons.
According to my simulations even 500 m/s is sufficient for landing at low altitude
Are you assuming the crew would have enough supplies
By definition, somewhere between 3 and 9 months worth.
And can you give me the results of your sim calculation with the same threshold altitude, an angle of 40 degrees, the shape was a cone with a rounded base with the height and diameter of the cone being 10 meters, and the mass was 60,000kg?According to my simulations even 500 m/s is sufficient for landing at low altitude
can you give me your assumptions for this: altitude where you start calculating drag and lift, angle of flight at that altitude, velocity at that altitude, cross section area, mass
No threshold, I use an exponential atmosphere model. I posted a link above. It contains all assumptions.
I should point out that their are MANY possible reason why we might wish to abort to surface other then loss of engines, anything which endangers the long term life-support capability, or compromises the heat-shield which would make Earth entry in that vehicle dangerous. So I can't agree that any abort to surface is necessarily coming down with no propulsion and must therefore be a bailout.Abort to orbit in most of those situations would be much safer since you're not risking landing in a compromised vehicle. The only reason you'd want to be on the surface rather than in orbit is if you literally can't get to orbit.
I should point out that their are MANY possible reason why we might wish to abort to surface other then loss of engines, anything which endangers the long term life-support capability, or compromises the heat-shield which would make Earth entry in that vehicle dangerous. So I can't agree that any abort to surface is necessarily coming down with no propulsion and must therefore be a bailout.Abort to orbit in most of those situations would be much safer since you're not risking landing in a compromised vehicle. The only reason you'd want to be on the surface rather than in orbit is if you literally can't get to orbit.
No. Aborting to orbit puts the vehicle through the least thermal load. They would be rescued by another vehicle. That was the plan for Shuttle post-Columbia, and it'd likely be the case for BFS at Mars.I should point out that their are MANY possible reason why we might wish to abort to surface other then loss of engines, anything which endangers the long term life-support capability, or compromises the heat-shield which would make Earth entry in that vehicle dangerous. So I can't agree that any abort to surface is necessarily coming down with no propulsion and must therefore be a bailout.Abort to orbit in most of those situations would be much safer since you're not risking landing in a compromised vehicle. The only reason you'd want to be on the surface rather than in orbit is if you literally can't get to orbit.
It depends on how the vehicle is compromised, if were in doubt of the vehicles thermal protection then aborting the launch immediately and landing propulsivly puts the vehicle through the least thermal load, less even then reaching orbit and then landing again.
Hi everyone, I just made a design concept, I'm a product designer, not engineer.
Newest version: http://imgur.com/a/15fO2
Initial concepts and idea development: http://imgur.com/a/EtH8F
Why are all the rocket tall and cylindrical? What are the disadvantages of launching disc like space ships?
Is it possible to build a single stage spaceship that is capable of launching from earth and landing on mars?
Hi everyone, I just made a design concept, I'm a product designer, not engineer.I appreciate your excellent work to create superb renderings.
Newest version: http://imgur.com/a/15fO2
Initial concepts and idea development: http://imgur.com/a/EtH8F
I appreciate your excellent work to create superb renderings.
A rocket engineer would require some changes to the model based on his/her calculations and experience. You should be aware that a team of persons in the L2 section of this forum are also creating renderings based on expert opinion and hints from SpaceX. It would be worth your joining L2 just to review the large amount of information and the history behind their work.
I have aldo proposed some concepts on this forum that have received good reviews and critiques, but I have mostly relied on sketches. I would love to provide some ideas to you for your review and your rendering machine.
Thank you Ionmars, I appreciate your comments and loved to see your concept, I'll talk more about it later.
As I don't have the technical skills, I based this concept on other studies, my view of SpaceX' design philosofy, etc.
I think proportions would be very different, with the proper calculations, wind tunnel testing and right materials, I was tryng more to find a possible logic on the design, but this concept is not more than thinking aloud.
I've seen lots of articles and fan proposals, but every time it appears a new information that should change the basic design, so it's impossible to track everything. For me this is a funny guessing game and will be very interesting to see who had the technical skills and luck to be closer to the truth, when all is revealed.
For my design, specifically, I tried to keep it simple, so it can be feasible on the development, costs and safety fields on a 15 years timeframe by now (more time than SpaceX exists).
IMO, not everything of what was said about the MCT will be achievable at the same time, it looks that the ship would become too heavy, expensive and complex, so I expect to see something brilliant and elegant, or something simplier, which is my approach.
I had seem some of your sketches before (nice to meet you finally) and I think that is the right bet if the MCT is the monolithic lander ship as the information goes. As you designed I also agree that they will rely on previous spacex's designs, evolutionary. I personaly bet on a modular ship instead. But as this is the main concept by this time, I'm thinking how to make a monolithic (almost) and refueling on orbit before going to mars. I assume it will be close to your design. But it'll have a difference, the propulsion module will not land back on earth as I presume you designed (I'm not sure if it's the case by the sketches I've seem by now). It would land here using the abort system, like the dragon. The new BFR, on the second MCT mission would launch the lander + fuel + more cargo then before, to meet the MCT (propulsion + mars landing + crew hab and radiation shielding)already on orbit to refuel and servicing it.
I just saw your cargo unloading solution and is close of what I was thinking (attached rough). I think a car would ride on a leg or the side fuselage instead. Could I see the rough dimensions of your concept?
Yes, BSenna. You and I are thinking alike.
I do not have a specific concept of the MCT design; I just react to new info as I discover it. I have some experience as a engineering project manager but my principle role now is "Idea Man." But as a team player I am ready to change my concepts when a better idea is presented. So if you and I were to work together, you should take the lead on what the MCT design is likely to be.
My arena would be an idea about a structure that the MCT architecture will require: an in-space propellant depot. If interested, it has 3 threads on this forum so far, starting with this:
http://forum.nasaspaceflight.com/index.php?topic=38146.0
...OK, IMO your architecture reflects what most engineers are thinking is reasonable. MCT will be the second stage of a two-stage system with BFR as a muscular stage I. Both stages will have methane and LOX as fuels and Raptor engines. SI will only be used to launch from Earth and will return as a reusable SI, much like F9R.
...
Great, I'll love to colaborate as much I can! I'll take a look. The monolithic version of the previous concept:
http://imgur.com/a/CCtZ5
Great, I'll love to colaborate as much I can! I'll take a look. The monolithic version of the previous concept:
http://imgur.com/a/CCtZ5
What do you guys think of a 'Return Capsule' plan of operations, where the tip of the MCT reenters with humans in Orion or Dragon-grade accomodations, but where the MCT hab + rocket itself never returns to the Earth Surface or to LEO, but stays in high orbit around Earth and receives repairs / refueling / cargo / passengers there?
What do you guys think of a 'Return Capsule' plan of operations, where the tip of the MCT reenters with humans in Orion or Dragon-grade accomodations, but where the MCT hab + rocket itself never returns to the Earth Surface or to LEO, but stays in high orbit around Earth and receives repairs / refueling / cargo / passengers there?
Why not have the "return capsule" just be part of the vehicle that ferries crew/cargo up and where the BFS never comes back down and gets all its servicing in LEO or at LM-1 or 2.
Musk is said to have ordered tooling for a 15 m fuel tank so I think this will be the diameter of main parts of SI and SII. Not so sure whether a wider 19 m capsule will sit on top.
Musk is said to have ordered tooling for a 15 m fuel tank so I think this will be the diameter of main parts of SI and SII. Not so sure whether a wider 19 m capsule will sit on top.
I see no supporting quote for this in the SX MCT info thread...
http://forum.nasaspaceflight.com/index.php?topic=37839.0
Source?
Great, I'll love to colaborate as much I can! I'll take a look. The monolithic version of the previous concept:
http://imgur.com/a/CCtZ5
I like your concept - and your illustration style - but I have two areas of concern:
- Abort capsule - I don't think it is necessary or helpful (lots of arguments over this I realize)
- The raptors are canted outwards and lose efficiency
PS. The L2 section of this forum carries a personal messaging service that makes inter-personal messaging very convenient. (No I get no commission) :)
...OK, IMO your architecture reflects what most engineers are thinking is reasonable. MCT will be the second stage of a two-stage system with BFR as a muscular stage I. Both stages will have methane and LOX as fuels and Raptor engines. SI will only be used to launch from Earth and will return as a reusable SI, much like F9R.
...
Great, I'll love to colaborate as much I can! I'll take a look. The monolithic version of the previous concept:
http://imgur.com/a/CCtZ5
The SII (MCT) will come in three varieties, a passenger vehicle. a cargo carrier and a fuel tanker. But they will have the same outer structure.
There will be differences in how the different versions will be utilized, but your stage 5, relaunch from Mars, seems pretty solid, at least for passenger and cargo versions. There are other observations to be made, but more appropriate for a section in L2.
Some differences: You show separation of a capsule from the main body when landing on Earth but both parts need to be landed for reusability and may likely land together. Also, you reasonably show refurbishment in step 8 but as time goes on Musk wants this to become minimal for rapid relaunch.
Musk is said to have ordered tooling for a 15 m fuel tank so I think this will be the diameter of main parts of SI and SII. Not so sure whether a wider 19 m capsule will sit on top.
I like your sketches.
Now it's your turn.
15 diameter will be great for aerobraking at Mars and back at Earth.
1/How much this diameter increase effect launch performance?
2/My estimate rocket will be only 70m tall(max 100m), how much make it easy to steering MCT during landing on Mars and Earth?
...because we've never ever had the ability to try SSRP, which allows you to land large payloads provided you use enough propellant.15 diameter will be great for aerobraking at Mars and back at Earth.
1/How much this diameter increase effect launch performance?
2/My estimate rocket will be only 70m tall(max 100m), how much make it easy to steering MCT during landing on Mars and Earth?
15m is still extremely small for aerobraking over 100 Mg on Mars. In mass per area, its an order of magnitude worse than anything we've ever sent to Mars.
Maybe a poll forecasting Musk's coming out of the closet date?
Maybe a poll forecasting Musk's coming out of the closet date?
How about a poll with the top ten BFS/BFR designs and mission architectures?
Musk already dropped a date for when he wants to reveal the architecture: September 26-30.
Musk already dropped a date for when he wants to reveal the architecture: September 26-30.Hopefully all the speculation on BFR/MCT system will end then.
Musk already dropped a date for when he wants to reveal the architecture: September 26-30.Hopefully all the speculation on BFR/MCT system will end then.
Maybe a poll forecasting Musk's coming out of the closet date?
How about a poll with the top ten BFS/BFR designs and mission architectures?
Who compiles the 1000 page pdf document detailing these?
Yes, BSenna. You and I are thinking alike.
I do not have a specific concept of the MCT design; I just react to new info as I discover it. I have some experience as a engineering project manager but my principle role now is "Idea Man." But as a team player I am ready to change my concepts when a better idea is presented. So if you and I were to work together, you should take the lead on what the MCT design is likely to be.
My arena would be an idea about a structure that the MCT architecture will require: an in-space propellant depot. If interested, it has 3 threads on this forum so far, starting with this:
http://forum.nasaspaceflight.com/index.php?topic=38146.0
You mean just like all speculation on the Falcon 9, Dragon 2, and Falcon Heavy have stopped?
You mean just like all speculation on the Falcon 9, Dragon 2, and Falcon Heavy have stopped?
People will stop suggesting completely alternative architectures and instead speculate about architecture upgrades. That will cut some of the 'Dastardly and Muttley in Their Flying Machines/Catch that pigeon' proposals.
Excuse my cynicism but even before the USAF award to develop a raptor or raptor variant for FH and F9 upper stage use it had been discussed for a long time here along with potential F9 or FH new reusable US. So I fully expect that whatever Musk announces in September some variants in both overall architecture (including depot/refuelling, upper stage types, dependencies on site preparation at Mars, Mars ISRU and depots, SEP on BFS, BFS return to Earth surface, LEO, LM-1/2, etc) and specific details (BFS engine arrangement, TPS, BFB recovery, BFS ECLSS, BFS cargo arrangemnt etc) will still get debated well outside guidelines of where SpaceX says it will be going.The end of speculation lies just beyond Andromeda Galaxy.
I took a look at the propellant depot project and it looks great, very refined. But unfortunately is too much for my technical limitations, so, I didn't undestood it completely, but I trust on you guys!Here's the way I see the Depot working:
I could start a 3d model, but my habilities are really limited. I have some doubts:
I assume that the MCTs which will trasport the fuel to LEO are of a similar size and mass of the crewed MCT, so, they're not full of fuel when reach the depot, right? If It's the case, would there be any advantage on leaving only a smaler fuel tank on each of these missions and returning to land to fly again on the same role, thus reducing the total fleet of MCT-SIIs) and also reducing the total mass of the complete transit to mars vehicle? Or The rockets on these MCTs are also necessary to throw the complete vehicle to Mars?
Once on Mars, just the crewd MCT would land, refuel and return to meet the rest of the vehicle?
Then, all 7 MCTs would land on Earth for refurbishing?
Here's the way I see the Depot working:
The first half dozen or so trips to Mars will carry cargo to set up habitats, ISRU machinery, and remote-contolled equipment. These preliminary forays will prepare a base for humans who will arrive later. A cargo MCT launching from Earth will be heavily loaded; it will use up essentially all its fuel just to get to the depot in orbit. It will require a refill before proceeding any further.
We will have a tanker variant of MCT, which has the same basic outer shell as a cargo or passenger version, but will be hollowed out as a big fuel carrier. The volume that would otherwise be cargo volume will be devoted to big fuel tanks. This could be just an expansion of the the main tanks that extend into the the cargo space. Its principle function will be to haul fuel from Earth to the depot, probably about 250 tonnes of CH4 and LOX in each trip. . Three or four loads of fuel must be transferred to the cargo-MCT before it can proceed to Mars.
We could just make a series of trips to the waiting cargo ship and sequentially fill it up. But having a depot allows us to fill up multiple cargo ships and launch as a fleet to Mars. This is the stated intent of SpaceX.
So tanker #1 arrives at the depot and docks at a berth. Tanker #2 arrives at the depot and transfers its fuel load to tanker #1, reserving just enough fuel to return to Earth via propulsive landing. These tankers are highly reusable and require little maintenance between flights, like an airplane. Tanker #2 refills at the launch site and launches again to the depot. It again transfers its fuel load to tanker #1 and returns to the launch site. Now tanker # 1 has enough fuel in its tanks to service one cargo MCT. And because we have six berths at the depot, we can perform this procedure three times simultaneously using six tanker MCTs.
Now three cargo ships arrive at the depot and dock beside three refilled tankers. The tankers transfer their fuel loads to the cargo ships and return to Earth.
We have three cargo carriers ready to trek to Mars with full bellies. :)
If you were to do a 3D model of the Depot it would just be a 2D (x,y) skeletal frame that is repeated into the third (z) dimension. Below is a sketch of the forward frame for one berth. The small circle is a "pad" for connecting to the nose of the MCT and the large circle is an imaginary volume representing one berth where one MCT is docked. The frames are connected by beams of a certain length in the z dimension. Then this 3D image is duplicated side-by-side to the right to make 5 additional berths until they join on the left side of the first berth. Voila!Blah,, blah... blah
...
...
Undesrtood!
We will have a tanker variant of MCT, which has the same basic outer shell as a cargo or passenger version, but will be hollowed out as a big fuel carrier. The volume that would otherwise be cargo volume will be devoted to big fuel tanks. This could be just an expansion of the the main tanks that extend into the the cargo space. Its principle function will be to haul fuel from Earth to the depot, probably about 250 tonnes of CH4 and LOX in each trip. . Three or four loads of fuel must be transferred to the cargo-MCT before it can proceed to Mars.
TMI burn window. Need time to get the propellants and MCT in orbit before the window opens. Depot would be designed to store the propellants over a long period of time and could be used by other in space craft.
We will have a tanker variant of MCT, which has the same basic outer shell as a cargo or passenger version, but will be hollowed out as a big fuel carrier. The volume that would otherwise be cargo volume will be devoted to big fuel tanks. This could be just an expansion of the the main tanks that extend into the the cargo space. Its principle function will be to haul fuel from Earth to the depot, probably about 250 tonnes of CH4 and LOX in each trip. . Three or four loads of fuel must be transferred to the cargo-MCT before it can proceed to Mars.
Cargo density is far lower than fuel density, so either the tanker is smaller or it has a lot of unused cargo volume.
I don't understand why you would need a depot though. Why not just have the tanker dock with the MCT and transfer propellent?
If you were to do a 3D model of the Depot it would just be a 2D (x,y) skeletal frame that is repeated into the third (z) dimension. Below is a sketch of the forward frame for one berth. The small circle is a "pad" for connecting to the nose of the MCT and the large circle is an imaginary volume representing one berth where one MCT is docked. The frames are connected by beams of a certain length in the z dimension. Then this 3D image is duplicated side-by-side to the right to make 5 additional berths until they join on the left side of the first berth. Voila!Blah,, blah... blah
...
...
Undesrtood!
I think after several cargo mission, first with FH and then MCT. When MCT successfully land on Mars. First crew will arrive without specified return day. There task will be prepare infrastructure and ask for additional resources in preparation of infrastructure for return trip. Setup ice collection and build and maintain devices to create fuel,build tank for fuel. At the moment finishing task and prepare rocket for return trip some of the crew will start their trip back to Earth. I think when first crew will arrive there will be not enough fuel to bring them back, they have to build their "return ticket".
Nice skill, BSenna.If you were to do a 3D model of the Depot it would just be a 2D (x,y) skeletal frame that is repeated into the third (z) dimension. Below is a sketch of the forward frame for one berth. The small circle is a "pad" for connecting to the nose of the MCT and the large circle is an imaginary volume representing one berth where one MCT is docked. The frames are connected by beams of a certain length in the z dimension. Then this 3D image is duplicated side-by-side to the right to make 5 additional berths until they join on the left side of the first berth. Voila!Blah,, blah... blah
...
...
Undesrtood!
Is this?
http://imgur.com/a/KJRna
An international depot presupposes everyone using the same fuel, which doesn't look like the case at all. Assuming ULA gets ACES tankers off the ground they look to supply LH2. SpaceX is of course going methane. The Russian Fenix looks to be a methane "Zenit" (Sputnik News) and clustered for a heavy, but do we know what they'll use once in space if they even use tankers? China??Good point. Can't please everybody, but BE-4 methane engine will be used for something. What?
Crazy idea number 854:No, it isn't feasible. But I do like crazy ideas.
Could a fuel depot like this dip a "very long" stiff hose into the atmosphere, pump air up, and make fuel by processing CO2 and/or capturing methane?
Granted it might take a really long time, but is it even technically possible? I assume the power supply to run this would have to be solar. since if it is anything else, you use more fuel than you make most likely.
The first stage of Vulcan and Blues own launcher, and their uppers use LH2. I haven't seen anything about a BE-4 Vac.Ok, so not compatible with a MCT depot.
No word of a depot from SpaceX.Well, if an MCT is going to act as the second stage of the BFR, it has to refuel in orbit. And if people launch on the MCT, they can't wait for multiple tankers (i.e. MCTs acting as tankers) to come and fill it up.
Depots have a big advantage with hydrogen since hydrogen wants to boil-off really bad, and it takes fancy equipment like a mult-layer sunshield and an active cooler to stop that. But both methane and oxygen are space-storable, meaning with the right type of paint and by keeping your tanks out of direct sunlight (point the long way, butt to the Sun), you can get passive zero boil-off.
So I really don't think SpaceX is planning a depot. But things change.
Well, if an MCT is going to act as the second stage of the BFR, it has to refuel in orbit. And if people launch on the MCT, it can't wait for multiple tanker MCTs to come and fill it up.
Waiting around in LEO when you're supposed to go to Mars is wasteful for the life support systems.Well, if an MCT is going to act as the second stage of the BFR, it has to refuel in orbit. And if people launch on the MCT, it can't wait for multiple tanker MCTs to come and fill it up.
Why not? They're going to be in space for the whole trip to Mars, what's a few more weeks in LEO? (Speaking of which, do we know if the refuelling is to be in LEO or somewhere else?)
Waiting around in LEO when you're supposed to go to Mars is wasteful for the life support systems.
I don't get the impression that Musk is planning a big rocket just to have a big rocket, but that the size will be perfectly well-justified based on the total IMLEO needed for full-scale MCT operations (on the order of a million tons per year IMLEO).Waiting around in LEO when you're supposed to go to Mars is wasteful for the life support systems.
Really? Seems like a much cheaper option than depots. Don't get me wrong, I'm all for depots - but because they imply a different architecture to "biggest rocket evar".
What if...ionmars'll gonna hate that... The fuel carriers 2, 3 and 4 transfer The fuel directly to carrier 1 each time, then 1 transfers to the incoming MCT? Os that feasible?I was just about to suggest exactly that. One tanker *is* the depot; you end up doing N launches to put a completely full tanker in orbit, then launch a crew+cargo MCT, fill it from the full tanker, and off we go..
What if...ionmars'll gonna hate that... The fuel carriers 2, 3 and 4 transfer The fuel directly to carrier 1 each time, then 1 transfers to the incoming MCT? Os that feasible?I was just about to suggest exactly that. One tanker *is* the depot; you end up doing N launches to put a completely full tanker in orbit, then launch a crew+cargo MCT, fill it from the full tanker, and off we go..
alternatively, you could daisy-chain:
launch tanker #1
launch tanker #2 to rendezvous with #1; transfer from #1 to #2; land #1
launch tanker #3 to rendezvous with #2; transfer from #2 to #3; land #2
(repeat until there is a full tanker in orbit)
launch cargo/crew vessel to rendezvous with tanker #N; transfer fuel, land #N
With that scheme, you do more pumping, but all tankers spend about the same amount of time in orbit..
Using a BFS tanker variant for orbital depot is thinking small.
And probably, the BFR 1st stage will not be designed for leaving the upper atmosphere.If you include the ionosphere in that you are correct, but like an F9 RTLS booster I expect it to go above 200km.
We've read elsewhere (Reddit) of really HUGE BFRs. So large that if you filled even a fraction of their height with propellant way over 50 engines would be needed to lift off. Here, the speculation has mostly centered around BFRs with 25-30ish engines. Still far larger than Saturn V or Nova class.
We've read elsewhere (Reddit) of really HUGE BFRs. So large that if you filled even a fraction of their height with propellant way over 50 engines would be needed to lift off. Here, the speculation has mostly centered around BFRs with 25-30ish engines. Still far larger than Saturn V or Nova class.
IF the optimum Raptor engine thrust size still remains ~230 metric tonnes force, is SX looking at possibly smaller, less engine/plumbing complex BFRs in the high teens to 21ish # of engines for most economical sized launchers? As long as the BFR can put the 2nd stage dry weight into LEO it's big enough. Or maybe even a couple launches plus in orbit assembly. Cargo, fuel, etc. could launch on subsequent flights. Of course these "small" BFRs would have to be so much less expensive to build, maintain and fly that many times more refueling flights per Mars transit launch still makes economic sense. Arguing against my own question, it seems to me that the largest launcher you can build and fly repeatedly would be the most economical. Flight operations costs are not zero, plus they would include support of LEO propellant & cargo transfer ops.
Given Elon's Day One focus on economics, I find it difficult to believe that size/complexity/cost tradeoffs aren't fundamental to their ongoing architectural analysis. September's reveal will be beyond interesting.
Shotwell mentioned about BFR a few months ago at the South Summit 2015 (Oct 7-9), in Madrid, " [Falcon Heavy] This one is about 4M pounds of thrust, and the mock... the vehicle that takes us to Mars will be three or four times that size"
https://youtu.be/omBF1P2VhRI?t=10m46s
(original video, mostly Spanish-language conference proceeding, but Shotwell's voice still appears beneath a title graphic for the first ten minutes, though not her face. The video I linked above seems to have been created a while after this one was promoted, and does a proper cut to her presentation alone)
I also vaguely remember her mentioning offhand that they were developing a 180-210t to LEO superheavy launcher. I've been trying to find the interview, but can't turn anything up.
We've read elsewhere (Reddit) of really HUGE BFRs. So large that if you filled even a fraction of their height with propellant way over 50 engines would be needed to lift off. Here, the speculation has mostly centered around BFRs with 25-30ish engines. Still far larger than Saturn V or Nova class.
Do you know how big some Nova designs were? There were dozens and dozens of them.
http://astronautix.com/lvs/novamm1c.htm
Nova MM 1C:
Gross mass: 11,516,800 kg (25,390,100 lb).
Payload: 444,000 kg (978,000 lb).
Height: 119.00 m (390.00 ft).
Diameter: 21.00 m (68.00 ft).
Thrust: 144,157.50 kN (32,407,895 lbf).
And this wasn't some internet amateurs' fantasy. It was a study by Martin Marietta.
The F-1A engine was actually built and developmental work done on the M-2.
...the Nova we're talking about is the one that LC39a and b were sized for.
...Nova class..."
Maybe the "100 t payload" is the total land mass, a 60 t dry mass mct 25 t 100 people (+their goods and consumables) and 15t cargo or more cargo and less people. You engineering fellows, is that feasible with 3-4 refueling cargos?
Maybe the "100 t payload" is the total land mass, a 60 t dry mass mct 25 t 100 people (+their goods and consumables) and 15t cargo or more cargo and less people. You engineering fellows, is that feasible with 3-4 refueling cargos?
Maybe the "100 t payload" is the total land mass, a 60 t dry mass mct 25 t 100 people (+their goods and consumables) and 15t cargo or more cargo and less people. You engineering fellows, is that feasible with 3-4 refueling cargos?
BTW, note the cryocooling requirement kicks in with just about any orbital refueling architecture, probably even for just 1 tanker flight and propellant transfer.
Which argues strongly for a depot from my point of view along with the idea that in the early days when only a few BFRs exist having a depot could allow for a significant increase in the number of BFSs that can launch in one window.
Maybe the "100 t payload" is the total land mass, a 60 t dry mass mct 25 t 100 people (+their goods and consumables) and 15t cargo or more cargo and less people. You engineering fellows, is that feasible with 3-4 refueling cargos?
Which argues strongly for a depot from my point of view along with the idea that in the early days when only a few BFRs exist having a depot could allow for a significant increase in the number of BFSs that can launch in one window.
An orbital depot is probably a requirement just because it would be very inefficient for BFS and/or the tankers to have to lug around the cryocooling equipment.
In regards to the MCT Depot, I think we are in the early planning stages where general issues can be discussed before any commitments will be made. I am trying to devise a few tools to help the discussion.
Attached is an simple Excel spreadsheet that calculates some data for this purpose. The inputs are the number of berths (Line 2), the assumed diameter of MCT (line3), and the minimum acceptable gravity force at the tank location closest to the depot center (line15). Some of the data of interest to me are depot diameter, the highest gravity force that will be recorded at the outermost edge of each tank, and the rotational acceleration required to achieve these gravity forces. These are just a few of the factors to be considered. Please check it out.
It will have to cryocool propellant on Mars, which is more difficult than cryocooling in heliocentric orbit.
PSA: Methane and oxygen don't need cryocooling to achieve zero boil off in orbit.
Thank you.
They don't if both tanks are cool, which can be done passively and without a big sunshield.PSA: Methane and oxygen don't need cryocooling to achieve zero boil off in orbit.
Thank you.
No but they do in transfer operations and to avoid the need for cryocooling to achieve zero boil off in orbit you need a attitude control and reflective shielding and radiative surface area that does not get that radiation reflected back at it.
They don't if both tanks are cool, which can be done passively and without a big sunshield.PSA: Methane and oxygen don't need cryocooling to achieve zero boil off in orbit.
Thank you.
No but they do in transfer operations and to avoid the need for cryocooling to achieve zero boil off in orbit you need a attitude control and reflective shielding and radiative surface area that does not get that radiation reflected back at it.
Pressure. Seriously, they already have to worry about hydrazine freezing in tanks. Just paint your spacecraft white and keep it out of the Sun except the skinny way. It'll naturally get that cold.Hydrazine freezes at 2 °C! (not relevant)
If the receiving tank is below boiling point, it'd have a lower pressure and could actually condense the propellant. No venting required.
Also remember that the figures you're using are at STP. Higher boiling point at 100psi ullage pressure.
Anyway, it's well understood that methane/LOx are "space storable" as XCOR advertises. Sick of arguing with people thinking that because boiloff s big for hydrogen that it must also be for methane and oxygen.
If the receiving tank is below boiling point, it'd have a lower pressure and could actually condense the propellant. No venting required.
Also remember that the figures you're using are at STP. Higher boiling point at 100psi ullage pressure.
Anyway, it's well understood that methane/LOx are "space storable" as XCOR advertises. Sick of arguing with people thinking that because boiloff s big for hydrogen that it must also be for methane and oxygen.
How do the rest of you account for this?Maybe the "100 t payload" is the total land mass, a 60 t dry mass mct 25 t 100 people (+their goods and consumables) and 15t cargo or more cargo and less people. You engineering fellows, is that feasible with 3-4 refueling cargos?
http://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20150003005.pdf offers 12 tons consumables budget for a 4-person conjunction-class (1030 day) Mars mission using closed-loop ECLSS. You're offering 25 tons for 100 people.
How do the rest of you account for this?Maybe the "100 t payload" is the total land mass, a 60 t dry mass mct 25 t 100 people (+their goods and consumables) and 15t cargo or more cargo and less people. You engineering fellows, is that feasible with 3-4 refueling cargos?
http://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20150003005.pdf offers 12 tons consumables budget for a 4-person conjunction-class (1030 day) Mars mission using closed-loop ECLSS. You're offering 25 tons for 100 people.
Maybe the "100 t payload" is the total land mass, a 60 t dry mass mct 25 t 100 people (+their goods and consumables) and 15t cargo or more cargo and less people. You engineering fellows, is that feasible with 3-4 refueling cargos?
http://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20150003005.pdf offers 12 tons consumables budget for a 4-person conjunction-class (1030 day) Mars mission using closed-loop ECLSS. You're offering 25 tons for 100 people.
In regards to the MCT Depot, I think we are in the early planning stages where general issues can be discussed before any commitments will be made. I am trying to devise a few tools to help the discussion.What is the safety circle in ths case?
Attached is an simple Excel spreadsheet that calculates some data for this purpose. The inputs are the number of berths (Line 2), the assumed diameter of MCT (line3), and the minimum acceptable gravity force at the tank location closest to the depot center (line15). Some of the data of interest to me are depot diameter, the highest gravity force that will be recorded at the outermost edge of each tank, and the rotational acceleration required to achieve these gravity forces. These are just a few of the factors to be considered. Please check it out.
Why has LOX/RP-1 (or even LOX/methane) historically not been used for delayed burn operations, which have almost always, if not always, used hydrazine?
How do the rest of you account for this?Maybe the "100 t payload" is the total land mass, a 60 t dry mass mct 25 t 100 people (+their goods and consumables) and 15t cargo or more cargo and less people. You engineering fellows, is that feasible with 3-4 refueling cargos?
http://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20150003005.pdf offers 12 tons consumables budget for a 4-person conjunction-class (1030 day) Mars mission using closed-loop ECLSS. You're offering 25 tons for 100 people.
As you can see in the joined graph (Ames research center trajectory browser), most synods offer 120 to 180 day missions to Mars. So if you do the math, you will find 52 tonnes of consumables for 180 days with 5 kg/person. As mentionned by others, you will only be having 100 passengers when there is a base in place, so the rest of the trip time is not applicable.How do the rest of you account for this?Maybe the "100 t payload" is the total land mass, a 60 t dry mass mct 25 t 100 people (+their goods and consumables) and 15t cargo or more cargo and less people. You engineering fellows, is that feasible with 3-4 refueling cargos?
http://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20150003005.pdf offers 12 tons consumables budget for a 4-person conjunction-class (1030 day) Mars mission using closed-loop ECLSS. You're offering 25 tons for 100 people.
Depends on your assumptions for "closed loop".
3 tons per person for ~1000 days isn't a very good closed loop. A person needs about 5 kg/day of life support consumables, so 3000 kg for 1000 days means you only save 40% with the closed loop... huh?
Also, the MCT is probably carrying supplies for 100-250 days (depending on transit length) not 1000. Remember there are supposed to be ~10 cargo missions per crew mission (or something like that).
...one of the reason I dislike that tool is that it isn't really made for looking for short transits, and the numbers you get can be very misleading (including how it optimizes things) unless you're really careful. Just keep in mind that it's not telling you the limits of what can be done. In actuality, 100 day transits are available except for maybe 20% of synods, and even then opportunities are available for less than 120 days.As you can see in the joined graph (Ames research center trajectory browser), most synods offer 120 to 180 day missions to Mars. So if you do the math, you will find 52 tonnes of consumables for 180 days with 5 kg/person. As mentionned by others, you will only be having 100 passengers when there is a base in place, so the rest of the trip time is not applicable.How do the rest of you account for this?Maybe the "100 t payload" is the total land mass, a 60 t dry mass mct 25 t 100 people (+their goods and consumables) and 15t cargo or more cargo and less people. You engineering fellows, is that feasible with 3-4 refueling cargos?
http://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20150003005.pdf offers 12 tons consumables budget for a 4-person conjunction-class (1030 day) Mars mission using closed-loop ECLSS. You're offering 25 tons for 100 people.
Depends on your assumptions for "closed loop".
3 tons per person for ~1000 days isn't a very good closed loop. A person needs about 5 kg/day of life support consumables, so 3000 kg for 1000 days means you only save 40% with the closed loop... huh?
Also, the MCT is probably carrying supplies for 100-250 days (depending on transit length) not 1000. Remember there are supposed to be ~10 cargo missions per crew mission (or something like that).
To reduce to 25 tonnes you need to go to dehydrated foods, and do some fierce water and atmospheric recycling. But that equipement will be essential on Mars, so it should be part of a colony package.
How do the rest of you account for this?Maybe the "100 t payload" is the total land mass, a 60 t dry mass mct 25 t 100 people (+their goods and consumables) and 15t cargo or more cargo and less people. You engineering fellows, is that feasible with 3-4 refueling cargos?
http://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20150003005.pdf offers 12 tons consumables budget for a 4-person conjunction-class (1030 day) Mars mission using closed-loop ECLSS. You're offering 25 tons for 100 people.
Depends on your assumptions for "closed loop".
3 tons per person for ~1000 days isn't a very good closed loop. A person needs about 5 kg/day of life support consumables, so 3000 kg for 1000 days means you only save 40% with the closed loop... huh?
Also, the MCT is probably carrying supplies for 100-250 days (depending on transit length) not 1000. Remember there are supposed to be ~10 cargo missions per crew mission (or something like that).
My guess is the breakdown of the 5 kg is :Why are you guys guessing and assuming and conjecturing when you could just be reading the NASA technical report on the subject I conveniently linked and picking apart individual numbers therein?
0,7 kg of oxygen, turned into CO2
3 kg of water
1,3 kg of food,
So there would be no allocation for recycling at all?
The oxygen can be recycled from CO2 using the sabatier process and some hydrogen, either stock of from the water coming from the food combustion. As long as the unit weighs less that ,7*100*180 = 14 tonnes, including solar cells to run it, then we're ahead on that.
The water should be easy, the ISS already does that. We need to treat 300 kg per day, plus whatever is used for sanitation and cleaning. We can allow for some of it to get dirty, since even dirty water will be a good Martian import. Again 5x100x180 = 52 tonnes, and we should be able to have a good cleaning system for much less mass than that.
The food will become compost. It'll be recycled on the colony.
Although overall water recovery rates are less than 100% for the assumed ECLSS system, there is a net surplus of water produced. This surplus occurs because additional water is added to the system in the form of water in the food that the crew consumes. Although the food is “dehydrated” it still contains approximately 28% water. The result is that, under the assumptions made for the study, no additional water needs to be added to satisfy water or oxygen generation requirements. Further closure of the ECLSS system will not reduce total logistics requirements.
Water reclamation from H2O contained within food promotes water-rich operating conditions for the partially closed ECLSS. As such, only 30 days of contingency water and oxygen were required to be delivered with the habitat, resulting in 362kg and 99kg of water and oxygen required
respectively
>
On the other hand, long-run hygiene and laundry needs in excess of this report are something I've heard speculation about.
One snippet from the report states that water needs are filled mostly by food water content based on presentday space-food menus:
A) Where do you get 5kg/day?How do the rest of you account for this?Maybe the "100 t payload" is the total land mass, a 60 t dry mass mct 25 t 100 people (+their goods and consumables) and 15t cargo or more cargo and less people. You engineering fellows, is that feasible with 3-4 refueling cargos?
http://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20150003005.pdf offers 12 tons consumables budget for a 4-person conjunction-class (1030 day) Mars mission using closed-loop ECLSS. You're offering 25 tons for 100 people.
Depends on your assumptions for "closed loop".
3 tons per person for ~1000 days isn't a very good closed loop. A person needs about 5 kg/day of life support consumables, so 3000 kg for 1000 days means you only save 40% with the closed loop... huh?
Also, the MCT is probably carrying supplies for 100-250 days (depending on transit length) not 1000. Remember there are supposed to be ~10 cargo missions per crew mission (or something like that).
Soylent includes fiber.
Remember, this is for 3 months, not for surface living.
People can pay more for more food, then. Problem solved. But for those willing to skimp, save $10,000-100,000 and go with Soylent on the way there, that option should be available to them.
Soylent tends to increase flatulence, from what I've read. We really don't want unpleasant conditions for all during the trip.
They wouldn't let another methane source go to waste, would they?Soylent tends to increase flatulence, from what I've read. We really don't want unpleasant conditions for all during the trip.
They can just open a window. ;D
A) Where do you get 5kg/day?
B) Wait, what? 10 cargo missions per crew mission... of MCT?
And how many propellant missions per crew mission is that?
If you limited it to 6km/s each = 5:1 mass ratio @ 380s (probably not reasonable, given 3.2 C3, 0.5 minimum to MTO, more for single-synod reuse, and largely propulsive SSRP EDL), 40 prop launches to feed 10 cargo launches
"Why would anyone want to live on Mars?? I wouldn't want to! This is a fool's errand and should be stopped!"
"Why would anyone want to just drink Soylent?? I wouldn't want to! This is a fool's errand and should be stopped!"
If you're going to Mars for years, drinking a protein shake on the way is going to be the very least of your psychological worries, I PROMISE you. Ridiculous.
Flatulence is a great resource for ISRO production of propellant. ;)
If you're going to Mars for years, drinking a protein shake on the way is going to be the very least of your psychological worries, I PROMISE you. Ridiculous.
If you're going to Mars for years, drinking a protein shake on the way is going to be the very least of your psychological worries, I PROMISE you. Ridiculous.I am reminded of a proverb involving the backs of camels, straws and breaking.
I am reminded of a proverb involving the backs of camels, straws and breaking.
People will tolerate a lot of discomfort, but what will finally break a person is usually something seeming relatively petty in the grand scheme of things.
Hmmm. Lots of unpleasant things I can do to these econocolonists.
"A few months" is still a very long time for a human to live through. Remember, there's not much to really do during the transfer. Sure, there's exercise, some leisure activities (reading, watching TV, various games, et cetera) and the occasional maintenance work to occupy your time. Boredom might be a serious morale issue and you do not want to compound that issue with a bland diet.I am reminded of a proverb involving the backs of camels, straws and breaking.
People will tolerate a lot of discomfort, but what will finally break a person is usually something seeming relatively petty in the grand scheme of things.
Yes, but we are talking about a few months of transfer. That the diet on Mars needs to be a lot better for colonists I think no one would dispute. I don't think that Soylent is the way to go but it is an option for the trip.
I agree that going to a super restricted diet wouldn't be good. But consumables mass could still be quite low without that extreme.A) Where do you get 5kg/day?
Basic human life support requirements per day are about 5kg - less than 1kg oxygen, less than 1kg dry food, and over 3kg water (including water in food).
This assumes either doing something other than normal bathing, or 100% recycling of cleaning water.QuoteB) Wait, what? 10 cargo missions per crew mission... of MCT?
I believe Musk tweeted that.QuoteAnd how many propellant missions per crew mission is that?
I don't know... depends on the size of MCT and the architecture.QuoteIf you limited it to 6km/s each = 5:1 mass ratio @ 380s (probably not reasonable, given 3.2 C3, 0.5 minimum to MTO, more for single-synod reuse, and largely propulsive SSRP EDL), 40 prop launches to feed 10 cargo launches
That assumes that the MCT arrives in LEO empty.
If BFR is really too large for any existing pad, I would expect MCT to either go to a higher energy orbit than LEO or arrive with significant propellant.
But cost per launch will nonetheless have to be incredibly low by current standards.
I drink Soylent for lunch. This is not "drinking 3000 Calories of vegetable oil."The first principle in Musk's mind is cost followed by almost eliminating training, which is also a major cost. That would include using normal diet, fresh like foods. Each MCT would need a core group that is the operator/maintainers but most of the group of 100 are just passengers along for the ride with little training. Many of the specialists/crew may not be colonists and return. Similar to a ocean passenger ship crew.
Building a colony transporter with the economics needed to make it feasible is not something that has ever been done before. It most CERTAINLY will involve thinking from first principles, as Musk is wont to do.
The NASA estimates are nowhere near a normal diet or fresh-like foods. Those foods are shelf-stabilized in some way, and there's no food freezer or refrigerator, and they're dehydrated wherever it's practical. 28% water content by mass is extremely low for foods that can be eaten without only a minimal kitchen.I drink Soylent for lunch. This is not "drinking 3000 Calories of vegetable oil."The first principle in Musk's mind is cost followed by almost eliminating training, which is also a major cost. That would include using normal diet, fresh like foods. Each MCT would need a core group that is the operator/maintainers but most of the group of 100 are just passengers along for the ride with little training. Many of the specialists/crew may not be colonists and return. Similar to a ocean passenger ship crew.
Building a colony transporter with the economics needed to make it feasible is not something that has ever been done before. It most CERTAINLY will involve thinking from first principles, as Musk is wont to do.
I drink Soylent for lunch. This is not "drinking 3000 Calories of vegetable oil."The problem is not thinking from first principles, it's rejecting actual research that's been done and the optimum that's been chosen in a program where every kilogram costs $30,000 extra, which has plenty of incentive to minimize mass and ramp up the ECLSS; And ensuring everyone that this is all an irrelevant, useless assessment, because you choose to drink Soylent for lunch sometimes.
Building a colony transporter with the economics needed to make it feasible is not something that has ever been done before. It most CERTAINLY will involve thinking from first principles, as Musk is wont to do.
I drink Soylent for lunch. This is not "drinking 3000 Calories of vegetable oil."The problem is not thinking from first principles, it's rejecting actual research that's been done and the optimum that's been chosen in a program where every kilogram costs $30,000 extra, which has plenty of incentive to minimize mass and ramp up the ECLSS; And ensuring everyone that this is all an irrelevant, useless assessment, because you choose to drink Soylent for lunch sometimes.
Building a colony transporter with the economics needed to make it feasible is not something that has ever been done before. It most CERTAINLY will involve thinking from first principles, as Musk is wont to do.
A space freezer can be an air lock opened to space for some foods.
Food has been consistently rated the most important morale boosting thing Astronauts get
Completely agree.I drink Soylent for lunch. This is not "drinking 3000 Calories of vegetable oil."The problem is not thinking from first principles, it's rejecting actual research that's been done and the optimum that's been chosen in a program where every kilogram costs $30,000 extra, which has plenty of incentive to minimize mass and ramp up the ECLSS; And ensuring everyone that this is all an irrelevant, useless assessment, because you choose to drink Soylent for lunch sometimes.
Building a colony transporter with the economics needed to make it feasible is not something that has ever been done before. It most CERTAINLY will involve thinking from first principles, as Musk is wont to do.
Yes, a lot of effort has gone into this subject. No need to throw that reseach out the airlock. NASA has a variety of dehydrated foods that are very mass efficient. So does private industry. While Soylent or something like it can be part of the food selection, there is no reason to make it the only food.
..meat can be freeze-dried, but reconstituting it turns it into a slurry.....
Completely agree.I drink Soylent for lunch. This is not "drinking 3000 Calories of vegetable oil."The problem is not thinking from first principles, it's rejecting actual research that's been done and the optimum that's been chosen in a program where every kilogram costs $30,000 extra, which has plenty of incentive to minimize mass and ramp up the ECLSS; And ensuring everyone that this is all an irrelevant, useless assessment, because you choose to drink Soylent for lunch sometimes.
Building a colony transporter with the economics needed to make it feasible is not something that has ever been done before. It most CERTAINLY will involve thinking from first principles, as Musk is wont to do.
Yes, a lot of effort has gone into this subject. No need to throw that reseach out the airlock. NASA has a variety of dehydrated foods that are very mass efficient. So does private industry. While Soylent or something like it can be part of the food selection, there is no reason to make it the only food.
Regarding weight and water content in food: how about freeze drying?While freeze-drying saw a number of earlier applications in food processing & medical products, eating freeze-dried food was basically popularized by products developed for NASA's short-duration spaceflights in the 60's. If you ever go to the NASM you'll probably walk away with souvenir 'Astronaut Ice Cream', developed under contract for NASA in 1968, or a package of freeze dried strawberries.
This is handled in the PBS link.
What's the sodium content of the NASA foods? The commercial versions have near edema inducing levels.
A system that cleans water (maybe via active coal, can be regenerated, and reverse osmosis) allows reduction of the mass without losing too much of the quality.
..meat can be freeze-dried, but reconstituting it turns it into a slurry.....
I spent a lot of summers of college and grad school working as a wilderness guide. I never had much problem with any kind of freeze dried meat: beef, chicken, ham, shrimp. And I ate a lot of it over the years.
ESA recently held a meeting to discuss possible human hibernation:Hibernation is a dark horse option. The tiny chance we can get it working multiplied by the huge difference it would make means scaling up a very large space program another 1% looks like a bad investment relative to taking the risk with low-hanging-fruit hibernation research... but it's much, much more speculative than MCT. It cannot be a critical-path element.
http://motherboard.vice.com/read/a-brief-history-of-cryosleep (http://motherboard.vice.com/read/a-brief-history-of-cryosleep)
Obviously, that is still fairly speculative technology. But given the enormous practical burden of carrying large amounts of food/water and handling the resulting huge amount of human waste, it warrants serious consideration for MCT. Hibernation would also eliminate potential psychological issues, like boredom, in transit.
It is stating the obvious IMO that at some point the Mars colony has to account for the system-level reality of growing real food and fully recycling the waste... Ya know, an ecosystem.
Assuming this much, I think it's useful to ask at what point in time do you start planning for the initial steps to happen? If you're going to have this discussion, it's also useful not to talk in absolutes, since it's likely even the first missions will at least have some experimental setup designed to produce something which can be consumed, and at the other end of the scale, even centuries from now, there will be some things which cannot be grown on Mars and are still imported from Earth.
Between those two points, theres a t50... at time point at which half the food is grown at Mars, and half is imported.
I think it's likely that if that t50 occurs during the operational lifetime of the MCT (say the next 3 decades), and it's probable that the MCT itself will incorporate a lot of that technology, in which case we can expect the design process would be in progress now.
There are other threads where agriculture on Mars would fit far better than here. Look in the general Mars section.
Sci-fi gets exactly three plausible means of interstellar and outer-system spaceflight for biological humans: Hibernation, new-physics FTL or new-physics CoE/CoM breaking engines, and generation ships.
There are other threads where agriculture on Mars would fit far better than here. Look in the general Mars section.
This is the active thread on those matters. Plenty of interesting posts. Please get the discussion there.
http://forum.nasaspaceflight.com/index.php?topic=35877.0
Why would an MCT arrive in LEO (with around 200 tons, per Shotwell) packing 50 tons of propellant in place of 50 tons of gear?
What does BFR being too large for any existing pad have to do with it?
Why would an MCT arrive in LEO (with around 200 tons, per Shotwell) packing 50 tons of propellant in place of 50 tons of gear?
I was imagining something significantly more than 50 tons.QuoteWhat does BFR being too large for any existing pad have to do with it?
Because if I run the rocket equation with 100 tons of payload, dV = 9500 m/s, Raptor's claimed Isp, and dry masses that I consider pretty conservative for SpaceX, I get a rocket that's not that big.
Which suggests to me that the delta-V will be a lot higher - either it'll go to a much higher energy orbit or have quite a bit left.
No it can't. "Space" isn't cold. It's a vacuum. If you keep your food in an open airlock, it will have the same temperature as any other section of the ship's hull...
Astronauts are crew. Crew morale is important. Colonists are cargo. Only their survival rates matter.
Why would an MCT arrive in LEO (with around 200 tons, per Shotwell) packing 50 tons of propellant in place of 50 tons of gear?
I was imagining something significantly more than 50 tons.QuoteWhat does BFR being too large for any existing pad have to do with it?
Because if I run the rocket equation with 100 tons of payload, dV = 9500 m/s, Raptor's claimed Isp, and dry masses that I consider pretty conservative for SpaceX, I get a rocket that's not that big.
Which suggests to me that the delta-V will be a lot higher - either it'll go to a much higher energy orbit or have quite a bit left.
Don't forget to save fuel for reuse.
You are not alone in getting such "not that big" results. Right now we're all over the map with possibly conflicting years old Elon statements and crazy big BFR/MCT rumors on Reddit. But the bottom line is that it does not even take the "mid sized" :) 15 million LB thrust vehicle to meet the LEO payload claims.
Postponing the architecture announcement for the 3rd time indicates to me that everything is in a high state of flux. They've likely done some next level of detail of engineering analysis and arrived at some key numbers different than expected, which iterates revisions. With Musk the decisions will not be just space cadet tech driven but will also have a strong best economic model (as he best believes it) influencing size/capability tradeoffs.
BFR is much more than 100 tons to LEO. That much is clearly established.
BFR is much more than 100 tons to LEO. That much is clearly established.
The 100 tonnes figure was the useful payload landed on Mars IIRC. This is by a ship/lander that will refuel and depart sans 100 tonnes payload. The mass to LEO will vary considerably(though always be much larger than 100 tonnes) depending how refueling is incorporated and how you account for the spaceship itself.
Hello.We think a bunch of Raptors. The Falcon 9 uses nine Merlin 1D's in the first stage and a single slightly different Merlin 1D (Vacuum Optimized) in the second stage.
I have a question, I know you had probably talk about this but I cant find it.
What kind of propulsion will have second stage?
Once they came back, they can never come back...
I'd also assumed RTLS in my past calculations but have come to believe that the 1st stage will instead land on a floating platform, as less propellant is needed with no boost-back.
Technical query for my BFS models.
What's a rational target range for T/W of the 2nd stage at ignition taking into account gravity losses, etc.?
I would expect it could be less than a 1st stage's preferred T/W.
How low can it be at ignition?
The 2nd stage of this supposed TSTO has near but not equal to Earth SSTO performance and is a SSTO when departing Mars' surface.
The 2nd stage of this supposed TSTO has near but not equal to Earth SSTO performance and is a SSTO when departing Mars' surface.
Can you clarify if you actually mean SSTLMO (to Low Mars Orbit) or SSTTEI (to trans-Earth-Injection). The latter is what everyone claiming integrated 2nd stage seems to be aiming for and it tends to yield a mars take off mass >1000 mt. The former is my position and involves SEP transit vehicles or refueling in mars orbit with SEP delivered propellants and keeps mars take-off mass to ~400 mt.
... How realistic is this?
...
I read that novel too. But... Colonists are people. It will only take one going koo-koo to jeopardize everyone and the ship.
...
I read that novel too. But... Colonists are people. It will only take one going koo-koo to jeopardize everyone and the ship.
And if you catch the kook in time, you run a quick experiment to determine how long someone lasts outside the airlock.
I like the way he listed space tourism as something *other* companies are doing.. and he specifically mentioned orbital space tourism in that list. I'd love to know when SpaceX decided they were too good for this market.
I like the way he listed space tourism as something *other* companies are doing.. and he specifically mentioned orbital space tourism in that list. I'd love to know when SpaceX decided they were too good for this market.
I like the way he listed space tourism as something *other* companies are doing.. and he specifically mentioned orbital space tourism in that list. I'd love to know when SpaceX decided they were too good for this market.
By too good, do you really mean not really interested in it because it's not what they want to do?
Or did you deliberately use the phrase "too bad" just to have a strawman dig at SpaceX?
I like the way he listed space tourism as something *other* companies are doing.. and he specifically mentioned orbital space tourism in that list. I'd love to know when SpaceX decided they were too good for this market.
Airlines and aircraft builders like Boeing and Airbus are different companies. The same applies to the builders and operators of ships, buses, trains and taxis. It was only a matter of time before space tourism was separated from launch vehicle construction.
The 3 km/s separation seems to be the more likely scenario and the T/W ratio sounds reasonable at 0.75 so we can extrapolate different masses for the 2nd stage at separation based on an engine count and the target Raptor thrust.
7 Raptors: 16100 kN thrust, 2200 mt mass at separation.
6 Raptors: 13800 kN thrust, 1880 mt mass at separation.
5 Raptors: 11500 kN thrust, 1570 mt mass at separation.
4 Raptors: 9200 kN thrust, 1260 mt mass at separation.
I believe the 5 engine configuration is getting on the small end, my bet would be to use a hexagonal 6 engine arrangement which would provide a space for a smaller central landing engine (I dub this mini engine 'Robin').
The 3 km/s separation seems to be the more likely scenario and the T/W ratio sounds reasonable at 0.75 so we can extrapolate different masses for the 2nd stage at separation based on an engine count and the target Raptor thrust.
7 Raptors: 16100 kN thrust, 2200 mt mass at separation.
6 Raptors: 13800 kN thrust, 1880 mt mass at separation.
5 Raptors: 11500 kN thrust, 1570 mt mass at separation.
4 Raptors: 9200 kN thrust, 1260 mt mass at separation.
I believe the 5 engine configuration is getting on the small end, my bet would be to use a hexagonal 6 engine arrangement which would provide a space for a smaller central landing engine (I dub this mini engine 'Robin').
Isn't that a bit too big? Wouldn't a 1260 mt BFS second stage, fully fueled, be enough to lift ~200 mt (50% of it being structural mass)?
And for the return trip, a 4 Raptor BFS would have a T/W > 1 in Mars anyway, even with full tanks and cargo.
I like your 'Robin' idea, I wonder if it could be a pressure fed, for safety reasons, lox/methane engine, a bit like the Superdracos but using a different propellant (Hyperdracos, anyone?). You don't need super efficient engines to reach LMO.
jsgirald:
Yes with a 1260 mt mass you would be looking at around a ~200 mt burn out mass at LEO assuming that acceleration after separation is 6 km/s. At 1880 mt separation mass the burn out mass it approaching 375 mt but that is an upper limit. A 6 engine configuration allows for engine out (which after symmetrically shut down drops us to 4 engines). As I expect the vehicle to be launched with cargo that requires a dry mass delivery in LEO of nearly 200 mt, thus the 4 engine configuration provides no margin for a shutdown not to mention the need to retain landing propellant when the goal is to delivery propellant to a depot. Lastly I see the robin engine as being a sub-scale Raptor running on Methane and Lox.
-snip-
Thank you, the intent is to think about the problem constructively. I'm sure that Elon's MCT architecture will be completely different but you never know where ideas come from.
* It's very difficult to parse your diagram. You should illustrate some more of the things you're captioning.
* I'm not sure anyone really considers an 8.4m diameter plausible for this vehicle anymore. What we know of the requirements imposed on the vehicle by most the mission architectures suggests there's too much rocket here for things to fit into 8.4m without making the vehicle extraordinarily long and thin and difficult to land. We examined 10, 12.5, and 15 meters and mostly concluded that we were looking at the upper part of that range, just based on how big the BFR would need to be to get it into LEO.
* If you could include more of your figures and timelines for the mission architecture, we would be able to critique that a lot more productively.
It's nice to see more people thinking about the problem constructively, in any case.
If you reuse each vehicle 10 times then slightly higher manufacturing costs are not as significant as other operational costs. So SpaceX assumption of the vehicle being operated sole as a reusable vehicle will weigh heavily on how it is designed and how it would be manufactured.
I'm late to this but I wonder, apart from Karou's estimate has anyone looked at physically how 100 people lying down map into a cylinder? I'm guessing about 8'x3'. A square 8.2m gives 24 spaces, a circle is likely to give substantially less.I got my numbers from NASA for the sleep stations on ISS. I assume that if one can live aboard ISS for months on end, it's good for MCT passengers. The sleep station is more than a bed, it doubles as a work/relaxation space. It doesn't have to be rectangle or box. Think of it being an enclosed first class suite on a AirBus A380.
I expect the accumulated water, food and air will be much heavier than the passenger (without heavy recycling of all consumables) but occupy less volume (given the density of those items.
5 m^3 per person is reasonable if the trip lasted a couple of weeks, but this is going to be at least 3-4 months or more.There's a difference between personal space and common space. The personal space aboard the ISS is less than 5 m3, the rest is work space and Scott Kelly has gone almost a year with only that amount of personal space. This is not an airplane, think of it more like a submarine with a lot more personal space (at least they're not hot racking). AFAICR, ISS in the beginning didn't have (may still not have) not enough sleep stations; 2 stations I believe in the beginning for three astronauts. One of them just had to sleep wherever they could be tied down.
Based on diagrams of the original Mars Direct plan, I had calculated (http://forum.nasaspaceflight.com/index.php?topic=37733.msg1393316#msg1393316) the volume of the Earth Return Vehicle crew cabin for each of the four astronauts: 36 m^3. And for 6 months, that (https://www.google.com/search?newwindow=1&q=mars+direct+erv+cabin+too+small) was thought to be too small.
ISS total habitable volume is 388 m^3 (http://www.nasa.gov/mission_pages/station/main/onthestation/facts_and_figures.html). For 6-7 people (permanent crew, living on the station for months at a time), that is 55-65 m^3 per person.
5 m^3 per person is reasonable if the trip lasted a couple of weeks, but this is going to be at least 3-4 months or more.Musk keeps saying 3 months or 100 days. Some opportunities require a little more energy, but 120 days is fine. So not more than 4 months (except with reduced crew).
Based on diagrams of the original Mars Direct plan, I had calculated (http://forum.nasaspaceflight.com/index.php?topic=37733.msg1393316#msg1393316) the volume of the Earth Return Vehicle crew cabin for each of the four astronauts: 36 m^3. And for 6 months, that (https://www.google.com/search?newwindow=1&q=mars+direct+erv+cabin+too+small) was thought to be too small.
ISS total habitable volume is 388 m^3 (http://www.nasa.gov/mission_pages/station/main/onthestation/facts_and_figures.html). For 6-7 people (permanent crew, living on the station for months at a time), that is 55-65 m^3 per person.
For a 100 crew size with a 15m diameter 50m^3 /person is a crew cabin of 30m tall. For 30m^3 it is 17m tall.5 m^3 per person is reasonable if the trip lasted a couple of weeks, but this is going to be at least 3-4 months or more.Musk keeps saying 3 months or 100 days. Some opportunities require a little more energy, but 120 days is fine. So not more than 4 months (except with reduced crew).
Based on diagrams of the original Mars Direct plan, I had calculated (http://forum.nasaspaceflight.com/index.php?topic=37733.msg1393316#msg1393316) the volume of the Earth Return Vehicle crew cabin for each of the four astronauts: 36 m^3. And for 6 months, that (https://www.google.com/search?newwindow=1&q=mars+direct+erv+cabin+too+small) was thought to be too small.
ISS total habitable volume is 388 m^3 (http://www.nasa.gov/mission_pages/station/main/onthestation/facts_and_figures.html). For 6-7 people (permanent crew, living on the station for months at a time), that is 55-65 m^3 per person.
I'd like to know where and when did he actually say this. Because I'm not finding it (https://www.google.com/webhp?sourceid=chrome-instant&ion=1&espv=2&ie=UTF-8#q=%22musk%22%20%22100%20days%22%20%22mct%22).5 m^3 per person is reasonable if the trip lasted a couple of weeks, but this is going to be at least 3-4 months or more.Musk keeps saying 3 months or 100 days. Some opportunities require a little more energy, but 120 days is fine. So not more than 4 months (except with reduced crew).
Based on diagrams of the original Mars Direct plan, I had calculated (http://forum.nasaspaceflight.com/index.php?topic=37733.msg1393316#msg1393316) the volume of the Earth Return Vehicle crew cabin for each of the four astronauts: 36 m^3. And for 6 months, that (https://www.google.com/search?newwindow=1&q=mars+direct+erv+cabin+too+small) was thought to be too small.
ISS total habitable volume is 388 m^3 (http://www.nasa.gov/mission_pages/station/main/onthestation/facts_and_figures.html). For 6-7 people (permanent crew, living on the station for months at a time), that is 55-65 m^3 per person.
I'd like to know where and when did he actually say this. Because I'm not finding it (https://www.google.com/webhp?sourceid=chrome-instant&ion=1&espv=2&ie=UTF-8#q=%22musk%22%20%22100%20days%22%20%22mct%22).5 m^3 per person is reasonable if the trip lasted a couple of weeks, but this is going to be at least 3-4 months or more.Musk keeps saying 3 months or 100 days. Some opportunities require a little more energy, but 120 days is fine. So not more than 4 months (except with reduced crew).
Based on diagrams of the original Mars Direct plan, I had calculated (http://forum.nasaspaceflight.com/index.php?topic=37733.msg1393316#msg1393316) the volume of the Earth Return Vehicle crew cabin for each of the four astronauts: 36 m^3. And for 6 months, that (https://www.google.com/search?newwindow=1&q=mars+direct+erv+cabin+too+small) was thought to be too small.
ISS total habitable volume is 388 m^3 (http://www.nasa.gov/mission_pages/station/main/onthestation/facts_and_figures.html). For 6-7 people (permanent crew, living on the station for months at a time), that is 55-65 m^3 per person.
Also, with that fast trajectory, what will the heating loads be as the vehicle gets aerocaptured by the Martian atmosphere? And can that thin atmosphere slow it down enough?
Right, I'm sure the crew size would be much less than 100 at first. And I used to be pretty sure 500m^3 was about the right size for cabin for MCT, squished, yes, but doable for the short trip. And I still stand by that as a possible minimum size per person, at least if you're really clever with how you utilize the space (with sleep schedules rotating, everyone having their own small personal space that they spend at least 12 hours a day in), but I no longer think SpaceX is thinking as small as 500m^3. Probably 1000m^3 or perhaps more. 2000m^3 is perhaps on the high end of what I think SpaceX is thinking of, but I wouldn't argue too much with it until we get more information.For a 100 crew size with a 15m diameter 50m^3 /person is a crew cabin of 30m tall. For 30m^3 it is 17m tall.5 m^3 per person is reasonable if the trip lasted a couple of weeks, but this is going to be at least 3-4 months or more.Musk keeps saying 3 months or 100 days. Some opportunities require a little more energy, but 120 days is fine. So not more than 4 months (except with reduced crew).
Based on diagrams of the original Mars Direct plan, I had calculated (http://forum.nasaspaceflight.com/index.php?topic=37733.msg1393316#msg1393316) the volume of the Earth Return Vehicle crew cabin for each of the four astronauts: 36 m^3. And for 6 months, that (https://www.google.com/search?newwindow=1&q=mars+direct+erv+cabin+too+small) was thought to be too small.
ISS total habitable volume is 388 m^3 (http://www.nasa.gov/mission_pages/station/main/onthestation/facts_and_figures.html). For 6-7 people (permanent crew, living on the station for months at a time), that is 55-65 m^3 per person.
If it is such a short duration of travel then a crew cabin size of 3000m^3 may work. Even a possible 2000m^3 may work. But for the first missions the MCT crew cabin will be the on surface HAB module as well so if the volume is 2000m^3 then the crew numbers would need to be less than 40 for initial trips. Crew sizes of 25 has been batted around a lot for these missions.
The heating loads can be dealt with using PICA-X, and yes the thin atmosphere is enough to slow it down since you can dive deeper in. You're right to mention aerocapture, though, as you likely would want to do an aerocapture pass (perhaps two, one to capture in a high orbit, then another to put you in low orbit before doing the final descent), not go straight in for a landing like most Mars probes have done (on a hyperbolic trajectory, it's harder to slow down in time).I'd like to know where and when did he actually say this. Because I'm not finding it (https://www.google.com/webhp?sourceid=chrome-instant&ion=1&espv=2&ie=UTF-8#q=%22musk%22%20%22100%20days%22%20%22mct%22).5 m^3 per person is reasonable if the trip lasted a couple of weeks, but this is going to be at least 3-4 months or more.Musk keeps saying 3 months or 100 days. Some opportunities require a little more energy, but 120 days is fine. So not more than 4 months (except with reduced crew).
Based on diagrams of the original Mars Direct plan, I had calculated (http://forum.nasaspaceflight.com/index.php?topic=37733.msg1393316#msg1393316) the volume of the Earth Return Vehicle crew cabin for each of the four astronauts: 36 m^3. And for 6 months, that (https://www.google.com/search?newwindow=1&q=mars+direct+erv+cabin+too+small) was thought to be too small.
ISS total habitable volume is 388 m^3 (http://www.nasa.gov/mission_pages/station/main/onthestation/facts_and_figures.html). For 6-7 people (permanent crew, living on the station for months at a time), that is 55-65 m^3 per person.
Also, with that fast trajectory, what will the heating loads be as the vehicle gets aerocaptured by the Martian atmosphere? And can that thin atmosphere slow it down enough?
3000m^3 is starting to get too big (IMHO) to be practical (at some point, you're better off lowering the price and sending more people if you have that much room).
I think they may start at trips less than 100. Musk originally gave a range for # of passengers, like 50-100 or something like that. So start with 50, say, and grow to 100 as you become more efficient.3000m^3 is starting to get too big (IMHO) to be practical (at some point, you're better off lowering the price and sending more people if you have that much room).
I am quite sure Elon Musk recently said just that. That 100 people may not be the upper limit forever and they may offer "economy class" trips with more. My guess is that they would need some advances in ECLSS to pull that off.
Right, I'm sure the crew size would be much less than 100 at first. And I used to be pretty sure 500m^3 was about the right size for cabin for MCT, squished, yes, but doable for the short trip. And I still stand by that as a possible minimum size per person, at least if you're really clever with how you utilize the space (with sleep schedules rotating, everyone having their own small personal space that they spend at least 12 hours a day in), but I no longer think SpaceX is thinking as small as 500m^3. Probably 1000m^3 or perhaps more. 2000m^3 is perhaps on the high end of what I think SpaceX is thinking of, but I wouldn't argue too much with it until we get more information.For a 100 crew size with a 15m diameter 50m^3 /person is a crew cabin of 30m tall. For 30m^3 it is 17m tall.5 m^3 per person is reasonable if the trip lasted a couple of weeks, but this is going to be at least 3-4 months or more.Musk keeps saying 3 months or 100 days. Some opportunities require a little more energy, but 120 days is fine. So not more than 4 months (except with reduced crew).
Based on diagrams of the original Mars Direct plan, I had calculated (http://forum.nasaspaceflight.com/index.php?topic=37733.msg1393316#msg1393316) the volume of the Earth Return Vehicle crew cabin for each of the four astronauts: 36 m^3. And for 6 months, that (https://www.google.com/search?newwindow=1&q=mars+direct+erv+cabin+too+small) was thought to be too small.
ISS total habitable volume is 388 m^3 (http://www.nasa.gov/mission_pages/station/main/onthestation/facts_and_figures.html). For 6-7 people (permanent crew, living on the station for months at a time), that is 55-65 m^3 per person.
If it is such a short duration of travel then a crew cabin size of 3000m^3 may work. Even a possible 2000m^3 may work. But for the first missions the MCT crew cabin will be the on surface HAB module as well so if the volume is 2000m^3 then the crew numbers would need to be less than 40 for initial trips. Crew sizes of 25 has been batted around a lot for these missions.
3000m^3 is starting to get too big (IMHO) to be practical (at some point, you're better off lowering the price and sending more people if you have that much room).
According to NASA, 5m^3 over a few months is about the size volume where trained astronauts will go bonkers and kill each other, the survivors turning into Reavers who will prey on hapless interplanetary vessels. I've stayed in a capsule hotel before - that's 2.7m^3 of space per capsule, and that's really not a lot.
http://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20070023306_2007019854.pdf
The only way 5m^3 will work is if the passengers are in hibernation or some advanced form of VR or other disruptive future tech. 10m^3 *might* be doable. Nuclear submarines have 10m^3 per person with hot bunking. Certainly you will want shifts so the passengers don't get in each other's faces too much.
A horizontally manufactured large diameter tank would be differently designed and possibly weigh more than a vertically manufactured tank like the SLS. Given the problems with getting the vertical manufacturing setup to stay in alignment it may be easier to do the manufacturing horizontally (lower floor weight support required, easier to maintain alignment).I always suspected a big part of NASA's vertical tank mfg approach was the tank sections would "sag" under their own weight in different ways and NASA (during Saturn/Apollo) did not feel it had the time to investigate and decided to side step the problem altogether (and continue to do so).
The only way 5m^3 will work is if the passengers are in hibernation or some advanced form of VR or other disruptive future tech. 10m^3 *might* be doable. Nuclear submarines have 10m^3 per person with hot bunking. Certainly you will want shifts so the passengers don't get in each other's faces too much.I thought only the British did hot bunking on nuclear submarines, and they'd abandoned it as well in their latest generation boats
Sorry, wrong link.
http://msis.jsc.nasa.gov/sections/section08.htm#Figure 8.6.2.1-1
http://ston.jsc.nasa.gov/collections/trs/_techrep/TM-2011-217352.pdf
http://ston.jsc.nasa.gov/collections/trs/_techrep/TM-2015-218564.pdf
I'm being flippant when I talk about the passengers/crew wigging out, but this is something crews on nuclear subs have to be carefully screened for. The British still do hot bunking - it's just the junior crew who have to suffer through it. About 1/3 of submariners are discharged for psychological reasons after their first tour. It may be lower (I didn't check the figures) but it's still a terrible psychological toll. Put your average hipster in there and you may find a gibbering mess coming out.
Well too bad, Boomers. The young folk are going to be the ones going to Mars.At those prices it sounds like Mars will be the first new territory by trust fund beneficiaries,retirees and ex caregivers.
Maybe you would've gotten the chance if you hadn't been messing around with drugs in your youth (then racking up a big national debt and environmental damage in your adult lives)
Well too bad, Boomers. The young folk are going to be the ones going to Mars.
Maybe you would've gotten the chance if you hadn't been messing around with drugs in your youth (then racking up a big national debt and environmental damage in your adult lives) instead of building spaceships like these fantastically hard-working and bright young engineers at SpaceX (and other Newspace companies), you would've been able to go. You can just deal with it. :P
(You aren't all bad, Boomers, but if you're going to dish it out, you better be able to take it...)
Sorry, wrong link.
http://msis.jsc.nasa.gov/sections/section08.htm#Figure (http://msis.jsc.nasa.gov/sections/section08.htm#Figure) 8.6.2.1-1
http://ston.jsc.nasa.gov/collections/trs/_techrep/TM-2011-217352.pdf (http://ston.jsc.nasa.gov/collections/trs/_techrep/TM-2011-217352.pdf)
http://ston.jsc.nasa.gov/collections/trs/_techrep/TM-2015-218564.pdf (http://ston.jsc.nasa.gov/collections/trs/_techrep/TM-2015-218564.pdf)
I'm being flippant when I talk about the passengers/crew wigging out, but this is something crews on nuclear subs have to be carefully screened for. The British still do hot bunking - it's just the junior crew who have to suffer through it. About 1/3 of submariners are discharged for psychological reasons after their first tour. It may be lower (I didn't check the figures) but it's still a terrible psychological toll. Put your average hipster in there and you may find a gibbering mess coming out.
Assuming 18 people can use the same communal space as 6, then 15m^3 looks possible.
Sorry, wrong link.Much more helpful but quite disturbing. I note 2 things.
http://msis.jsc.nasa.gov/sections/section08.htm#Figure 8.6.2.1-1
http://ston.jsc.nasa.gov/collections/trs/_techrep/TM-2011-217352.pdf
http://ston.jsc.nasa.gov/collections/trs/_techrep/TM-2015-218564.pdf
Assuming 18 people can use the same communal space as 6, then 15m^3 looks possible.
1)Yes, you start out looking at the minimum to bound the problem. Obviously, there's not really a maximum.Sorry, wrong link.Much more helpful but quite disturbing. I note 2 things.
http://msis.jsc.nasa.gov/sections/section08.htm#Figure 8.6.2.1-1
http://ston.jsc.nasa.gov/collections/trs/_techrep/TM-2011-217352.pdf
http://ston.jsc.nasa.gov/collections/trs/_techrep/TM-2015-218564.pdf
Assuming 18 people can use the same communal space as 6, then 15m^3 looks possible.
5 m^3 is the very low end of the range for missions about 3 months.
But more worryingly there is a sense in the 2 reports that this is just a placeholder as there has simply been
been no research in this area to find out what the minimum really is.
I suspect the average hipster is already a gibbering mess. Putting on in a spaceship will just make him/her gibberouser.
I never had to hot bunk on the sub I was on, but some of the crew did. This was in the mid to late 1970's, in an FBM submarine that was large for its day. I worked and bunked in the missile compartment, the largest space on the boat, so my conditions were better than most. But it's true that even there, after a few weeks of submerged patrol, some people did start to wig out, and there were some major confrontations that developed over the most trivial triggering incidents.It's kind of worrying that would happen despite what I presume was fairly extensive screening for the roles. :(
Sorry, wrong link.Much better.
http://msis.jsc.nasa.gov/sections/section08.htm#Figure 8.6.2.1-1
http://ston.jsc.nasa.gov/collections/trs/_techrep/TM-2011-217352.pdf
http://ston.jsc.nasa.gov/collections/trs/_techrep/TM-2015-218564.pdf
...it should be noted that the last report (2015) is referencing ~5m^3 of person space for missions 912 days in duration. That's an order of magnitude greater duration than the MCT's 3 month trip. Also, everything seems quite hand-wavy as to why /exactly/ that much space is needed.Sorry, wrong link.Much better.
http://msis.jsc.nasa.gov/sections/section08.htm#Figure 8.6.2.1-1
http://ston.jsc.nasa.gov/collections/trs/_techrep/TM-2011-217352.pdf
http://ston.jsc.nasa.gov/collections/trs/_techrep/TM-2015-218564.pdf
The bottom line of those 2 reports seems to be that that they are recommending just the personal space be somewhere between 7-8 m^3.
...
Sorry, wrong link.Much better.
http://msis.jsc.nasa.gov/sections/section08.htm#Figure 8.6.2.1-1
http://ston.jsc.nasa.gov/collections/trs/_techrep/TM-2011-217352.pdf
http://ston.jsc.nasa.gov/collections/trs/_techrep/TM-2015-218564.pdf
The bottom line of those 2 reports seems to be that that they are recommending just the personal space be somewhere between 7-8 m^3.
Unless this MCT comes with a big blow up section that's going to have a pretty sizeable impact on vehicle size.
That's an order of magnitude greater duration than the MCT's 3 month trip. Also, everything seems quite hand-wavy as to why /exactly/ that much space is needed.
You'll have fewer crew for the return journey. Thems the rules or this whole thing doesn't work.
Strong guarantees of a round-trip ticket are necessary to get a million people to pay $500k to go to Mars.
...it should be noted that the last report (2015) is referencing ~5m^3 of person space for missions 912 days in duration. That's an order of magnitude greater duration than the MCT's 3 month trip. Also, everything seems quite hand-wavy as to why /exactly/ that much space is needed.Sorry, wrong link.Much better.
http://msis.jsc.nasa.gov/sections/section08.htm#Figure 8.6.2.1-1
http://ston.jsc.nasa.gov/collections/trs/_techrep/TM-2011-217352.pdf
http://ston.jsc.nasa.gov/collections/trs/_techrep/TM-2015-218564.pdf
The bottom line of those 2 reports seems to be that that they are recommending just the personal space be somewhere between 7-8 m^3.
...
In the 2011 report, most of those problems seem better mitigated by things other than simply having more per-person volume. For instance, sending always two or more MCTs at once (with margin for both) would provide full backup capability without adding extra space required just for backup use. Customizability is something that is also not volume-dependent. Having sufficient mental tasks can be accomplished by providing stimulating training during the transit, including perhaps group projects to prepare for life on Mars. VR headsets could help with mental stimulation, as well as having the entire Netflix catalogue available.
Proper layout of the cabin itself as well as good sleep schedule management (with some freedom to change sleep shift if you want) could help relieve congestion in common areas without needing any more volume per person.
...again, those reports don't at all seem definitive about what exactly is the minimum per-passenger space needed for a short 90-100 day transit.
...and remember also that for the same mass, allocating less volume will improve the radiation protection automatically.
That's huge emphasizing the need to make ECLSS closed loop as much as possible.
90 day transit is insanely expensive in propellant, and while it might be possible to do it from LEO at huge cost in propellant. But the longest leg of the journey is what you need to plan around and that is likely to be the Earth return leg.
Do your remember the discussion we had back on page 39 & 40 when we found that just 100 days Earth return launches from Mars surface would require 8.8 km/s DeltaV and significant aero-braking or propulsive capture at Earth.
Transit times of 120-180 days for Earth return are the range that's actually achievable.These are the figures I was thinking about. Do you happen to have the Delta V to give a 100 day to Mars trip?
If the crew modules are unloaded for surface habitation, then you can use some of the cargo vehicles. If you're talking evacuating the entire planet, that would take longer than a single synod anyway, an impossible standard if you really believe in establishing a city.You'll have fewer crew for the return journey. Thems the rules or this whole thing doesn't work.
Why even bother returning the vehicles? We could just scale up Mars One by a factor of 30?
"But it's necessary for the price to drop to $500k..."
Strong guarantees of a round-trip ticket are necessary to get a million people to pay $500k to go to Mars.
That's huge emphasizing the need to make ECLSS closed loop as much as possible.
90 day transit is insanely expensive in propellant, and while it might be possible to do it from LEO at huge cost in propellant. But the longest leg of the journey is what you need to plan around and that is likely to be the Earth return leg.
Do your remember the discussion we had back on page 39 & 40 when we found that just 100 days Earth return launches from Mars surface would require 8.8 km/s DeltaV and significant aero-braking or propulsive capture at Earth.QuoteTransit times of 120-180 days for Earth return are the range that's actually achievable.These are the figures I was thinking about. Do you happen to have the Delta V to give a 100 day to Mars trip?
That also hint SX might like to make it an uncrewed return, but then how do you get the crew back for the next one?
The bunks will almost certainly be vertical. Say 1.5m in from the wall, 1.5m of wall circumference, and then you can do an inner ring of bunks with similar dimensions. With such an arrangement you can easily fit 30 people in a 12m diameter floor, half that for your 8.3m radius tank.That might have worked on Shuttle but how were you going to handle launch?
Also, the central corridor and doors are wasted space. Doors need space to open and if they open into a corridor can smack an unsuspecting passenger or snag some cargo being pushed along. All you need is a standard door-sized hole to make a corridor.Perhaps but a single depressurization event anywhere in this connected space now threatens everyone.
The bunks will almost certainly be vertical. Say 1.5m in from the wall, 1.5m of wall circumference, and then you can do an inner ring of bunks with similar dimensions. With such an arrangement you can easily fit 30 people in a 12m diameter floor, half that for your 8.3m radius tank.That might have worked on Shuttle but how were you going to handle launch?
QuoteAlso, the central corridor and doors are wasted space. Doors need space to open and if they open into a corridor can smack an unsuspecting passenger or snag some cargo being pushed along. All you need is a standard door-sized hole to make a corridor.Perhaps but a single depressurization event anywhere in this connected space now threatens everyone.
Yes it's probably the number of headaches caused by people being hit by closing compartment doors will be higher than the number of blow outs on any given flight but the consequences if one did happen are much more severe.
BTW
100 people weigh somewhere around 9-10 tonnes. Open loop LSS is 50 tonnes of consumables (hopefully 1/2 that or less with SoA ECLSS closing most of the loops)
But whats the power & thermal situation?
I mean what are current figures of merit for PV arrays and radiators? In 2008 80-100W/Kg for a rigid array was SoA. The same report gave thin film systems around 2000W/Kg It lists ISS as roughly 1W/Kg, a staggeringly low number IMHO.
http://www.spacefuture.com/archiveearly_commercial_demonstration_of_space_solar_power_using_ultra_lightweight_arrays.shtml
But what about radiators.
On Earth a human being has been simulated by a 400W incandescent lamp. So figure 40Kw for humans alone? It's also likely to be low temperature heat, so radiation efficiency is likely to be low.
Does anyone have some actual numbers? IIRC ISS is about 40W/m^2 but that sounds like garbage.
400W? No, more like 100W. 400W is like a champion cyclist going full tilt. 40W/m^2 sounds about right on average for low-temperature ammonia radiator panels. Google is your friend. Each radiator panel can dump a maximum of about 200W/m^2.
After reading the NASA reports/specs for "orbital" modules and reading the posts here, there seems to be a disjoint. The MCT has to ascend and land which means passengers will need a seat. Moreover, the MCT will be on Earth and Mars (in the vertical position) therefore orientation is predetermined. This means that the vertical height will be fixed, most likely 7 feet. The area (not volume) for a crew/sleep station should the size of a bed and storage. The bed would articulate into a seat and workstation (aka cubicle). Think of like a first class seat on a long haul aircraft. Bed at (simulated) night and cubicle by day. Not that different from my life if you minus my commute. :)
Kaoru
The bunks will almost certainly be vertical. Say 1.5m in from the wall, 1.5m of wall circumference, and then you can do an inner ring of bunks with similar dimensions. With such an arrangement you can easily fit 30 people in a 12m diameter floor, half that for your 8.3m radius tank. Also, the central corridor and doors are wasted space. Doors need space to open and if they open into a corridor can smack an unsuspecting passenger or snag some cargo being pushed along. All you need is a standard door-sized hole to make a corridor.Obviously, I thought about making the bunks vertical (a la NASA) but that mode is only good for orbit/space. What about when landed on Mars. You can't move 100 people plus cargo when no or a limited hab is in place. You'll need a place and do a migration over time; that means normally oriented beds.
The seat will need to oriented properly for takeoff and landing, but in microgravity during flight it can be moved to the wall. That will allow the personal space to be more vertical than horizontal. There will need to be enough room for a computer and storage of some personal items along the walls. 1.5m by 1m by 2.25m high totaling 3.375m^3 should be plenty of volume for personal space.You still need a bed in the horizontal position if one of the modes is to use the module as a temporary hab on Mars. It is *great* idea to allow the seat/bed to be reconfigured for more space while on orbit. With an articulating frame you can have a normal (horizontal) seat, horizontal bed (for Mars surface; temporary), and a vertical bed for on orbit. I'll see if I can work that into my module.
Sleeping and working in shifts can easily result in a crewperson spending 50% of their time in their personal space. With that in mind, 15m^3 per person would be about the same as 27m^3 per person if there was only one shift. That's a comfortable volume for a few months in microgravity. Going from 15m^3 down to 12m^3 would still be reasonable.
Inflatable mattress would work on Mars if needed for sleeping. Low mass, small volume when not inflated.The seat will need to oriented properly for takeoff and landing, but in microgravity during flight it can be moved to the wall. That will allow the personal space to be more vertical than horizontal. There will need to be enough room for a computer and storage of some personal items along the walls. 1.5m by 1m by 2.25m high totaling 3.375m^3 should be plenty of volume for personal space.You still need a bed in the horizontal position if one of the modes is to use the module as a temporary hab on Mars. It is *great* idea to allow the seat/bed to be reconfigured for more space while on orbit. With an articulating frame you can have a normal (horizontal) seat, horizontal bed (for Mars surface; temporary), and a vertical bed for on orbit. I'll see if I can work that into my module.
Sleeping and working in shifts can easily result in a crewperson spending 50% of their time in their personal space. With that in mind, 15m^3 per person would be about the same as 27m^3 per person if there was only one shift. That's a comfortable volume for a few months in microgravity. Going from 15m^3 down to 12m^3 would still be reasonable.
Kaoru
I think this is no problem since you will land on Earth and you could easy recover from long voyage. What is important is trip to Mars, when arrived you have to be ready to work hard make colony alive.That's an order of magnitude greater duration than the MCT's 3 month trip. Also, everything seems quite hand-wavy as to why /exactly/ that much space is needed.
90 day transit is insanely expensive in propellant, and while it might be possible to do it from LEO at huge cost in propellant. But the longest leg of the journey is what you need to plan around and that is likely to be the Earth return leg.
Do your remember the discussion we had back on page 39 & 40 when we found that just 100 days Earth return launches from Mars surface would require 8.8 km/s DeltaV and significant aero-braking or propulsive capture at Earth.
Transit times of 120-180 days for Earth return are the range that's actually achievable.
From the comments given on my model, I got some ideas on how to improve space and access. The idea RonM gave me got me thinking on how to have the seat/bed articulate horizontal to vertical. For a display/workstation, I always planned to make it part of the seat/bed with an overhead console that contains air vent/fan (on orbit will always be on; got to move air around), lighting, connections (for pressure suit), and communications. The pressure suit can be store in the overhead console/bin. Think like an airplane. However, this seat/bed with overhead console/storage would be fully integrated together. This means that it will be functional as a horizontal bed, seat, or vertical bed. Using this fold-away notion other equipment can now fold into this space for use.
One possible uses for this space is for exercise. For example, an exercise treadmill will be necessary for everyone to stave off the effects of weightlessness. However, having equipment as such per passenger station or in a common area has pros and cons. I would like hear people opinions on how the space could be utilized (without costing a lot of mass) vis-à-vis common areas like food prep, eating, hygiene, exercise, etc.
Update: I forgot to mention my idea for the doors. Originally I oriented the doors so that I could do a grab rail/lift rail that extends from the top to the bottom. However, I like the idea of having a large central core but having the 8 doors latching onto the 4 rails shrinks the core space. I got a brilliant idea from the Model X, make the doors double hinged (aka a folding door) and automatic (with sensors). This effectively doubles the core space and doubles the elevator/lift area.
Kaoru
2000kcal diet works out to 100Watts average. Conversion efficiencies irrelevant, it all becomes heat. :P
For a radiator, the figure of merit you want is mass per unit area. State of the art experimental radiators can do 1.5-2kg/m^2. I think older heritage designs are more like 6kg/m^2.That's what I was looking for. Indicates propellant and HSF consumables are still the big mass items.
More realistically, 273Kelvin rejection temp (0Celsius) and 100kW of heat and a more conservative (but still aggressive) 2kg/m^2 requires 635kg.I think the Shuttle payload bay door radiators worked around 60-80c
But note that the body of the spacecraft itself emits heat, though it's insulated for reentry.True. I wanted to start with the human thermal load as we know the design is expected to support 100 people.
Which is a good argument for putting the crew quarters inside a cargo bay with doors that can open to space like Shuttle.Maybe, but that would add substantial complexity and significant power requirements.
Forgive my naivety.
In the novel Sundiver (David Brin), they dumped excess heat by converting to electricity,
Forgive my naivety.No
In the novel Sundiver (David Brin), they dumped excess heat by converting to electricity, and using that to power a laser that they simply shone in to space.
Is that even theoretically possible?
Nope! You need a larger radiator because you're rejecting the heat at a lower temperature (all methods of extracting useful energy from heat require some sort of thermal gradient or thermodynamically equivalent).Forgive my naivety.No
In the novel Sundiver (David Brin), they dumped excess heat by converting to electricity, and using that to power a laser that they simply shone in to space.
Is that even theoretically possible?
The classic version of this is the Theremo Electric Modules used in RTGs to turn the decay heat from Plutonium directly into electricity. You're looking at about 8% efficiency.
So you need a smaller radiator, but you still need a pretty big one.
Forgive my naivety.
In the novel Sundiver (David Brin), they dumped excess heat by converting to electricity, and using that to power a laser that they simply shone in to space.
Is that even theoretically possible?
Here's another quick grab of improvements to my passenger model. The biggest change is the diameter which is now 10.1 meters, the same as the Saturn V first stage. With this diameter, each deck holds 24 sleep stations and 24 seats (fold-away). That means only 4 decks for 96 passengers. The deck features an airlock door. Being a work in progress, still got fill out the sleep stations and put in a fold-away ladder for up and down on surface access.
Kaoru
The 100-person MCT will only be operational much later in the 21st century.And you know this how exactly? Musk seems quite clear on getting this up and running fairly soon.
By then, Mars will have landing pads and domes / underground tunnels / modules/ whatever to handle an influx of colonists. Many of the initial colonists will earn their money by continually building colony habitats for the newcomers.Quite possibly.
Now of course if you want to design for an artificial gravity environment during the trip, then that's going to change things and naval designs may be the better option (ie your current layout).It's a fair point that Musk does not not seem to have mentioned AG, although it would simplify some problems and effectively give everyone a workout without exercise machines.
The 100-person MCT will only be operational much later in the 21st century.And you know this how exactly? Musk seems quite clear on getting this up and running fairly soon.Quote
Just because we COULD do AG doesn't mean we SHOULD. If your trip is short, you don't need it and probably shouldn't use it.
Unless you are designing for extended surface operations, you don't need the under-bed drawers, and you don't need the "dressing space" between the beds and the door. Designing for zero-g allows you to cheat by using a lot less space than normal. People don't stretch out in bed in zero gee - they float in foetal position, meaning your sleeping compartments are that much smaller. They can also go anywhere - wall, floor, ceiling. Lockers too can go anywhere. You don't need "floor" space where you can have a dual usage area - vertical corridors don't need landings, etc. Since your walls are all just partitions, you can just take them down and stow them for atmospheric entry and landing.I agree with you if the passenger module was going to be on orbit exclusively. However, the modes that the passenger module will be employed is for both on orbit and on surface. My take is that it will take lots of man power to build the first habitat on Mars. This means the passenger module, as part of the Mars landing, will be the habitat for some time. When on orbit, orientation or areas as you describe are meaningless. However, on the surface it does matter so my layout inherently supports both on orbit and on surface modes.
The 100-person MCT will only be operational much later in the 21st century. By then, Mars will have landing pads and domes / underground tunnels / modules/ whatever to handle an influx of colonists. Many of the initial colonists will earn their money by continually building colony habitats for the newcomers.
Now of course if you want to design for an artificial gravity environment during the trip, then that's going to change things and naval designs may be the better option (ie your current layout).
Here's another quick grab of improvements to my passenger model. The biggest change is the diameter which is now 10.1 meters, the same as the Saturn V first stage. With this diameter, each deck holds 24 sleep stations and 24 seats (fold-away). That means only 4 decks for 96 passengers. The deck features an airlock door. Being a work in progress, still got fill out the sleep stations and put in a fold-away ladder for up and down on surface access.
Kaoru
Just because we COULD do AG doesn't mean we SHOULD. If your trip is short, you don't need it and probably shouldn't use it.Most people would not consider 3-4 months a short trip.
Just because we COULD do AG doesn't mean we SHOULD. If your trip is short, you don't need it and probably shouldn't use it.Most people would not consider 3-4 months a short trip.
However the practical work on AG is so limited it would absurd to bet your plans on it working.
If space exploration were done logically it's a technique that would have been tried decades ago and its benefits and problems already found. Yet in 2016 it still has had no full scale test. :(
Zero-g certainly allows a much more compact ship than would otherwise be the case.
How was the electricity produced? Conversion of energy produces heat.
Just because we COULD do AG doesn't mean we SHOULD. If your trip is short, you don't need it and probably shouldn't use it.Most people would not consider 3-4 months a short trip.
However the practical work on AG is so limited it would absurd to bet your plans on it working.
If space exploration were done logically it's a technique that would have been tried decades ago and its benefits and problems already found. Yet in 2016 it still has had no full scale test. :(
The cold pack dumps its heat to the surrounding air. The heat is still present, it's just been moved, namely from the pack and to the surrounding air. It's why the backs of fridges are hot. Because the fridge is merely dumping the heat, not unmaking it.
"Endothermic" does not mean "laughs in the face of entropy", rather it states that a certain reaction requires heat to happen and it most certainly does not remove the heat, merely moves it around.
a bolo made of two [[my edit]MCTsBFSs] connected by tether at the nose. Spinup and spindown by synchronized use of RCS thrusters.
ThermocoupleForgive my naivety.
In the novel Sundiver (David Brin), they dumped excess heat by converting to electricity,
How was the electricity produced? Conversion of energy produces heat.
the commode weighs 50Kg. ::) A 110lb toilet. I get it, gas,water and solid separation is complex. But 50Kg. Really?
I know what you're talking about WRT endothermic reactions. These do not violate the laws of thermodynamics because entropy still increases, and you soon run out of the little reservoir of low entropy.The cold pack dumps its heat to the surrounding air. The heat is still present, it's just been moved, namely from the pack and to the surrounding air. It's why the backs of fridges are hot. Because the fridge is merely dumping the heat, not unmaking it.
"Endothermic" does not mean "laughs in the face of entropy", rather it states that a certain reaction requires heat to happen and it most certainly does not remove the heat, merely moves it around.
You are simply wrong. The heat pump on a refrigerator does move the heat, but that is not what happens in an endothermic reaction. I used to teach both chemistry and physics, as well as biology. When carbon oxidizes, potential energy is given off as heat. In an endothermic reaction, ambient heat is absorbed and changed into potential energy in the new chemical bonds that are created. The energy changes from thermal energy to potential energy. A cold pack does not dump heat; it soaks it up. You are absolutely and totally W...R...O...N...G ! ! ! Go look it up. Exo thermic means puts out heat. Endo thermic literally means that heat goes into the reaction. Quit arguing and go do the research.
I know what you're talking about WRT endothermic reactions. These do not violate the laws of thermodynamics because entropy still increases, and you soon run out of the little reservoir of low entropy.
This other device you're talking about, a device that's able to generate useful energy by absorbing waste heat indefinitely would clearly NOT obey the laws of thermodynamics.
Please realize that this device that you keep defending is not possible.
I agree that AG should not be in the critical path. With that said...Zero-g certainly allows a much more compact ship than would otherwise be the case.
Please convince me of this.
Thermocouple
Thermopile
Radioisotope_thermoelectric_generator
Stirling_engine
May not be mass and or volume efficient inside the MCT. Radiators might be better.
The electrical energy produced could then be used to help run the heat pump ( A/C ). I don't know if this could generate more or less power than what is needed to run the heat pump ( A/C ).
Please realize that this device that you keep defending is not possible.I have not defended any device. I stated above that I do not know the methodology by which it supposedly works, however, such a device is theoretically possible.
This is dumb. Can we get back on topic instead of giving a platform to people (who ought to understand thermodynamics) to demonstrate that they don't understand thermodynamics?
You can't make your radiator smaller by trying to get free energy. Argue about it in the New Physics section if you like.
I agree that AG should not be in the critical path. With that said...Zero-g certainly allows a much more compact ship than would otherwise be the case.
Please convince me of this.
Essentially, the difference between playing Tetris and playing Tetris without being able to rotate the bricks.
Here's another quick grab of improvements to my passenger model. The biggest change is the diameter which is now 10.1 meters, the same as the Saturn V first stage. With this diameter, each deck holds 24 sleep stations and 24 seats (fold-away). That means only 4 decks for 96 passengers. The deck features an airlock door. Being a work in progress, still got fill out the sleep stations and put in a fold-away ladder for up and down on surface access.
Kaoru
I'd like to say a few words about laundry.You can vacuum dry most things, at low pressure the water will boil away from its own heat. You can put the water back into the eclss. Cleaning a sleeping bag may be a chore though.
Assuming people wear their clothes for a week each that's a minimum of 1500 sets of clothes you'll be taking to Mars unless something gets done about reusing them.
Remember this is not the ISS, where stocks can be substantial and resupply is (relatively) easy. :(
Before we get to all the tricky techno solutions (BTW still waiting for my waterless ultrasonic clothes cleaner) could I suggest a simple, clear heavish grade plastic bag you stick your stuff in with a waterproof closure, hook up to water & soap connectors then knead the water & soap in to hand wash.
Personally I think it could be quite theraputic. :)
Now the rinsing and drying stuff looks a bit harder. A waste connector to flush the water then transfer to a warm dry air cabinet to gradually suck the water off the fabric?
Naturally this all has to be plumbed into the ECLSS to to work.
This other device you're talking about, a device that's able to generate useful energy by absorbing waste heat indefinitely would clearly NOT obey the laws of thermodynamics.
Please realize that this device that you keep defending is not possible.
No, it's not dumb. Because most people don't understand the laws of thermodynamics. But should be elsewhere, certainly.
So you're talking about increased packing efficiencies from being able to store things in the ceiling/floor racks?Essentially, the difference between playing Tetris and playing Tetris without being able to rotate the bricks.Zero-g certainly allows a much more compact ship than would otherwise be the case.Please convince me of this.
Before we get to all the tricky techno solutions (BTW still waiting for my waterless ultrasonic clothes cleaner) could I suggest a simple, clear heavish grade plastic bag you stick your stuff in with a waterproof closure, hook up to water & soap connectors then knead the water & soap in to hand wash.
Now the rinsing and drying stuff looks a bit harder. A waste connector to flush the water then transfer to a warm dry air cabinet to gradually suck the water off the fabric?
{snip for length }We are considing the possible design for removing excess heat from the MCT , that is part of the MCT's design.
And I did say help in powering the heat pump, meaning it would not be able to supply all the needed power. The rest of the power would need to come from the MCT's power source ( solar panels ).
Jim in an above post I did stay a Stirling engines and not heat heat pump for producing power from the waste heat, the heat pump just moves heat from one point to another. For the tech we have today I agree with you that the heat pump would most likely produce more heat than the Stirling engine would convert to electrical power. That making this not work on a space craft meant to land on a planet. But that does not mean sometime in the future it will not be possible to have such a system be able to work on a craft such as the MCT concept. People are working on the concept, it may or may not ever work.
And I did say help in powering the heat pump, meaning it would not be able to supply all the needed power. The rest of the power would need to come from the MCT's power source ( solar panels ).
The only way of removing heat from the MCT is by radiators or dumping mass. There is no way around it. A "heat pump" by definition moves heat from a cold region to a hot region and it consumes energy in the process which generates more heat. A heat pump does not use direct heat for power.
1. Jim in an above post I did stay a Stirling engines and not heat heat pump for producing power from the waste heat,
2. the heat pump just moves heat from one point to another. For the tech we have today I agree with you that the heat pump would most likely produce more heat than the Stirling engine would convert to electrical power. That making this not work on a space craft meant to land on a planet. But that does not mean sometime in the future it will not be possible to have such a system be able to work on a craft such as the MCT concept. People are working on the concept, it may or may not ever work.
And some heat pumps do use heat for power. Look up motor home refrigerators and how they use heat from propane to work
True. Mostly that the bag solution already exists, which I'd forgotten when I originally posted. Something like it is sold as a gadget for backpackers and lightweight travelers.Before we get to all the tricky techno solutions (BTW still waiting for my waterless ultrasonic clothes cleaner) could I suggest a simple, clear heavish grade plastic bag you stick your stuff in with a waterproof closure, hook up to water & soap connectors then knead the water & soap in to hand wash.
Now the rinsing and drying stuff looks a bit harder. A waste connector to flush the water then transfer to a warm dry air cabinet to gradually suck the water off the fabric?
There's a big gap between a sci-fi waterless ultrasonic cleaner and "stick it in a bag with some soapy water".
A water efficient combination washer/dryer seems pretty trivial compared to everything else that'll need to work. It doesn't need to be as fast or heavy as anything we use in the home. Hell, if you must, it can even be spun by hand (or foot (http://1.bp.blogspot.com/-DnJHnHn3X24/VXsTw5BQOII/AAAAAAAAEhM/vsIyWM4jLS0/s1600/Maquina%2Bde%2Blavar%2Bbicicleta.jpg))Well there was something in the 2011 ISS budget for this but it never happened. :(
And I did say help in powering the heat pump, meaning it would not be able to supply all the needed power. The rest of the power would need to come from the MCT's power source ( solar panels ).
The only way of removing heat from the MCT is by radiators or dumping mass. There is no way around it. A "heat pump" by definition moves heat from a cold region to a hot region and it consumes energy in the process which generates more heat. A heat pump does not use direct heat for power.
You can vacuum dry most things, at low pressure the water will boil away from its own heat. You can put the water back into the eclss. Cleaning a sleeping bag may be a chore though.Agreed. The challenge is to retain the water for reuse.
There are methods which convert heat directly to electricity (though with low efficiencies so far).
There are methods than convert heat DIFFERENCE directly into electricity. These are called heat pumps and are disscussed above.
There are methods which convert heat directly to electricity (though with low efficiencies so far).
No, there are NOT. There are methods than convert heat DIFFERENCE directly into electricity.
In order to have that considerable heat DIFFERENCE, big radiators are needed.
It would seem we shed 10grams of skin per day, for a total of 3.6 kg per year. So 100 crew would produce 1 kg per day of skin flakes, and for a 200 day trip 200 kg.You can vacuum dry most things, at low pressure the water will boil away from its own heat. You can put the water back into the eclss. Cleaning a sleeping bag may be a chore though.Agreed. The challenge is to retain the water for reuse.
Bizarre as it may seem I think the simplest option may be to use some vacuum accumulator tanks. Vent them to vac for use, then use them to collect the vapor for later reuse on a pulse basis.
Sizewise I guess the sleeping bag would be the challenge for any system. Obviously zero g helps a lot in this but since I presume people will be staying in the ship on Mars and hence touching the surface a lot during sleep it will need to be washed more frequently.
Consider how long you'd be comfortable sleeping in the same bed linen.
There are methods than convert heat DIFFERENCE directly into electricity. These are called heat pumps and are disscussed above.
There are methods which convert heat directly to electricity (though with low efficiencies so far).
No, there are NOT. There are methods than convert heat DIFFERENCE directly into electricity.
In order to have that considerable heat DIFFERENCE, big radiators are needed.
Less discussed are direct methods like IR photovoltaic cells which dont use heat difference and don't require radiators themselves.
There are methods than convert heat DIFFERENCE directly into electricity. These are called heat pumps and are disscussed above.
There are methods which convert heat directly to electricity (though with low efficiencies so far).
No, there are NOT. There are methods than convert heat DIFFERENCE directly into electricity.
In order to have that considerable heat DIFFERENCE, big radiators are needed.
Less discussed are direct methods like IR photovoltaic cells which dont use heat difference and don't require radiators themselves.
Nope. As previously mentioned, there aren't. Sorry if you misunderstood, but this is not the place for thermodynamic engineering 101; I would start with the Wikipedia entry on Carnot heat engines and work your way up from there. It may help to consider how IR photovoltaic cells would work at an equilibrium temperature, or consider where the energy is supposed to go after you 'generate' it with magical IR photovoltaic cells that do work (spent inside, it heats up the interior). Please seek other resources than this thread.
8)There are methods than convert heat DIFFERENCE directly into electricity. These are called heat pumps and are disscussed above.
There are methods which convert heat directly to electricity (though with low efficiencies so far).
No, there are NOT. There are methods than convert heat DIFFERENCE directly into electricity.
In order to have that considerable heat DIFFERENCE, big radiators are needed.
Less discussed are direct methods like IR photovoltaic cells which dont use heat difference and don't require radiators themselves.
Nope. As previously mentioned, there aren't. Sorry if you misunderstood, but this is not the place for thermodynamic engineering 101; I would start with the Wikipedia entry on Carnot heat engines and work your way up from there. It may help to consider how IR photovoltaic cells would work at an equilibrium temperature, or consider where the energy is supposed to go after you 'generate' it with magical IR photovoltaic cells that do work (spent inside, it heats up the interior). Please seek other resources than this thread.
Well , perhaps just a bit of thermo: The Carnot equation is Work=Heat*(1-Tcold/Thot)
So for a ship wall at 20C, or 293K, and exchanging heat with an exterior at 200K, the maximum work that could be done (or rate of energy extraction and transformation) is 1-200/293=32% so the rest of the energy, 68% would leave the ship at external radiators. This is for an ideal Carnot cycle. Real cycles are about 50% of these efficiences or less. so we might extract 15%. Of course the radiators themselves would exchange heat with deep space according to the Stephan boltzman law of radiation, and would need to be very large since the temperature difference with space (200K to 180 K) would be very small. In fact these radiators would be much larger than if we had just radiated away all the heat in the first place. How much larger? Well radiateor surface is to the fourth of the temperature difference, so 293 to 180 vs 200 to 180 is 5 times more, so 5^4 is 625 times.
So we would get 15% back, at the cost of radiators 625 times larger.
Not worth it, is it?
Forgive my naivety.
In the novel Sundiver (David Brin), they dumped excess heat by converting to electricity, and using that to power a laser that they simply shone in to space.
Is that even theoretically possible?
Nope because the electrical side gets hot and that needs its own set of radiators. Basically you make more waste heat than you throw overboard.
I read a paper ages ago which described a "laser radiator" which basically a very complex gasdynamic laser. Essentially it produced a beam of light with a blackbody radiation curve which carried the waste heat away (so not really a radiator which shoots heat away as a laser beam).
It would seem we shed 10grams of skin per day, for a total of 3.6 kg per year. So 100 crew would produce 1 kg per day of skin flakes, and for a 200 day trip 200 kg.Keep in mind the ISS has been in continuous occupation for 1-2 decades and this is not a show stopper.
So basically at the end of the trip you have two extra crew members made of dead skin!!!! Plus at least the same amount of hair, so you really need to clean the ship and wash the linen. I guess the skin mites will get a lot of it and transform it into CO2 and skin mite shit. Beaurk.
Most of the skin will wash off in showers though. So bacteria will get it instead?
And there is all that urea from sweat that condenses in the linen as well. Definitively need washing machines ;-)I kind of like the washing bags idea because
Explain how heat can conduct or radiate without a temperature differential.Forgive my naivety.
In the novel Sundiver (David Brin), they dumped excess heat by converting to electricity, and using that to power a laser that they simply shone in to space.
Is that even theoretically possible?
Nope because the electrical side gets hot and that needs its own set of radiators. Basically you make more waste heat than you throw overboard.
I read a paper ages ago which described a "laser radiator" which basically a very complex gasdynamic laser. Essentially it produced a beam of light with a blackbody radiation curve which carried the waste heat away (so not really a radiator which shoots heat away as a laser beam).
Actually, it can be done, but some claim it's a violation of the 2nd law. However, in reality, it is not. An electroentropic or magnetoentropic device can do it. It generates electricity without using a temperature differential by directly converting heat into electricity.
Weird thought here;A big scaled up Methane Falcon heavy type design was originally the thought of BFR according to Musk in the R*ddit AMA, but in the AMA, Musk said they opted to instead go the single-core route with LOTS of Raptors.
Anybody ever consider that the BFR could be an up scaled Falcon 9 configuration, using 9 Raptor engines for the first stage and possibly a couple of BFR's as strap on boosters? Or even a configuration using six or more Falcon 9 first stages as strap ons as an alternative?
It seems, based on dV requirements, that BFR will likely need to land on a drone ship. Would the existing drone ships be big enough? If not, how big would a new drone ship need to be?What delta-V requirements? Are you thinking BFR is going to launch all the way to Mars? BFR would likely just be launching to LEO. No, return to launch site is much more likely. This way you can have much faster turnaround time of the BFR stage.
It seems, based on dV requirements, that BFR will likely need to land on a drone ship.
Would putting the linens out the airlock and in full sunlight be the ultimate "dry cleaning?" I have no idea if this would work or not, but i love the idea of the MCT strung with clothes lines and festooned with dirty laundry.Interesting idea. Something that should be testable on the ISS.
Matthew
It seems, based on dV requirements, that BFR will likely need to land on a drone ship. Would the existing drone ships be big enough? If not, how big would a new drone ship need to be?What delta-V requirements? Are you thinking BFR is going to launch all the way to Mars? BFR would likely just be launching to LEO. No, return to launch site is much more likely. This way you can have much faster turnaround time of the BFR stage.
According to Musk, Falcon 9 1st stage separation velicity drops from 9 km/s with drone landing to 6km/s with dry landing.
Indeed.According to Musk, Falcon 9 1st stage separation velicity drops from 9 km/s with drone landing to 6km/s with dry landing.
Source? That seems incredibly unlikely. 9km/s at staging is enough to carry the fully fuelled ~100 tonne second stage into LEO without firing its engines. That would make the F9 first stage a 110 tonne to LEO SSTO!
(Sorry, a 110 tonne to LEO reusable SSTO.)
AIUI people have said a 100 day trip needs 8.8Km/s from LEO. Raptors Isp is suggested at 363secs in Vac, which should be OK in the Mars atmosphere as well, given its pressure.
363 sec. in vac. for SL Raptor. 380 sec. in vac. for the vac. Raptor.AIUI people have said a 100 day trip needs 8.8Km/s from LEO. Raptors Isp is suggested at 363secs in Vac, which should be OK in the Mars atmosphere as well, given its pressure.
I seem to recall that Tom Mueller said that they aimed for 380 secs of vac Isp.
363 sec. in vac. for SL Raptor. 380 sec. in vac. for the vac. Raptor.Running with 380sec gives a mass ratio of 10.57:1 and a mass of about 714 tonnes to LEO, IE about 14x the size of the expendable FH.
According to Musk, Falcon 9 1st stage separation velicity drops from 9 km/s with drone landing to 6km/s with dry landing.
Source? That seems incredibly unlikely. 9km/s at staging is enough to carry the fully fuelled ~100 tonne second stage into LEO without firing its engines. That would make the F9 first stage a 110 tonne to LEO SSTO!
(Sorry, a 110 tonne to LEO reusable SSTO.)
It's flying away from the pad in this case at 5000 km/h. In the upcoming flight it'll be going 8 or 9000 km/h, or roughly 5000 mph, in the wrong direction.http://shitelonsays.com/transcript/postlanding-teleconference-with-elon-musk-2015-12-22
Apologies, my fault, the difference was 9000 km/hr down to 6000km/hr. This was based on a Musk tweet on 1/17/2016. Still, I think the dV penalty of dry landing of BFR makes drone ship landing worth consideration.It's not a "penalty" if the first stage is sized to allow for it from the beginning.
Yet another lesson in why we should all be using metric, and we should be using m/s.
Apologies, my fault, the difference was 9000 km/hr down to 6000km/hr. This was based on a Musk tweet on 1/17/2016. Still, I think the dV penalty of dry landing of BFR makes drone ship landing worth consideration.It's not a "penalty" if the first stage is sized to allow for it from the beginning.
200C is actually fine for cotton as long as it's fairly quick. Keep it out there too long and it'd degrade at those temperatures. Anyway, the actual temperature would depend on the color of the cloth and other aspects.Would putting the linens out the airlock and in full sunlight be the ultimate "dry cleaning?" I have no idea if this would work or not, but i love the idea of the MCT strung with clothes lines and festooned with dirty laundry.Interesting idea. Something that should be testable on the ISS.
Matthew
The joker in the pack is not the vacuum. It's the temperature in the lock. IIRC in full sunlight you're talking c200c, in full shadow -200c.
I think cold would do less damage but I'd still suspect it would do so much damage they would literally crumble in your hands.
But note it's still basically an open loop system and the goal should always be to close the loops.
Not right. More like 5-7km/s from LEO if you aerocapture at the other end. 8.8km/s would be doing a propulsive capture, which simply isn't going to happen. SpaceX will be doing aerocapture for the majority of the capture delta-v.Indeed.According to Musk, Falcon 9 1st stage separation velicity drops from 9 km/s with drone landing to 6km/s with dry landing.
Source? That seems incredibly unlikely. 9km/s at staging is enough to carry the fully fuelled ~100 tonne second stage into LEO without firing its engines. That would make the F9 first stage a 110 tonne to LEO SSTO!
(Sorry, a 110 tonne to LEO reusable SSTO.)
I'd settle for a reusable TSTO.
BTW has anyone mentioned the mass fraction this thing needs?
AIUI people have said a 100 day trip needs 8.8Km/s from LEO. ...
Even if the BFR is relatively a smaller proportion of the stack, it is hard to see cutoff not being at least a little bit down range. So, I would claim you can push the design for a somewhat smaller penalty for dry landing - but not penalty free.
Not right. More like 5-7km/s from LEO if you aerocapture at the other end. 8.8km/s would be doing a propulsive capture, which simply isn't going to happen. SpaceX will be doing aerocapture for the majority of the capture delta-v.Indeed.According to Musk, Falcon 9 1st stage separation velicity drops from 9 km/s with drone landing to 6km/s with dry landing.
Source? That seems incredibly unlikely. 9km/s at staging is enough to carry the fully fuelled ~100 tonne second stage into LEO without firing its engines. That would make the F9 first stage a 110 tonne to LEO SSTO!
(Sorry, a 110 tonne to LEO reusable SSTO.)
I'd settle for a reusable TSTO.
BTW has anyone mentioned the mass fraction this thing needs?
AIUI people have said a 100 day trip needs 8.8Km/s from LEO. ...
This is what happens when people (not saying this is you, john smith) use an online tool or online numbers without reading the fine print or understanding where the numbers come from.
http://trajbrowser.arc.nasa.gov/
Not right. More like 5-7km/s from LEO if you aerocapture at the other end. 8.8km/s would be doing a propulsive capture, which simply isn't going to happen. SpaceX will be doing aerocapture for the majority of the capture delta-v.Indeed.According to Musk, Falcon 9 1st stage separation velicity drops from 9 km/s with drone landing to 6km/s with dry landing.
Source? That seems incredibly unlikely. 9km/s at staging is enough to carry the fully fuelled ~100 tonne second stage into LEO without firing its engines. That would make the F9 first stage a 110 tonne to LEO SSTO!
(Sorry, a 110 tonne to LEO reusable SSTO.)
I'd settle for a reusable TSTO.
BTW has anyone mentioned the mass fraction this thing needs?
AIUI people have said a 100 day trip needs 8.8Km/s from LEO. ...
This is what happens when people (not saying this is you, john smith) use an online tool or online numbers without reading the fine print or understanding where the numbers come from.
http://trajbrowser.arc.nasa.gov/
In a 2-dimensional simplification, we go over the dV's required here:
http://forum.nasaspaceflight.com/index.php?topic=37808.msg1436054#msg1436054
http://forum.nasaspaceflight.com/index.php?topic=37808.msg1436460#msg1436460
0) You are correct, we find only about 3.2+2.2=5.4km/s for LEO to Mars 100d-transit escape burn, before accounting for EDL & without accounting for capture. EDL dV needs seem to be contentious, I add a conservative 2km/s for 7.4km/s total. Note again this is a circular coplanar simplification with launch direct to ecliptic; Reality is likely to be a little bit worse.
1) We have never aerocaptured at Mars before
2) We expect the Mars aerocapture to be especially difficult because the Martian atmosphere's thickness is apparently somewhat variable.
3) The required aerocaptures on a 100d Earth-Mars trip are rather extreme for aerocaptures, probably requiring MAC or some novel means; I have read published suggestions that this is unfeasible in a traditional heatshielded lander, though I cannot verify them. It needs to dissipate 9.3km/s of velocity in one pass to get into a high elliptical, but captured orbit, and more if you're doing a direct entry. A minimum aerocapture from slow Hohmann transfer is more like 1.0km/s.
4) A 100d direct return trip with 'free' aerocapture at Earth is likely to require extreme amounts of dV for the Mars launch + escape burn, the link mentioned 8.8km/s.
When MCT departs for Mars the propellant tanks will be mostly empty, having only enough for Mars EDL + reserve. To keep these from boiling off during Mars transit it will require active cooling. My thought is that perhaps it would be better to store these is separate tanks. A tank in a tank? The point being that the inner tank would have additional insulation to reduce the active cooling energy required. My first thought was that the tank in a tank could be a central column inside methane and LOX tanks. Another thought was that these could be large insulated pressure vessels in a lower equipment bay. After landing these tanks could be removed and used as habitation module: large insulated pressure vessels.
Has this been suggested already? Is it a reasonable idea?
Thanks
Long EZ builder / pilot
Not right. More like 5-7km/s from LEO if you aerocapture at the other end. 8.8km/s would be doing a propulsive capture, which simply isn't going to happen. SpaceX will be doing aerocapture for the majority of the capture delta-v.Indeed.According to Musk, Falcon 9 1st stage separation velicity drops from 9 km/s with drone landing to 6km/s with dry landing.
Source? That seems incredibly unlikely. 9km/s at staging is enough to carry the fully fuelled ~100 tonne second stage into LEO without firing its engines. That would make the F9 first stage a 110 tonne to LEO SSTO!
(Sorry, a 110 tonne to LEO reusable SSTO.)
I'd settle for a reusable TSTO.
BTW has anyone mentioned the mass fraction this thing needs?
AIUI people have said a 100 day trip needs 8.8Km/s from LEO. ...
This is what happens when people (not saying this is you, john smith) use an online tool or online numbers without reading the fine print or understanding where the numbers come from.
http://trajbrowser.arc.nasa.gov/
In a 2-dimensional simplification, we go over the dV's required here:
http://forum.nasaspaceflight.com/index.php?topic=37808.msg1436054#msg1436054
http://forum.nasaspaceflight.com/index.php?topic=37808.msg1436460#msg1436460
0) You are correct, we find only about 3.2+2.2=5.4km/s for LEO to Mars 100d-transit escape burn, before accounting for EDL & without accounting for capture. EDL dV needs seem to be contentious, I add a conservative 2km/s for 7.4km/s total. Note again this is a circular coplanar simplification with launch direct to ecliptic; Reality is likely to be a little bit worse.
1) We have never aerocaptured at Mars before
2) We expect the Mars aerocapture to be especially difficult because the Martian atmosphere's thickness is apparently somewhat variable.
3) The required aerocaptures on a 100d Earth-Mars trip are rather extreme for aerocaptures, probably requiring MAC or some novel means; I have read published suggestions that this is unfeasible in a traditional heatshielded lander, though I cannot verify them. It needs to dissipate 9.3km/s of velocity in one pass to get into a high elliptical, but captured orbit, and more if you're doing a direct entry. A minimum aerocapture from slow Hohmann transfer is more like 1.0km/s.
4) A 100d direct return trip with 'free' aerocapture at Earth is likely to require extreme amounts of dV for the Mars launch + escape burn, the link mentioned 8.8km/s.
umm. Are you calculating delta-v by calculationg sums of partial delta-v:s from a delta-v map even when you can do the whole thing in one burn?
In those cases Oeberth effect gives huge benefit and the required delta-v is only the Pythagoran sum of partial delta-v's that are done together, not direct sum of those.
3.2km/s + 2.2 km/s done together at pegiree, high speed velocity, costs only sqrt(3.2^2 + 2.2^2) = 3.88 km/s.
When MCT departs for Mars the propellant tanks will be mostly empty, having only enough for Mars EDL + reserve. To keep these from boiling off during Mars transit it will require active cooling. My thought is that perhaps it would be better to store these is separate tanks. A tank in a tank? The point being that the inner tank would have additional insulation to reduce the active cooling energy required. My first thought was that the tank in a tank could be a central column inside methane and LOX tanks. Another thought was that these could be large insulated pressure vessels in a lower equipment bay. After landing these tanks could be removed and used as habitation module: large insulated pressure vessels.My model shown in my previous post addresses this by an Integrated Vehicle Fluids (IVF) system. Though speculation on my part, something like this would have to be implemented. First, the Dracos (RCS) would be converted to CH4/LOX with electric ignition since hydrazine is hard to come by on Mars. This means that pressurized CH4/LOX is required (doing away with helium). The IVF is what would maintain these tanks using the boil off of the main tanks. The IVF would use some of the CH4/LOX boil off to run the Internal Combustion Engine (ICE) that generates electricity and runs compressors to pressurize those tanks and chillers to cool/maintain the fuels in the main tanks. The trick is to build a system that is light (plumbing/compressors can be heavy) and have really good heat radiators.
Has this been suggested already? Is it a reasonable idea?
Thanks
Long EZ builder / pilot
I did not know cotton was good to 200c. Working out it's actual final temperature inside an airlock would be difficult enough but once you're in direct sunlight then you're into the spectral absorbance and emissivity of the fabrics. Has anyone even measured these factors?Quote from: John Smith 19200C is actually fine for cotton as long as it's fairly quick. Keep it out there too long and it'd degrade at those temperatures. Anyway, the actual temperature would depend on the color of the cloth and other aspects.
Interesting idea. Something that should be testable on the ISS.
The joker in the pack is not the vacuum. It's the temperature in the lock. IIRC in full sunlight you're talking c200c, in full shadow -200c.
I think cold would do less damage but I'd still suspect it would do so much damage they would literally crumble in your hands.
But note it's still basically an open loop system and the goal should always be to close the loops.
This is what happens when people (not saying this is you, john smith) use an online tool or online numbers without reading the fine print or understanding where the numbers come from.True.
http://trajbrowser.arc.nasa.gov/
An item is that the colonization policy would be a strict 50/50 male/female ratio. For there and back mission the ratio could be anything but for colonization it would have to be 50/50.
An item is that the colonization policy would be a strict 50/50 male/female ratio. For there and back mission the ratio could be anything but for colonization it would have to be 50/50.
Excuse me but that was not the sort of ratio seen in historical colonization efforts and I really doubt that it would pertain today. What justification do you have for imposing such a ratio? Is sexual orientation going to be a selection criteria for colonization? Is intention to have children going to be a selection criteria? Is fertility? Is philosophical commitment to monogamy going to be a selection criteria?
The discussion about the SES flight showed that subcooling is not compatible with self pressurizing.
Subcooled prop won't work with VaPak style self pressurization but it does not prevent autogenous pressurization where the propellant is pumped by the engine and then some of it gets boiled and ducted back to the prop tank to keep it pressurized.
Correct, and that's basically how it's ALREADY done, except you'd use hot/warm methane/oxygen instead of hot/warm helium for the system we're talking about.Subcooled prop won't work with VaPak style self pressurization but it does not prevent autogenous pressurization where the propellant is pumped by the engine and then some of it gets boiled and ducted back to the prop tank to keep it pressurized.
Thanks for the explanation. Do I understand this correctly? The tanks would be brought to flight pressure with inert gas, probably from ground support equipment. Then in flight the pressure would be maintained using hot propellant from the engines?
A speculation piece re the MCT on SpaceFlightInsider.comNote the kicker at the end
http://www.spaceflightinsider.com/organizations/space-exploration-technologies/spacexs-mars-colonial-transporter-rumors-realities/
Right. Doesn't have to be a lot, should be less than BFR/MCT. Small modular reactors are supposed to be around $5000/kW (much of that site fees, etc), and I expect SpaceX would be working with some other company, so there'd be development cost sharing for the reactor portion. There are some as small as 11MWe, so about the right size.This article
...got to solve that heat transfer problem, though, with Mars' thin atmosphere.
What about getting the permits to launch a nuclear reactor, though?
I was in institution-scale solar for years. There's a bunch of over thinking the MCT ground use-panels on this thread (plus it's OT, we have a thread for solar panels on Mars in that area).
Instead of complicated schemes to deploy PV panels or schemes like reflectors to get more power out them, it is far easier and cheaper to just have a longer spool of thin film panels to make up for inefficiencies.
Additionally, no automated unfurling system is needed. Just set the spool on the ground and roll it out- over the rocks and everything, then plug it in to the junction box (on board power will be DC). Hammer in some ground stakes every 2m.
A UV coating will be needed. They make that stuff in "space application" strength, so there's no new technology there.
...However, while fine in space, I do have to ask how you intend to structurally support all those PV panels while the MCT is on the martian surface? They'd have to have support arms underneath, or else lay flat on the dirt and roll out like a tongue ...
I was in institution-scale solar for years. There's a bunch of over thinking the MCT ground use-panels on this thread (plus it's OT, we have a thread for solar panels on Mars in that area).
Instead of complicated schemes to deploy PV panels or schemes like reflectors to get more power out them, it is far easier and cheaper to just have a longer spool of thin film panels to make up for inefficiencies.
Additionally, no automated unfurling system is needed. Just set the spool on the ground and roll it out- over the rocks and everything, then plug it in to the junction box (on board power will be DC). Hammer in some ground stakes every 2m.
A UV coating will be needed. They make that stuff in "space application" strength, so there's no new technology there.Quote...However, while fine in space, I do have to ask how you intend to structurally support all those PV panels while the MCT is on the martian surface? They'd have to have support arms underneath, or else lay flat on the dirt and roll out like a tongue ...
I absolutely NAILED this one last year!
http://www.nasa.gov/mission_pages/station/research/experiments/2139.html
Back to BFR/MCT. I believe I have heard about MCT using densified subcooled propellant but may remember wrong. ...MCT would not be volume restricted in the same way (as F9) so it seems to me they will not use subcooling to enable self pressurizing.
'Subcooling' is actually easier on Mars because the sub part refers to temperatures below the nominal boiling point at 1atm. Or boiling points at subatmospheric pressures. On Mars the low pressure of ambient atmosphere means you can drop pressures inside your LV (during prop load) and storage tanks below 1atm without fear of imploding things. Wiki says average surface pressure on Mars is 600Pa. LOX boils at 59K in this pressure, only five Kelvins above freezing point.
'Subcooling' is actually easier on Mars because the sub part refers to temperatures below the nominal boiling point at 1atm. Or boiling points at subatmospheric pressures. On Mars the low pressure of ambient atmosphere means you can drop pressures inside your LV (during prop load) and storage tanks below 1atm without fear of imploding things. Wiki says average surface pressure on Mars is 600Pa. LOX boils at 59K in this pressure, only five Kelvins above freezing point.
That's pretty wishful thinking.What about getting the permits to launch a nuclear reactor, though?
Why would that be difficult? Provided you haven't run it before launch, it's a big lump of metal and ceramic, with some low-level radioactive fuel which is routinely transported around the world. (And the fuel rods/pellets can be launch separately from the reactor itself.) Should be easier than getting permission to launch an RTG.
The MIT report on a potential 6Kw/Kg is interesting. That would make the ISS array weigh about 36Kg. However t there is much more to a full array than just the blanket.
I mean what are current figures of merit for PV arrays and radiators? In 2008 80-100W/Kg for a rigid array was SoA. The same report gave thin film systems around 2000W/Kg It lists ISS as roughly 1W/Kg, a staggeringly low number IMHO.
http://www.spacefuture.com/archiveearly_commercial_demonstration_of_space_solar_power_using_ultra_lightweight_arrays.shtml
'Subcooling' is actually easier on Mars because the sub part refers to temperatures below the nominal boiling point at 1atm. Or boiling points at subatmospheric pressures. On Mars the low pressure of ambient atmosphere means you can drop pressures inside your LV (during prop load) and storage tanks below 1atm without fear of imploding things. Wiki says average surface pressure on Mars is 600Pa. LOX boils at 59K in this pressure, only five Kelvins above freezing point.
I mean what are current figures of merit for PV arrays and radiators? In 2008 80-100W/Kg for a rigid array was SoA. The same report gave thin film systems around 2000W/Kg It lists ISS as roughly 1W/Kg, a staggeringly low number IMHO.We have some distinctions to make:
Raw solar cells can be made as good as 40-100kW/kg.
The actual absorption thickness is on the order of 50-100nm, actually. If we're talking about fundamental limits, here.Actually I was. IIRC KE Drexlers Masters thesis was on the idea of mfg solar sails by vacuum deposition onto a wax layer, then dissolving the wax layer. In principle excellent performance but far too fragile to survive launch.
...think like a solar sail designer!!!
Neat idea, I was thinking of something similar (but I was just thinking regular acid etching... wax is cleverer than what I was thinking). You could launch it with some sort of waxlike material that would sublimate away once heated in space.The actual absorption thickness is on the order of 50-100nm, actually. If we're talking about fundamental limits, here.Actually I was. IIRC KE Drexlers Masters thesis was on the idea of mfg solar sails by vacuum deposition onto a wax layer, then dissolving the wax layer. In principle excellent performance but far too fragile to survive launch.
...think like a solar sail designer!!!
Drexlers thesis should still be available atNeat idea, I was thinking of something similar (but I was just thinking regular acid etching... wax is cleverer than what I was thinking). You could launch it with some sort of waxlike material that would sublimate away once heated in space.The actual absorption thickness is on the order of 50-100nm, actually. If we're talking about fundamental limits, here.Actually I was. IIRC KE Drexlers Masters thesis was on the idea of mfg solar sails by vacuum deposition onto a wax layer, then dissolving the wax layer. In principle excellent performance but far too fragile to survive launch.
...think like a solar sail designer!!!
Drexlers thesis should still be available at Stamford, however it'll date from the late 70's so I've no idea if it's been scanned and available online.Neat idea, I was thinking of something similar (but I was just thinking regular acid etching... wax is cleverer than what I was thinking). You could launch it with some sort of waxlike material that would sublimate away once heated in space.The actual absorption thickness is on the order of 50-100nm, actually. If we're talking about fundamental limits, here.Actually I was. IIRC KE Drexlers Masters thesis was on the idea of mfg solar sails by vacuum deposition onto a wax layer, then dissolving the wax layer. In principle excellent performance but far too fragile to survive launch.
...think like a solar sail designer!!!
Most masters' theses are not published in peer-reviewed journals, today. The overwhelming majority of papers published in the late 70's have not been scanned and made available online and transferred to a paper repository we can access.Excellent work. That's the one I was thinking of.
Fortunately, this is an exception to the rule:
https://dspace.mit.edu/handle/1721.1/16234
https://dspace.mit.edu/bitstream/handle/1721.1/16234/06483741-MIT.pdf?sequence=2
'Subcooling' is actually easier on Mars because the sub part refers to temperatures below the nominal boiling point at 1atm. Or boiling points at subatmospheric pressures. On Mars the low pressure of ambient atmosphere means you can drop pressures inside your LV (during prop load) and storage tanks below 1atm without fear of imploding things. Wiki says average surface pressure on Mars is 600Pa. LOX boils at 59K in this pressure, only five Kelvins above freezing point.
Granted you must think about tank implosion, but also about explosion. The temperatures you cited are open system, but you are not going to put LOX in an open vat and allow it to evaporate/boil away. It is going into closed tanks and will have to be chilled to lower temperature as well as under some pressure in order to fit more density into the tank. Lacking the 1 atm. of pressure on the outside of the tank, the amount of interior pressure the tanks can withstand would be offset by an equal counter-pressure. So we still arrive back at the question of how cold you can get the prop when you are on Mars. You still have the limitations of a much lower electrical power supply than the North American electrical grid and you also have to dump the heat you withdraw from the prop as you chill it. With almost no atmosphere to run through a radiator from a heat pump, the heat has to be dumped into the ground or allowed to radiate directly into space or something. So I am back to wondering whether creating super-cryo prop in situ on Mars is going to be significantly harder than more typical cryo temperature prop.
Most masters' theses are not published in peer-reviewed journals, today. The overwhelming majority of papers published in the late 70's have not been scanned and made available online and transferred to a paper repository we can access.Just had a chance to skim it.
Fortunately, this is an exception to the rule:
https://dspace.mit.edu/handle/1721.1/16234
https://dspace.mit.edu/bitstream/handle/1721.1/16234/06483741-MIT.pdf?sequence=2
Can I get a factcheck / primer on this?
Is 'slush propellant' the same concept as 'subcooled propellant', or is it a subset? Does propellant ever actually see 1atm?
Here is my engineering module (upper stage Raptor engines & tanks) to date. The person at the top shows scale. The airlock/doors are dynamic which mean they actually open/close. The engine canting mount also works, canting the Raptor 15 degrees with the engines retaining gimbal room. The nacelles internally have ablative nozzle extension which also cants with the engine. Next up is landing legs and filling in the engineering spaces with IVF, Draco methlox tankage, plumbing. Also have to finish the upper deck with ECLSS, tankage (water & nitrogen), batteries, etc. The only thing I don't like is the separated tanks; I should have made a common bulkhead/dome. Less weight and height but would have more time to resize everything.
Kaoru
Here is my engineering module (upper stage Raptor engines & tanks) to date. [...]
What use do you envision for the airlock tunnel when the airlock itself is not in use? I mean, it is a lot of space in a space constrained design...
Edit: since it runs by the engines, wouldn't it make sense to allow engine compartment access from the tunnel?
You hit the nail on the head. The space where the engine nacelles and tankage (gaseous methane/oxygen, nitrogen, etc.) are are unpressurized. However, the IVF and other equipment would be pressurized and accessible. The intent is to be able to service (everything that's serviceable) from the inside; important for long term trips where everything is in-situ.
What use do you envision for the airlock tunnel when the airlock itself is not in use? I mean, it is a lot of space in a space constrained design...
Edit: since it runs by the engines, wouldn't it make sense to allow engine compartment access from the tunnel?
The base outer design of the MCT will most likely be something like a jumbo sized Dragon 2 capsule, (assuming the Dragon 2 works as planned) combined with a separate living quarters module and strongback structure to which propulsion, fuel and cargo modules would be attached. Sending a single MCT at a time to Mars, with 100 colonists each, would take many decades to achieve the population intended. Likely, when the actual colonial effort starts, they will effectively be building a "Liner" style system, carrying hundreds of people for each trip.
The base outer design of the MCT will most likely be something like a jumbo sized Dragon 2 capsule, (assuming the Dragon 2 works as planned) combined with a separate living quarters module and strongback structure to which propulsion, fuel and cargo modules would be attached. Sending a single MCT at a time to Mars, with 100 colonists each, would take many decades to achieve the population intended. Likely, when the actual colonial effort starts, they will effectively be building a "Liner" style system, carrying hundreds of people for each trip.
It is quite reasonable, the problem here is Musk saying that they intend to 'land the whole thing'.
Also, the 100 colonists per ship looks more a like long term goal (for future versions of the system) rather than a requirement. A bit like the 80000 persons/year.
He's also said that the design is currently fluid. It's likely to change before they start cutting metal for it.
...Edit: since it runs by the engines, wouldn't it make sense to allow engine compartment access from the tunnel?
{laughs} I really like the idea, but suspect that the amount of maintainable systems you could access would be limited, and number of situations where you have the vehicle safe in space but with an engine failure would be even more limited.
The only scenario that's realistic is on the Martian surface, where they landed safely but either something glitched or they just want to do a check-out and service before lift-off for Earth-return. In which case, you'd be better off designing for access to the entire engine system from outside (as you would on Earth for a first stage after a launch-abort).
Guys, what happened to "Keep It Simple Stupid"?
...
Adding things like internally stored cargo
I am sure you can appreciate maximum access when changing on orbit a couple of Raptors after an unfortunate MMOD event on one of a couple of hundreds of craft you total fleet comprises.
I am sure you can appreciate maximum access when changing on orbit a couple of Raptors after an unfortunate MMOD event on one of a couple of hundreds of craft you total fleet comprises.
You aren't going to be transporting spare Raptors inside the passenger ship, nor moving them through a crew-access tunnel. (Nor even the major parts of the engine.) They'd be stored in one of the cargo ships. So the repair crew still need to exit the ship.
Having a crew access tunnel that opens inside the engine area offers a trivial advantage over an EVA-airlock anywhere else on the ship.
This most likely means a hybrid solution between solar/batteries and an ICE running the pumps/compressors and generator. All of these technologies will be used on orbit and on surface and are absolutely critical/high maintenance. This maintenance requirement assumes access to the technology timely/effectively and in any mode, in other words, in a pressurized environment.
I'm still devoted to finishing my model. I must be since I bought the Pro version of SketchUp to the tune of $700. My justification to my wife was that my daughter (who's doing computer science/game development) can use it for game modelling. She seemed to buy that.
My .02 worth or more,
Kaoru
I'm still on the free Sketchup ;-) Am I missing anything except for the converters?I was on the SketchUp Pro trial which expired and then I used SketchUp Make for a while to see the differences. While on the Pro trial, I implemented dynamic components for the engines, airlocks, and hatches; they actually move when clicked. Also, I used the solid tools (intersect/union, etc.) for implementing a lot of the components. When I went to the free version, those tools were unavailable. I could still model but it was more time consuming/tedious and I couldn't modify the dynamic components I done. I researched around for various modelling software and SketchUp is the only one that was applicable with a reasonable price (perpetual license; no subscription required). All other software was subscription based and way to expensive anyway (aka Autodesk wanted $700 per year).
Did you check the fuel and oxygen requirements for the ICE (Internal Combustion Engine)? for the 100 kW power level I chose for my own design, I calculated about 2 tonnes of solar arrays and converters.
When I checked the power to mass ratio of CH4+3O2=CO2+2H2O in an ICE, I found I would only get 7.1 Mj/kg, or 7100 MJ per tonne. This works out at about 1 tonne per day of combustion required for 40 kW of power + 60 kW of heat, so it was really bad choice. The oxygen is very heavy.
You might choose to pressurize your RCS ullage gas tanks with methane and oxygen and avoid pressurants altogether?
looking forward to seeing the whole model!
An ICE is a really bad idea. Batteries work great. Solar panels also work great. 100 days is much too long to be using a ICE. An ICE throws away 70-80% of your energy. At least a fuel cell only throws away half. And you already have to throw away about half your energy to produce the methane and oxygen in the first place. So you're left with a round-trip efficiency of between 10 and 25%, ie you're left with only one tenth to one fourth of the energy you started with. Bad bad bad. Solar panels, on the other hand, are producing energy and are also improving all the time.
No, because the IVF is lasting for hours, maybe a few days. It'd need like a 50th the fuel of a 100 day trip that we're talking about.An ICE is a really bad idea. Batteries work great. Solar panels also work great. 100 days is much too long to be using a ICE. An ICE throws away 70-80% of your energy. At least a fuel cell only throws away half. And you already have to throw away about half your energy to produce the methane and oxygen in the first place. So you're left with a round-trip efficiency of between 10 and 25%, ie you're left with only one tenth to one fourth of the energy you started with. Bad bad bad. Solar panels, on the other hand, are producing energy and are also improving all the time.
I take it this applies to ACES IVF system as well? AIUI, its powered by a Roush Racing built ICE.
“We can take our first launch, big fuel tanks, supplies, food, water, if it’s a manned mission. And the next mission will bring up the spacecraft or the astronauts in their capsule. With this advanced upper stage, which can fly around for weeks, it’s up there waiting, we can put these pieces together, and outside the deep part of Earth’s gravity well with that much impulse and propellant, we can do anything.
“We can go out and tap the resources that are in space. We can asteroid mine, we can build the infrastructure required for a real and permanent human presence. Fuel depots, water depots, commercial human habitats. This is truly a game-changer and I couldn’t be more excited about what this will do for the future of space, all enabled by that advanced, high-performance, ultralong-duration upper stage,” Bruno said.
That's not what Boeing says.They say weeks, and mission descriptions out as asteroids sound longer than that.
http://spaceflightnow.com/2015/04/16/ula-gets-futuristic/Quote“We can take our first launch, big fuel tanks, supplies, food, water, if it’s a manned mission. And the next mission will bring up the spacecraft or the astronauts in their capsule. With this advanced upper stage, which can fly around for weeks, it’s up there waiting, we can put these pieces together, and outside the deep part of Earth’s gravity well with that much impulse and propellant, we can do anything.
“We can go out and tap the resources that are in space. We can asteroid mine, we can build the infrastructure required for a real and permanent human presence. Fuel depots, water depots, commercial human habitats. This is truly a game-changer and I couldn’t be more excited about what this will do for the future of space, all enabled by that advanced, high-performance, ultralong-duration upper stage,” Bruno said.
An ICE is a really bad idea. Batteries work great. Solar panels also work great. 100 days is much too long to be using a ICE. An ICE throws away 70-80% of your energy. At least a fuel cell only throws away half. And you already have to throw away about half your energy to produce the methane and oxygen in the first place. So you're left with a round-trip efficiency of between 10 and 25%, ie you're left with only one tenth to one fourth of the energy you started with. Bad bad bad. Solar panels, on the other hand, are producing energy and are also improving all the time.
I take it this applies to ACES IVF system as well? AIUI, its powered by a Roush Racing built ICE.
It's hard to imagine Elon Musk taking SpaceX down the internal combustion engine road (i.e., Tesla, Giga-battery factories, Solar City, ZBO Methlox, minimal mass-to-orbit limitations, etc.), but the remaining ACES/IVF concepts are great and almost can be assumed to be part of BFS/MCT.Elon did say that you can make all modes of transportation electric on the exception of rockets. A rocket engine by definition is a combustion engine, adding an internal combustion engine to MCT is no big deal.
It's hard to imagine Elon Musk taking SpaceX down the internal combustion engine road (i.e., Tesla, Giga-battery factories, Solar City, ZBO Methlox, minimal mass-to-orbit limitations, etc.), but the remaining ACES/IVF concepts are great and almost can be assumed to be part of BFS/MCT.Elon did say that you can make all modes of transportation electric on the exception of rockets. A rocket engine by definition is a combustion engine, adding an internal combustion engine to MCT is no big deal.
Kaoru
It's hard to imagine Elon Musk taking SpaceX down the internal combustion engine road (i.e., Tesla, Giga-battery factories, Solar City, ZBO Methlox, minimal mass-to-orbit limitations, etc.), but the remaining ACES/IVF concepts are great and almost can be assumed to be part of BFS/MCT.Elon did say that you can make all modes of transportation electric on the exception of rockets. A rocket engine by definition is a combustion engine, adding an internal combustion engine to MCT is no big deal.
Kaoru
I imagine the advantage of using an ICE is much bigger with LH/LOX than methane/LOX. It is not that hard to get methane/LOX to zero boil off and purely for electric power I believe solar is the better solution, especially beyond LEO.
It's hard to imagine Elon Musk taking SpaceX down the internal combustion engine road (i.e., Tesla, Giga-battery factories, Solar City, ZBO Methlox, minimal mass-to-orbit limitations, etc.), but the remaining ACES/IVF concepts are great and almost can be assumed to be part of BFS/MCT.Elon did say that you can make all modes of transportation electric on the exception of rockets. A rocket engine by definition is a combustion engine, adding an internal combustion engine to MCT is no big deal.
Kaoru
Adding things like internally stored cargo
Where else would you put the cargo for a vehicle capable of EDL on a planet with an atmosphere?
Guys, what happened to "Keep It Simple Stupid"?
...
Not to disagree with the rest of your post, I'd like to suggest that by the time this thing gets built, Musk's and his engineers' concept of Simple will be interesting to behold.
Also, completely aside, I'd note that that phrase, addresses stupid, which people building MCT are not
Here is the latest model of my design speculation. For context, the attached image shows the engineering module focusing on the lower part. Specifically I'm only showing the pressure vessel which would contain (not shown) IVF, Sabatier reactor, batteries, pumps/compressors, etc. which are serviceable. The very lower part would be storage/cargo containing hoses/lines, etc. to hook up the systems externally/on surface or to another module via the CBM ring. Outside the pressure vessel, would contain structure and tankage for various systems, mostly CH4, O2, CO2, etc.. The top deck (which is not done yet) will contain removable water tanks, ECLSS, and other life support systems. Total height as shown is 26.5 meters (87 feet) which obviously does not include the payload (aka crew decks or cargo/consumables/propellants).
Kaoru
Hi! Elon just said this on Twitter, do we think he could be talking below the $500,000 point?
https://twitter.com/elonmusk/status/718598761832968192
"Tickets to orbital hotels, the moon and Mars will be a lot less than people think."
Hi! Elon just said this on Twitter, do we think he could be talking below the $500,000 point?
https://twitter.com/elonmusk/status/718598761832968192
"Tickets to orbital hotels, the moon and Mars will be a lot less than people think."
First time Elon talking about orbital hotels and the moon, to my knowledge.
I don't think the Moon is a necessary step, but I think if you've got a rocket and spacecraft capable of going to Mars, you might as well go to the Moon as well - it's along the way. That's like crossing the English Channel, relative to Mars. So, it's like, if you have these ships that could cross the Atlantic, would you cross the English Channel? Probably. It's definitely not necessary, but you'd probably end up having a Moon base just because, like, why not, ya know.
Hi! Elon just said this on Twitter, do we think he could be talking below the $500,000 point?
https://twitter.com/elonmusk/status/718598761832968192
"Tickets to orbital hotels, the moon and Mars will be a lot less than people think."
First time Elon talking about orbital hotels and the moon, to my knowledge. Might indicate a shift towards more "attainable" goals. As for "what people think". NASA will pay $58m per seat on average to LEO. That's the reality.
Chris, and I'll report my own post so Chris reads it. You seem to know things about SpaceX's Mars plans. Is it true you've kept some of it to yourself or is it all in L2?
I'm confused, what's going on on Mexico?
What do you think?
Kaoru
This is just an over-complicated design IMO. There is no need for a hatch at the bottom like that. And all that around the engines - the exhaust deflector housing - is going to need some *SERIOUS* active cooling to not melt during Raptor burns. Simplify. Remove the bottom hatch/tunnel. Remove the tunnel in the middle of the tanks. Increase tank volume.I value your feedback as it questions my design choices, which in turn, validates the reason why I made those design choices. Of course, my design is still very much a thought exercise until I model it. Until then, my design choices are obviously unclear so I'll make a mental note to show why I decided that way (or clarify the choice).
It looks like you decided on 4 raptors in such an arrangement early on (with engine cowlings - why?) and you don't want to let go of it. That shape doesn't make much sense.
I have a hard time envisioning a Mars base constructed with elements limited in size by having to pass through a tunnel such as the one in your plans.If you look at the dimensions of BEAM, it's not small by any standard (once expanded) but small enough to pass thru my modeled tunnel. My thoughts stem from the requirement in building any habitat, either here on Earth or Mars, is that your materials have to be transported to your job site. Here on Earth we have the luxury of heavy machinery. On Mars, you're not going to have that luxury thus moving building materials is going to be manual labour which in turn means the size/weight will have to be limited. Using BEAM like components is like using IKEA flat packs, some assembly required... :D Since a BEAM like component will fit thru the tunnel with a simple lift/pulley (or magnetic lift) system, you eliminate heavy equipment (like a crane, cargo doors, ramps). With a small motorized dolly, building a habitat from small modules is more practical but time consuming.
The train idea is interesting, and you've obviously quite a attached to it. But I don't think the added complexity and compromises is worth it.Thank you for finding the "space train" idea interesting, but I'm not attached to it. It's only choices which I made that, IMHO, is a possible and plausible MCT architecture.
If an inflatable unit can fit the tunnel, is the lander going to be high enough for it to turn 90 degrees after exiting the spacecraft to get out? If some items can't make the 90 degree turn, they would have to be offloaded from the top via a crane. Then weight and balance can become a problem. Maybe if the fuel is placed higher, payload at the bottom, with people at the top, and they can come down the tunnel.Great insight! I'm glad you mentioned it because I've already put some thought into the two necessary components to handle this. The first component is the tunnel and the lift system which I believe can be made as a electromagnetic rails/lift system with a special platform. This I already mentioned. What I didn't mention is my design for the legs. All the designs I've seen use the same style of SpaceX landing legs. The reality is that they would have to be completely redesigned to work on Mars. First they have to extend and then retract on demand. Secondly, and more germane to your insight, the legs will have to auto-level lifting the weight of the lander (in Earth gravity). Obviously, to aid in the unloading of items on Mars the legs would have the strength to jack up the ship and/or there is a height restriction on the cargo matched to the clearance the legs provide. In either case the legs will require an interesting design due to these requirements. My yet-to-be-modeled legs will be pure speculation but founded in the above realization.
I have a hard time envisioning a Mars base constructed with elements limited in size by having to pass through a tunnel such as the one in your plans.If you look at the dimensions of BEAM, it's not small by any standard (once expanded) but small enough to pass thru my modeled tunnel. My thoughts stem from the requirement in building any habitat, either here on Earth or Mars, is that your materials have to be transported to your job site. Here on Earth we have the luxury of heavy machinery. On Mars, you're not going to have that luxury thus moving building materials is going to be manual labour which in turn means the size/weight will have to be limited. Using BEAM like components is like using IKEA flat packs, some assembly required... :D Since a BEAM like component will fit thru the tunnel with a simple lift/pulley (or magnetic lift) system, you eliminate heavy equipment (like a crane, cargo doors, ramps). With a small motorized dolly, building a habitat from small modules is more practical but time consuming.
Since your going to be on Mars for awhile, time to move and assemble a lot of small modules/components is not an issue... You'll have plenty of time. This is why my model speculates that it (the service module/lander) will be a temporary habitat while on surface.
Kaoru
We have some evidence/conjecture for the notion that BFR is going to be 15m diameter. MCT is going to require lots of space for vacuum bell nozzles to achieve high Isp with enough engines for redundancy, so it will probably be about the same 15m diameter. A side-loading cargo dispenser at 15m overall diameter might have six <=5m diameter cargo pods arrayed along the outer edge, of indefinite length, around a central structural core + crane system. We know that ISS modules have already been designed at 4.1 to 4.5m diameter for 10 to 20 tons of mass, and that MCT is targeting 100 tons 'useful cargo' to the Martian surface. This provides a cargo footprint that is usefully similar to the Shuttle's payloads.I have a hard time envisioning a Mars base constructed with elements limited in size by having to pass through a tunnel such as the one in your plans.If you look at the dimensions of BEAM, it's not small by any standard (once expanded) but small enough to pass thru my modeled tunnel. My thoughts stem from the requirement in building any habitat, either here on Earth or Mars, is that your materials have to be transported to your job site. Here on Earth we have the luxury of heavy machinery. On Mars, you're not going to have that luxury thus moving building materials is going to be manual labour which in turn means the size/weight will have to be limited. Using BEAM like components is like using IKEA flat packs, some assembly required... :D Since a BEAM like component will fit thru the tunnel with a simple lift/pulley (or magnetic lift) system, you eliminate heavy equipment (like a crane, cargo doors, ramps). With a small motorized dolly, building a habitat from small modules is more practical but time consuming.
Since your going to be on Mars for awhile, time to move and assemble a lot of small modules/components is not an issue... You'll have plenty of time. This is why my model speculates that it (the service module/lander) will be a temporary habitat while on surface.
Kaoru
While your right that standardized cargo modules that can be linked together to form habitats, your size is much too small. The BEAM module is a mere 2.36 m in diameter and 1.7 tall when compressed for a volume of ~7.5 m^3. ISS cargo has a fairly low average density so your looking at many dozens of such modules needing to be unloaded from each ship and then linked together. As these modules are already too large to be moved by anything other then cranes their is every incentive to go bigger as the usefulness of the modules increases as well as the speed of unloading.
I favor a module size comparable to a TEU shipping container which is 6.1 m x 2.44 m x 2.59 m totaling 38.5 m^3 roughly 5 times larger then BEAM when compressed and about twice as large as BEAM when expanded. If a module this size expanded with the same ratio as BEAM the resulting interior space would be 80 m^3 a very generous and spacious habitat indeed.
Containers this size will necessitate a side door and a cargo bay of around 500 m^3 in which containers can be stacked and secured by bolting them to structural hard-points as in modern containerized cargo on ships, planes, trains etc etc. This is very similar to the Space Shuttle which had 300 m^3 internal cargo bay and similar hard-point mountings. Cargo would be loaded/unloaded by a gantry crane in the roof of the cargo bay and extends out to clear the edge of the vehicle. First a flat bed truck is unloaded, then modules are unloaded onto the truck which take them away.
Remember we need to think about the entire SYSTEM in MCT including logistics of ground transportation on Mars, we can't just dump stuff right at the landing site, we need the habitat a safe distance of a few miles away.
While your right that standardized cargo modules that can be linked together to form habitats, your size is much too small. The BEAM module is a mere 2.36 m in diameter and 1.7 tall when compressed for a volume of ~7.5 m^3. ISS cargo has a fairly low average density so your looking at many dozens of such modules needing to be unloaded from each ship and then linked together. As these modules are already too large to be moved by anything other then cranes their is every incentive to go bigger as the usefulness of the modules increases as well as the speed of unloading.This is good feedback, though I don't believe my tunnel/small cargo pallets is too small. I sized it to be practical for a spacecraft landing on Mars. While the idea of "cargo containers" are practical here on Earth using airplanes, it's not even remotely practical for mass limited landing spacecraft on Mars. First, such rigid containers represent a lot of dead weight. Add to that the weight of the crane, cargo doors, and the mechanics to support all of that. Of course, you still have the problem of transporting all the cargo to the base site. Obviously, the base site will not be next to the MCT as it has to launch again and you would want to avoid hot exhaust and flying rocks/debris.
I favor a module size comparable to a TEU shipping container which is 6.1 m x 2.44 m x 2.59 m totaling 38.5 m^3 roughly 5 times larger then BEAM when compressed and about twice as large as BEAM when expanded. If a module this size expanded with the same ratio as BEAM the resulting interior space would be 80 m^3 a very generous and spacious habitat indeed.
Containers this size will necessitate a side door and a cargo bay of around 500 m^3 in which containers can be stacked and secured by bolting them to structural hard-points as in modern containerized cargo on ships, planes, trains etc etc. This is very similar to the Space Shuttle which had 300 m^3 internal cargo bay and similar hard-point mountings. Cargo would be loaded/unloaded by a gantry crane in the roof of the cargo bay and extends out to clear the edge of the vehicle. First a flat bed truck is unloaded, then modules are unloaded onto the truck which take them away.
Remember we need to think about the entire SYSTEM in MCT including logistics of ground transportation on Mars, we can't just dump stuff right at the landing site, we need the habitat a safe distance of a few miles away.
Yes, the BFR will most likely be 15 m diameter. However, BFS may not be 15 m diameter because of mass limitations. A 15 m diameter upper stage/BFS is a lot of dry mass, which mean more engines and more fuel.
Yes, the BFR will most likely be 15 m diameter. However, BFS may not be 15 m diameter because of mass limitations. A 15 m diameter upper stage/BFS is a lot of dry mass, which mean more engines and more fuel.
Que? A shorter, wider stage is closer to a sphere, therefore more mass-efficient than a longer narrower stage.
We have some evidence/conjecture for the notion that BFR is going to be 15m diameter. MCT is going to require lots of space for vacuum bell nozzles to achieve high Isp with enough engines for redundancy, so it will probably be about the same 15m diameter. A side-loading cargo dispenser at 15m overall diameter might have six <=5m diameter cargo pods arrayed along the outer edge, of indefinite length, around a central structural core + crane system. We know that ISS modules have already been designed at 4.1 to 4.5m diameter for 10 to 20 tons of mass, and that MCT is targeting 100 tons 'useful cargo' to the Martian surface. This provides a cargo footprint that is usefully similar to the Shuttle's payloads.
Lower these standard cargo pods to the ground, and you can have vehicles drive out of their ends, just like a new automobile might drive out of an ISO container coming off the shipyard stacks in Baltimore. ISO containers are built of the cheapest materials that will take the load of intermodal shipping, and weigh about 10% of their rated maximum load.
It may not be part of the existing design plan, but why couldn't MCT have some temporarily expandable section, a la Bigelow, which could be temporarily expanded during transit to Mars in order to provide necessary interior space during the months-long journey, and which could then be un-expanded or jettisoned before Mars arrival?
It may not be part of the existing design plan, but why couldn't MCT have some temporarily expandable section, a la Bigelow, which could be temporarily expanded during transit to Mars in order to provide necessary interior space during the months-long journey, and which could then be un-expanded or jettisoned before Mars arrival?
Another alternative to using that would be to just launch it separately and have MCT dock with it in LEO before heading out.
Would it then be possible to just leave it in orbit around Mars then pick it up again on the way back? (fully reusable and all that...)
I thought MCT is supposed to fly to Mars directly, with no stops in between (like what Zubrin originally suggested)
I thought MCT is supposed to fly to Mars directly, with no stops in between (like what Zubrin originally suggested)It will probably get refueled in Earth orbit, unlike MD. Otherwise you either have a vehicle that is "mass-starved" for its intended purpose or a launcher that is impractically huge for ground handling.
The Moon MethaLOX ISRU thread got me thinking about the MCT as passenger service vehicle from Earth to Moon surface. For a short trip < 1week the volume needed per passenger is lot less than the volume per passenger needed for a Mars trip. If The MCT is designed for a crew volume area of 2000-3000 m^3 to support the transport of 100 passengers to Mars how many person could be sent on a short trip to just the Moon?
My estimate was numbers of passengers as low as 250 and as high as 750.
2000m^3/8m^3(per person) = 250
3000m^3/4m^3(per person) = 750
BFS?
BFS?
Big Freakin' Spaceship
BFS?
It may not be part of the existing design plan, but why couldn't MCT have some temporarily expandable section, a la Bigelow, which could be temporarily expanded during transit to Mars in order to provide necessary interior space during the months-long journey, and which could then be un-expanded or jettisoned before Mars arrival?Because 1) un-expanding is problematic. Needs to be carefully refolded. 2) jettisoning it makes it expensive.
BFS?
MCT = BFR + BFS.
People previously assumed just the Mars stage was the "MCT", but apparently Musk uses that name for the whole system. IMO, it's less confusing for everyone to adopt Musk's nomenclature, since he's not likely to adopt ours; unfortunately most posters persist in using "MCT" for just the BFS.
They still do, you know? They think the development will cost less because it is only one vehicle design.BFS?
MCT = BFR + BFS.
People previously assumed just the Mars stage was the "MCT", but apparently Musk uses that name for the whole system. IMO, it's less confusing for everyone to adopt Musk's nomenclature, since he's not likely to adopt ours; unfortunately most posters persist in using "MCT" for just the BFS.
I remember back a year or so ago when people argued that Elon's MCT nomenclature implied a single monolithic vehicle which was clearly absurd. Unfortunately most speculation today is hardly much improved with a compulsive desire to make simplistic and monstrously large 'direct' approaches.
Landing the same spacecraft on Mars that was launched from Earth and returning it to Earth or LEO would be "off the wall". Even with LEO refueling before Mars transit.
I think there will be additional "enhancements" in the plan.
Elon says 2025, but in Elon time that means mid 2030s, even his 2:1 schedule slip would be fantastic.
It probably would. If you're going to do refueling, and you're going to be doing ISRU, and you're going to be building a high-performance TSTO fully reusable HLV, then you basically can skip much the rest. No need for a Battlestar Galactica style architecture with 10 different elements to satisfy the 10 different NASA centers, incorporating everyone's pet technology, etc.They still do, you know? They think the development will cost less because it is only one vehicle design.BFS?
MCT = BFR + BFS.
People previously assumed just the Mars stage was the "MCT", but apparently Musk uses that name for the whole system. IMO, it's less confusing for everyone to adopt Musk's nomenclature, since he's not likely to adopt ours; unfortunately most posters persist in using "MCT" for just the BFS.
I remember back a year or so ago when people argued that Elon's MCT nomenclature implied a single monolithic vehicle which was clearly absurd. Unfortunately most speculation today is hardly much improved with a compulsive desire to make simplistic and monstrously large 'direct' approaches.
Another thing that virtually no one here talks about is MARS ORBITAL RENDEZVOUS.
If the BFS can be refueled in LEO then it can likewise be refueled in LMO if propellant can be brought their.
This would break the DeltaV budget for Earth return into two legs, assent to LMO and then TEI resulting in a vastly smaller vehicle AND a faster Earth return then would be possible with a single direct launch of even a huge vehicle.
To make the vehicle capably of departing for Mars it must start at EML-1 fully fueled which again will allow for a faster transit then a large vehicle starting in LEO.
To get the propellants to LMO and the vehicle to EML-1 you use the same solution a SEP tug, fist it moves the BFS to EML-1, picks up and drops off fuel between LEO and EML-1, then makes a fuel run all the way out to mars to rendezvous with the BFS a second time and finally returns to Earth to repeat the cycle.
The BFS would only need around 4-5 km/s DeltaV capability in this scenario which makes it hugely smaller and simpler, to launch to LEO the BFR is a 2 stage rocket like F-9 but with reusable 2nd stage which can carry payloads other then the BFS on top.
Another thing that virtually no one here talks about is MARS ORBITAL RENDEZVOUS.Orbital rendezvous and refueling does make things more mass efficient, but I think ferrying fuel from earth may not be the right approach. Instead consider a tanker variant of the BFS which is sent to Mars and uses local ISRU to loft fuel to LMO for refueling the returning BFS.
If the BFS can be refueled in LEO then it can likewise be refueled in LMO if propellant can be brought their.
This would break the DeltaV budget for Earth return into two legs, assent to LMO and then TEI resulting in a vastly smaller vehicle AND a faster Earth return then would be possible with a single direct launch of even a huge vehicle.
To make the vehicle capably of departing for Mars it must start at EML-1 fully fueled which again will allow for a faster transit then a large vehicle starting in LEO.
To get the propellants to LMO and the vehicle to EML-1 you use the same solution a SEP tug, fist it moves the BFS to EML-1, picks up and drops off fuel between LEO and EML-1, then makes a fuel run all the way out to mars to rendezvous with the BFS a second time and finally returns to Earth to repeat the cycle.
The BFS would only need around 4-5 km/s DeltaV capability in this scenario which makes it hugely smaller and simpler, to launch to LEO the BFR is a 2 stage rocket like F-9 but with reusable 2nd stage which can carry payloads other then the BFS on top.
Another thing that virtually no one here talks about is MARS ORBITAL RENDEZVOUS.Orbital rendezvous and refueling does make things more mass efficient, but I think ferrying fuel from earth may not be the right approach. Instead consider a tanker variant of the BFS which is sent to Mars and uses local ISRU to loft fuel to LMO for refueling the returning BFS.
If the BFS can be refueled in LEO then it can likewise be refueled in LMO if propellant can be brought their.
This would break the DeltaV budget for Earth return into two legs, assent to LMO and then TEI resulting in a vastly smaller vehicle AND a faster Earth return then would be possible with a single direct launch of even a huge vehicle.
To make the vehicle capably of departing for Mars it must start at EML-1 fully fueled which again will allow for a faster transit then a large vehicle starting in LEO.
To get the propellants to LMO and the vehicle to EML-1 you use the same solution a SEP tug, fist it moves the BFS to EML-1, picks up and drops off fuel between LEO and EML-1, then makes a fuel run all the way out to mars to rendezvous with the BFS a second time and finally returns to Earth to repeat the cycle.
The BFS would only need around 4-5 km/s DeltaV capability in this scenario which makes it hugely smaller and simpler, to launch to LEO the BFR is a 2 stage rocket like F-9 but with reusable 2nd stage which can carry payloads other then the BFS on top.
The Battlestar Galactica thing made me think and put some numbers, just plug data in the rocket equation to see what happens. I'm assuming the 'land the whole thing' architecture because, in order to transport a lot of cargo and people to Mars, you need a big and heavy lander anyway, so it does make sense to use the same ship to make the interplanetary part as well. I'm also assuming a pure chemical rocket propulsion system, although G Shotwell mentioned other possibilities being studied.
As a reference I used Saturn V (I'm taking Apollo XVII numbers), obviously the only comparable succesfull system so far, yes the Nova designs where more like it, but they remained paper rockets, no ppt in those happy days.
Plugging the Saturn IC data in the rocket equation shows that deltaV was just over 3.3 Km/s, it also staged at a pretty low altitude (good for the RTLS thing). Let's assume that the MCT booster does likewise, that leaves about 7 Km/s deltaV to the second stage (aka BFS) which is ok if you plan to refuel and do a TMI burn.
As per word of Musk, the ship will be capable of putting 100 t of useful mass on Mars, so let's make it another 100 t for structure (engines, TPS, legs, you name it). So that leaves a 200 t dry mass vehicle, capable of 7 km/s deltaV, with engines giving a Isp of 380 s you need a whopping 1150 t of propellant and a total BFS mass at lift off of about 1350 t. This also pushes the BFR mass to 3200 t, more than a Saturn V GLOW.
Indeed, it's not a small ship, so the term Battlestar does apply here.
The other question is how they will want to size the Raptor. At the currently stated 2300kN a single Raptor would be rather underpowered for throwing 100t into LEO, although probably not worse than the Centaur US. A single 2300 kN engine would be ideal for TMI and Mars return though, and dragging extra engines to Mars and back isn't ideal.
The other question is how they will want to size the Raptor. At the currently stated 2300kN a single Raptor would be rather underpowered for throwing 100t into LEO, although probably not worse than the Centaur US. A single 2300 kN engine would be ideal for TMI and Mars return though, and dragging extra engines to Mars and back isn't ideal.
You're going to need multiple engines to propulsively land BFS back at KSC, Brownsville or wherever, which is absolutely critical for the economics of the architecture. I'm assuming extendable/discardable nozzle extensions on the BFS raptor, which may be a distinctive variant from the initial upper stage raptor vac.
Besides, if I was the commander of a 100 individual interplanetary spaceship, I'd either want engines with a ludicrous reliability or a few with a wide gimbal so I have engine out redundancy. Nobody wants to be marooned.
They are going to be carting a few unnecessary engines with them, I bet the most edible of my cowskin hats on it.
Another thing that virtually no one here talks about is MARS ORBITAL RENDEZVOUS.
If the BFS can be refueled in LEO then it can likewise be refueled in LMO if propellant can be brought their.
This would break the DeltaV budget for Earth return into two legs, assent to LMO and then TEI resulting in a vastly smaller vehicle AND a faster Earth return then would be possible with a single direct launch of even a huge vehicle.
To make the vehicle capably of departing for Mars it must start at EML-1 fully fueled which again will allow for a faster transit then a large vehicle starting in LEO.
To get the propellants to LMO and the vehicle to EML-1 you use the same solution a SEP tug, fist it moves the BFS to EML-1, picks up and drops off fuel between LEO and EML-1, then makes a fuel run all the way out to mars to rendezvous with the BFS a second time and finally returns to Earth to repeat the cycle.
The BFS would only need around 4-5 km/s DeltaV capability in this scenario which makes it hugely smaller and simpler, to launch to LEO the BFR is a 2 stage rocket like F-9 but with reusable 2nd stage which can carry payloads other then the BFS on top.
Questions arise...
How long does the SEP tug take to bring the BFS and/or the propellant to EML-1 and return to LEO?
How many SEP trips to EML-1 per single BFS launch to Mars and return?
How efficient are SEP solar panels after numerous long trips thru the Van Allen belts?
How many BFTanker flights to re-fuel the SEP tug itself for its various transits to support one Mars round trip mission? Including bringing up the BFS propellant that the SEP carries all the way to Mars orbit.
Besides the added expense, resources and complexity of developing the SEP tug.
SpaceX pushes mass fractions to the limit so a 100t empty mass is probably conservative, but lets go with that. However, Saturn V was a 3 stage to orbit system, so the comparisons to a TSTO aren't really accurate.Well, actually I thought 100t estimate to be optimistic!
BFR will likely have mass fractions much closer to Falcon 9 than Saturn V, but with a better ISP than either. Falcon 9 stages at around 2.2 to 2.5 kms, and the upper stage adds about 5.3 kms to get it to LEO. With Raptor's higher ISP, it only takes about 625t of prop to get a 200t ship+payload into LEO.
If they can stage at 3.5 kms (with enough prop in the BFR to boostback to reentry at a survivable Mach 6 or 2 kms), then the upper stage/ship only has to add about 4.5 kms to LEO, which happens to be about the same performance it needs to sent 100t through TMI and do a 1 kms (ish) EDL burn at Mars. That only takes about 500t of prop.
The other question is how they will want to size the Raptor. At the currently stated 2300kN a single Raptor would be rather underpowered for throwing 100t into LEO, although probably not worse than the Centaur US. A single 2300 kN engine would be ideal for TMI and Mars return though, and dragging extra engines to Mars and back isn't ideal.
Besides, if I was the commander of a 100 individual interplanetary spaceship, I'd either want engines with a ludicrous reliability or a few with a wide gimbal so I have engine out redundancy. Nobody wants to be marooned.Mostly agree with that, better safe than sorry. But being Raptors an overkill, I also wonder if they will use Raptors for landing, either on Mars and on Earth. Probably smaller and lightweight pressure fed engines for landing might work too.
They are going to be carting a few unnecessary engines with them, I bet the most edible of my cowskin hats on it.
Another thing that virtually no one here talks about is MARS ORBITAL RENDEZVOUS.
If the BFS can be refueled in LEO then it can likewise be refueled in LMO if propellant can be brought their.
This would break the DeltaV budget for Earth return into two legs, assent to LMO and then TEI resulting in a vastly smaller vehicle AND a faster Earth return then would be possible with a single direct launch of even a huge vehicle.
To make the vehicle capably of departing for Mars it must start at EML-1 fully fueled which again will allow for a faster transit then a large vehicle starting in LEO.
To get the propellants to LMO and the vehicle to EML-1 you use the same solution a SEP tug, fist it moves the BFS to EML-1, picks up and drops off fuel between LEO and EML-1, then makes a fuel run all the way out to mars to rendezvous with the BFS a second time and finally returns to Earth to repeat the cycle.
The BFS would only need around 4-5 km/s DeltaV capability in this scenario which makes it hugely smaller and simpler, to launch to LEO the BFR is a 2 stage rocket like F-9 but with reusable 2nd stage which can carry payloads other then the BFS on top.
Questions arise...
How long does the SEP tug take to bring the BFS and/or the propellant to EML-1 and return to LEO?
How many SEP trips to EML-1 per single BFS launch to Mars and return?
How efficient are SEP solar panels after numerous long trips thru the Van Allen belts?
How many BFTanker flights to re-fuel the SEP tug itself for its various transits to support one Mars round trip mission? Including bringing up the BFS propellant that the SEP carries all the way to Mars orbit.
Besides the added expense, resources and complexity of developing the SEP tug.
For fast transits you want to go nuclear. See for example the DRM 5 Addendum 2. In fact I don't think one synod return is possible without EP or refueling in LMO, because the delta-v is SSTO-level or higher. From both options, EP or LMO refueling, I would pick EP because it also saves you a ton of mass in LEO and it doesn't require a huge launch infrastructure on Mars. Note that you would not refuel the BFS in LMO, the EP would bring it back. In fact all in-space propulsion would be done with EP, however the BFS would do aerocapture at Mars.
If SpaceX plans to use a reactor on Mars they could use it for NEP too, if it has sufficiently low mass.
Plugging the Saturn IC data in the rocket equation shows that deltaV was just over 3.3 Km/s, it also staged at a pretty low altitude (good for the RTLS thing). Let's assume that the MCT booster does likewise, that leaves about 7 Km/s deltaV to the second stage (aka BFS) which is ok if you plan to refuel and do a TMI burn.
As per word of Musk, the ship will be capable of putting 100 t of useful mass on Mars, so let's make it another 100 t for structure (engines, TPS, legs, you name it). So that leaves a 200 t dry mass vehicle, capable of 7 km/s deltaV, with engines giving a Isp of 380 s you need a whopping 1150 t of propellant and a total BFS mass at lift off of about 1350 t.
RB: You completely inverted the meaning of 'Battlestar Galactica'. That was Zubrin's pejorative term for a single huge vehicle that carries all propellant from Earth which was the NASA plan in the 90 day report.
Zubrin's argument was two fold, use insitu propellant on mars to reduce outbound propellant needs (which we all agree Musk is doing) AND using a number of smaller modular elements that serve specific roles and provide redundancy. The Battlestart Galactica term was specifically about not making a single all inclusive budget busting vehicle designed to perform multiple tasks.
The BFR/BFS is only big because SpaceX's cargo requirement is hugely ambitious (100t to Mars surface), it wouldn't be so big if the cargo requirement was reduced to be inline with what Zubrin or NASA imagined.
The BFR/BFS is only big because SpaceX's cargo requirement is hugely ambitious (100t to Mars surface), it wouldn't be so big if the cargo requirement was reduced to be inline with what Zubrin or NASA imagined.
I'd say NASA is currently downsizing its Mars plans, not in terms of payload landed on Mars (still ~80t), but in terms of the size of the individual elements. It's now considering 27t or even 18t landers (instead of 40t), chemical in-space stages with masses around 40t, SEPs with a few hundred kw. Nothing monstrous. In fact the only thing monstrous left is SLS.
All that I suppose in an effort to save cost. SpaceX does exactly the opposite, it's going very big (of course SpaceX attempts to colonize Mars instead of sending 4 people every few years).
That's why I fear that SpaceX will present a paper at IAC but since it's so diametrically opposed to NASA's evolvable Mars campaign it won't have any impact. In fact it might even give congress more legitimation to push forward with SLS (Look! SpaceX wants to build an ever bigger rocket!).
(Note that all Musk has ever said about the Moon is, paraphrasing, "it's on the way, I guess we should look at how to get there, but it's not anything like a goal of the architecture to be able to go to the Moon".)
I, for one, would like to see a new interplanetary transportation architecture be designed to support more than one possible destination, I guess... ;)
Imagine Hubble in the Smithonian.(Note that all Musk has ever said about the Moon is, paraphrasing, "it's on the way, I guess we should look at how to get there, but it's not anything like a goal of the architecture to be able to go to the Moon".)
I, for one, would like to see a new interplanetary transportation architecture be designed to support more than one possible destination, I guess... ;)
Just because going to the Moon is not a goal of the MCT architecture doesn't mean the system won't be capable of being used to go to the Moon. I'm sure people will come up with all sorts of non-Mars uses for the MCT architecture once it's in existence (or even before).
Imagine Hubble keeps operating nearly indefinitely. Much better, IMHO... ;)Imagine Hubble in the Smithonian.(Note that all Musk has ever said about the Moon is, paraphrasing, "it's on the way, I guess we should look at how to get there, but it's not anything like a goal of the architecture to be able to go to the Moon".)
I, for one, would like to see a new interplanetary transportation architecture be designed to support more than one possible destination, I guess... ;)
Just because going to the Moon is not a goal of the MCT architecture doesn't mean the system won't be capable of being used to go to the Moon. I'm sure people will come up with all sorts of non-Mars uses for the MCT architecture once it's in existence (or even before).
Imagine Hubble keeps operating nearly indefinitely. Much better, IMHO... ;)
That's why I fear that SpaceX will present a paper at IAC but since it's so diametrically opposed to NASA's evolvable Mars campaign it won't have any impact. In fact it might even give congress more legitimation to push forward with SLS (Look! SpaceX wants to build an ever bigger rocket!).
Imagine Hubble keeps operating nearly indefinitely. Much better, IMHO... ;)
Nah, we all love it but it's old. Build a one significantly newer and 4x the size. Hell, with that kind of capacity you could have a dedicated manned observatory station, or even throw a hubble-esque telescope into a solar orbit.
The manned portion could float nearby. As far as that issue ruling out that particular use of non-Mars MCT.Imagine Hubble keeps operating nearly indefinitely. Much better, IMHO... ;)
Nah, we all love it but it's old. Build a one significantly newer and 4x the size. Hell, with that kind of capacity you could have a dedicated manned observatory station, or even throw a hubble-esque telescope into a solar orbit.
You wouldn't want a manned observatory station - too much vibration and other interference. Better to have an unmanned observatory which can reached and upgraded quickly and cheaply, as needed. Which is exactly what a "cislunar" BFS could provide.
Will MCT use methalox for all parts of its Earth-Mars trip? If so, then is that volatile methalox supposed to last all the way until the Mars EDL burn? Or will hypergolics be used for the ending burn instead?
Likewise, what about for return trip back to Earth? There likely won't be a way to manufacture traditional hypergolics on Mars for the return trip, so it sounds like Methalox will need to survive from Mars all the way to Earth.
How will that methalox be preserved across the entire journey?
Will MCT use the same very-low-temperature cryo approach as Falcon FT?
It occurred to me that it's *technically feasible* to build BFR and BFS at the cape and have them both self-ferry to Brownsville... in fact this is feasible from any eastern-seaboard location.There is speculation that MCT might be built/tested/flown all at the Cape.
There is no way SpaceX would do that unless the range is ready to support daily flights.
I think this counts as MCT speculation:
While reading this:It occurred to me that it's *technically feasible* to build BFR and BFS at the cape and have them both self-ferry to Brownsville... in fact this is feasible from any eastern-seaboard location.There is speculation that MCT might be built/tested/flown all at the Cape.
There is no way SpaceX would do that unless the range is ready to support daily flights.
Perhaps this side-steps the problem of needing a large skilled workforce in an out-of-the-way place like Brownsville.
I think this counts as MCT speculation:
While reading this:It occurred to me that it's *technically feasible* to build BFR and BFS at the cape and have them both self-ferry to Brownsville... in fact this is feasible from any eastern-seaboard location.There is speculation that MCT might be built/tested/flown all at the Cape.
There is no way SpaceX would do that unless the range is ready to support daily flights.
Perhaps this side-steps the problem of needing a large skilled workforce in an out-of-the-way place like Brownsville.
Shipping large rockets by sea is a problem that has been solved many times. It's not only *technically feasible* but SOP.
(Note that all Musk has ever said about the Moon is, paraphrasing, "it's on the way, I guess we should look at how to get there, but it's not anything like a goal of the architecture to be able to go to the Moon".)
I, for one, would like to see a new interplanetary transportation architecture be designed to support more than one possible destination, I guess... ;)
Regarding the BFS TPS, I always thought the best place to put it was on the bottom.
[...] and it would enable [the] simplist cargo/passenger stowage.
If we put the TPS on top,
I think it could even be inflatable.
(or make a big slow target for a helicopter mid air grab.)
I think it could even be inflatable.
Doubtful. It would need to be repacked on Mars to allow for the Earth-return EDL. Or you'd need to carry two expendable inflatable heat-shields (one for Mars, one for return to Earth.)
(It just doesn't fit the typical SpaceX architecture.)
If acceleration is always in the same direction, then you only need to support cargo from one direction. If not, straps / bolts have to be beefier and more numerous. Not a huge issue. Slightly more problematic is passengers after the flip maneuver and landing. Their high-g chairs need to flip around. I think some kind of swinging hammock would work, but as I said, acceleration in only one direction is simpler to design for.Regarding the BFS TPS, I always thought the best place to put it was on the bottom.What's the reasoning for that?
[...] and it would enable [the] simplist cargo/passenger stowage.
I've assumed if the TPS isn't at the bottom, it'll be down the ventral side. With a biconic or lifting-body forward shape, allowing lift/drag type EDL, extending the deceleration, reducing the amount of supersonic retropropulsion required, and lowering the final speed before the landing burn, both reducing fuel demand.Certainly all valid points. I think everyone agrees that lift is pretty much required to minimize prop.
It also allows a longer glide, improving safety for suboptimal entry timing. (For eg, uneven upper atmosphere means the BFS decelerates earlier than expected, losing speed before reaching the intended landing area. BFS lands short, necessitating a rescue scenario. With a longer glide, you can stretch the descent.)
Doubtful. It would need to be repacked on Mars to allow for the Earth-return EDL. Or you'd need to carry two expendable inflatable heat-shields (one for Mars, one for return to Earth.)
...
(It just doesn't fit the typical SpaceX architecture.)
...
Errr, no. IMO, ULA's engine pod for Vulcan will end up being be too large for helicopter aerial recovery, the BFS would be too large for capture even by a giant transport plane.
what do you think about tripropellant engine for long duration missions?Ain't going to happen. BFR+MCT will be 100% LOx/LCH4. Elon is staying away from H2 because of it's difficulties.
i know that lox+RP+H2 gives some good results but could lox+methane+H2 do that?
for example (engine is the same):
for BFR first stage: start on methane 66-75% + 33-25% H2 for best thrust, then high isp flight on pure H2 and landing on pure methane (no need for high insulation for H2)
for MCT: start to mars on pure H2 and landing on mars on pure methane. start from mars on methane or mix if it possible to produce H2. landing on earth on pure methane.
Matt, that's exactly what a common bulkhead propellent tank design does. Look up the Saturn V second stage tank common bulkhead.
http://www.alternatewars.com/Games/KSP/Tut2/KSP_Tutorial_2-6.htm
What you said is true for hydrolox. It is not true for methalox. It takes 3.4kg of lox to burn 1kg of liquid methane, but lox is only about 2.6 times denser. If you have equal volume tanks you will have methane left over.
If your outer cylinder diameter is 10m, the same-length inner cylinder needs to be 7.5m in diameter to hold enough lox for complete combustion.
Pressure vessel mass to volume ratios don't change much with total volume, and narrower tanks give better airodynamics. I think we will see BFR at around 10 m to 12.5 m in diameter and 90 m tall.Your logic is correct, but your statement isn't correct.
The BFR is likely in my opinion, to have a center engine plus two concentric rings of engines
Pressure vessel mass to volume ratios don't change much with total volume, and narrower tanks give better airodynamics. I think we will see BFR at around 10 m to 12.5 m in diameter and 90 m tall.Your logic is correct, but your statement isn't correct.
Pressure vessel mass to volume ratios don't change much with total volume, and narrower tanks give better airodynamics. I think we will see BFR at around 10 m to 12.5 m in diameter and 90 m tall.
Pressure vessel mass to volume ratios don't change much with total volume, and narrower tanks give better airodynamics. I think we will see BFR at around 10 m to 12.5 m in diameter and 90 m tall.
Efficient rocket stages are pretty much just giant tanks. Think again.Pressure vessel mass to volume ratios don't change much with total volume, and narrower tanks give better airodynamics. I think we will see BFR at around 10 m to 12.5 m in diameter and 90 m tall.
I doubt the tank is going to dominate the structural mass of the vehicle.
Efficient rocket stages are pretty much just giant tanks. Think again.Pressure vessel mass to volume ratios don't change much with total volume, and narrower tanks give better airodynamics. I think we will see BFR at around 10 m to 12.5 m in diameter and 90 m tall.
I doubt the tank is going to dominate the structural mass of the vehicle.
Efficient rocket stages are pretty much just giant tanks. Think again.Pressure vessel mass to volume ratios don't change much with total volume, and narrower tanks give better airodynamics. I think we will see BFR at around 10 m to 12.5 m in diameter and 90 m tall.
I doubt the tank is going to dominate the structural mass of the vehicle.
BFS requires rocket-stage-like performance and so will have to be built like a rocket stage. Tank mass is very important.Efficient rocket stages are pretty much just giant tanks. Think again.Pressure vessel mass to volume ratios don't change much with total volume, and narrower tanks give better airodynamics. I think we will see BFR at around 10 m to 12.5 m in diameter and 90 m tall.
I doubt the tank is going to dominate the structural mass of the vehicle.
I thought he meant BFS, my fault.
Pressure vessel mass to volume ratios don't change much with total volume, and narrower tanks give better airodynamics. I think we will see BFR at around 10 m to 12.5 m in diameter and 90 m tall.
IF the Rvac engine approaches the 5m width some have computed, 10m will be too small.
I figure the BFR/BFS as just under 100m tall but maybe 15m wide, certainly at least 12.5m. If narrower, then taller.
In any case, it will look very different from the pencil necked geek Falcon9.
May I ask if anyone has a conception how MCT is supposed to RETURN from Mars?Neither. Or rather, something in between.
It was mentioned a few times that the return leg, surface to LMO and LMO to TEI, constitutes the most challenging aspect ('long pole') for the entire architecture. Yet I have read almost no discussion of how this could be achieved. Basically, there seem to be two extreme possibilities, plus a range of intermediates.
1) Mars surface is reached by a larger version of the Dragon, with engines suitable for precision landing. But can the same engines take the craft back to orbit?
2) Mars surface is reached by a larger version of the F9R, with engines suitable for both orbital launch and precision landing. But can such a craft launch without a launch complex and its infrastructure? And can such a craft return to earth without replenishing propellants from a tanker in LMO?
Apologies if I have overlooked any information pertaining to these points!
For example, one of the chips I like to use the most in residential downlighting applications, masses 3kg per 100 units. including shipping packaging. Something even lighter, like a Bridgelux COB array comes in at 130-150 lumens per watt, ~1300-2500 lumens delivered, ....
Or for 400kg extra, have a backup solar array that can provide 20-40kW indefinitely. A solar array and a generator have about the same mass, but the solar array isn't going to explode, won't wear out or require maintenance or fluid, and needs no fuel and can operate indefinitely, and doesn't have exhaust.
Just. Stop. Stop trying to make "internal combustion engine in space" happen. It's not happening.
What about in a dust storm on mars? Your solar panels probably aren't going to be much use then.
May I ask if anyone has a conception how MCT is supposed to RETURN from Mars?
It was mentioned a few times that the return leg, surface to LMO and LMO to TEI, constitutes the most challenging aspect ('long pole') for the entire architecture. Yet I have read almost no discussion of how this could be achieved. Basically, there seem to be two extreme possibilities, plus a range of intermediates.
1) Mars surface is reached by a larger version of the Dragon, with engines suitable for precision landing. But can the same engines take the craft back to orbit?
2) Mars surface is reached by a larger version of the F9R, with engines suitable for both orbital launch and precision landing. But can such a craft launch without a launch complex and its infrastructure? And can such a craft return to earth without replenishing propellants from a tanker in LMO?
Apologies if I have overlooked any information pertaining to these points!
If you're landing away from the pad, you're in a survival situation. You're going to be dead very soon, power or no, without a rescue party. You may have even needed to bail out of the vehicle.Or for 400kg extra, have a backup solar array that can provide 20-40kW indefinitely. A solar array and a generator have about the same mass, but the solar array isn't going to explode, won't wear out or require maintenance or fluid, and needs no fuel and can operate indefinitely, and doesn't have exhaust.
Just. Stop. Stop trying to make "internal combustion engine in space" happen. It's not happening.
What about in a dust storm on mars? Your solar panels probably aren't going to be much use then.
What about in a dust storm on mars? Your solar panels probably aren't going to be much use then.
Batteries whilst the storm lasts, shrouds to close over the panels before the storm happens/orient the panels away from the wind/roll the panels back in/put something over them. Any of these objects are less resource intensive than an ICE.
If you're landing away from the pad, you're in a survival situation. You're going to be dead very soon, power or no, without a rescue party. You may have even needed to bail out of the vehicle.Or for 400kg extra, have a backup solar array that can provide 20-40kW indefinitely. A solar array and a generator have about the same mass, but the solar array isn't going to explode, won't wear out or require maintenance or fluid, and needs no fuel and can operate indefinitely, and doesn't have exhaust.
Just. Stop. Stop trying to make "internal combustion engine in space" happen. It's not happening.
What about in a dust storm on mars? Your solar panels probably aren't going to be much use then.
Wasting tons on the corner-case of a corner-case (in a way that still is inferior to other solutions) is not a good idea.
An internal combustion engine on MCT is a bad idea that simply won't die. People keep bringing it back up. It's a zombie idea.
If you're landing away from the pad, you're in a survival situation. You're going to be dead very soon, power or no, without a rescue party. You may have even needed to bail out of the vehicle.You keep saying that. But there are engineering methods to define when something makes sense and when it doesn't. It's called a trade. So help him make the technical trade on mass and simply let the numbers speak for themselves.
Wasting tons on the corner-case of a corner-case (in a way that still is inferior to other solutions) is not a good idea.
An internal combustion engine on MCT is a bad idea that simply won't die. People keep bringing it back up. It's a zombie idea.
Interesting, so what's being said is that it is impossible for a scenario, say an explosion, which both sets up an uncontrolled rotation and takes out the roll control system. This prevents the solar panels from operating with any efficiency, if at all. So you're willing to bet the life and safety of the crew of the MCT on one single mode of power generation.
I deal with critical life support where failure isn't an option, and would result in certain death. A proposal that's analogous to a single mode of power generation would simply not be entertained.
I already have done that. Doesn't seem to matter, the zombie idea keeps coming up.If you're landing away from the pad, you're in a survival situation. You're going to be dead very soon, power or no, without a rescue party. You may have even needed to bail out of the vehicle.You keep saying that. But there are engineering methods to define when something makes sense and when it doesn't. It's called a trade. So help him make the technical trade on mass and simply let the numbers speak for themselves.
Wasting tons on the corner-case of a corner-case (in a way that still is inferior to other solutions) is not a good idea.
An internal combustion engine on MCT is a bad idea that simply won't die. People keep bringing it back up. It's a zombie idea.
Interesting, so what's being said is that it is impossible for a scenario, say an explosion, which both sets up an uncontrolled rotation and takes out the roll control system. This prevents the solar panels from operating with any efficiency, if at all. So you're willing to bet the life and safety of the crew of the MCT on one single mode of power generation.What. I can tell you didn't actually read what I wrote, just what people snipped. I don't care that you work with critical life support systems, because your analysis is ridiculous.
I deal with critical life support where failure isn't an option, and would result in certain death. A proposal that's analogous to a single mode of power generation would simply not be entertained.
I would expect the lines to be pressurized. And I'd expect the time to resolution to be sometime just past the capacity of the battery backup, because that's exactly how Murphy works. And yes, there's a careful balance in trade offs between redundancy and practicality. I'm not saying that an IC engine is the solution (though I could be), rather I'm saying its foolhardy to have only one method to generate power.Interesting, so what's being said is that it is impossible for a scenario, say an explosion, which both sets up an uncontrolled rotation and takes out the roll control system. This prevents the solar panels from operating with any efficiency, if at all. So you're willing to bet the life and safety of the crew of the MCT on one single mode of power generation.
I deal with critical life support where failure isn't an option, and would result in certain death. A proposal that's analogous to a single mode of power generation would simply not be entertained.
Surely that scenario would also create problems for a liquid fuel system too, since the forces in the lines would keep changing directions. Additionally how long would you expect such a scenario to last? would it be longer than the on-board battery supply?
Do you have an IC in your rebreather to act as a backup to the battery? No? Of course not, because that'd be ridiculous. It'd be dangerous, complicated, and would probably end up adding more failure modes than it'd save. You use backup batteries.Sigh. I'm not sure why you are being so aggressive here. It's a conversation. I did mention that the focus is not on using an IC, rather that there should be a completely independent power generator that shares no commonality with solar.
Do you have an IC in your rebreather to act as a backup to the battery? No? Of course not, because that'd be ridiculous. It'd be dangerous, complicated, and would probably end up adding more failure modes than it'd save. You use backup batteries.
NO ONE is saying not to have redundancies and margin. What we're saying is ICE is the wrong solution.
As I said, the backup would be batteries. But you'd also build-in redundancy in the solar array. For life support, no doubt you'd actually have a few oxygen candles and chemical scrubbers, or rely on the large volume of the MCT as a buffer.Do you have an IC in your rebreather to act as a backup to the battery? No? Of course not, because that'd be ridiculous. It'd be dangerous, complicated, and would probably end up adding more failure modes than it'd save. You use backup batteries.Sigh. I'm not sure why you are being so aggressive here. It's a conversation. I did mention that the focus is not on using an IC, rather that there should be a completely independent power generator that shares no commonality with solar.
But to answer your question, my rebreather uses three battery systems, all independent and in physically different enclosures. Each is rated to power the system longer than the consumables could last (O2, diluent, and scrubber). With the loss of one the system can feed off the others. The system uses a digital bus called DiveCAN derived from the ubiquitous auto CAN bus and even with a physical cut between main computer and O2 delivery system, O2 will continue to inject at the current set point. I also have a 100% independent system (literally everything is independent) which allows me to fly the unit manually in the advent of complete system failure. Finally there's an orifice in the O2 side that delivers, depth independent, metabolic rate O2 into the loop (0.6 l/m).
Anyway, I've no desire to argue with you. I respect your posts. I just humbly suggest, IMO, that the real MCT will no doubt have an independent power generator to solar.
Methane/oxygen fuel cellsDo you have an IC in your rebreather to act as a backup to the battery? No? Of course not, because that'd be ridiculous. It'd be dangerous, complicated, and would probably end up adding more failure modes than it'd save. You use backup batteries.
NO ONE is saying not to have redundancies and margin. What we're saying is ICE is the wrong solution.
Yeah I'm struggling to think what other options there are beyond solar/battery power combo and an ICE, unless there is some useful power that can be dragged out of the cooling system (I'm assuming that there won't be nuclear power or RTG on the crew transports)
Is there a guess on the minimum power required to maintain basic functionality of the craft? (I.e. basic life support, basic thermal control, minimal comms, emergency lighting only, basic GNC, etc.)Passive thermal control, manually operated life support or oxygen candles and chemical scrubbers. Windows. gravity-stabilized orientation or manual operation of thruster valves (not a terribly good idea, but possible). Except for comms, basically everything could be done with zero electrical power (though relying on manual operations).
Methane/oxygen fuel cellsDo you have an IC in your rebreather to act as a backup to the battery? No? Of course not, because that'd be ridiculous. It'd be dangerous, complicated, and would probably end up adding more failure modes than it'd save. You use backup batteries.
NO ONE is saying not to have redundancies and margin. What we're saying is ICE is the wrong solution.
Yeah I'm struggling to think what other options there are beyond solar/battery power combo and an ICE, unless there is some useful power that can be dragged out of the cooling system (I'm assuming that there won't be nuclear power or RTG on the crew transports)
Hydrogen/oxygen fuel cells
Stirling engine based on catalytic Methane/Oxygen reaction
Turbine based on Methane/Oxygen combustion
Or use independent batteries. The simplest, most reliable solution.Methane/oxygen fuel cellsDo you have an IC in your rebreather to act as a backup to the battery? No? Of course not, because that'd be ridiculous. It'd be dangerous, complicated, and would probably end up adding more failure modes than it'd save. You use backup batteries.
NO ONE is saying not to have redundancies and margin. What we're saying is ICE is the wrong solution.
Yeah I'm struggling to think what other options there are beyond solar/battery power combo and an ICE, unless there is some useful power that can be dragged out of the cooling system (I'm assuming that there won't be nuclear power or RTG on the crew transports)
Hydrogen/oxygen fuel cells
Stirling engine based on catalytic Methane/Oxygen reaction
Turbine based on Methane/Oxygen combustion
Regenerative fuel cells have very low round-trip efficiency. Use frakking batteries, you'll get a lot more power for your trouble and no moving parts (regenerative fuel cells need pumps, etc).Methane/oxygen fuel cellsDo you have an IC in your rebreather to act as a backup to the battery? No? Of course not, because that'd be ridiculous. It'd be dangerous, complicated, and would probably end up adding more failure modes than it'd save. You use backup batteries.
NO ONE is saying not to have redundancies and margin. What we're saying is ICE is the wrong solution.
Yeah I'm struggling to think what other options there are beyond solar/battery power combo and an ICE, unless there is some useful power that can be dragged out of the cooling system (I'm assuming that there won't be nuclear power or RTG on the crew transports)
Hydrogen/oxygen fuel cells
Stirling engine based on catalytic Methane/Oxygen reaction
Turbine based on Methane/Oxygen combustion
I was subconsciously including all combustion in that... but yep. Once you can split your water back into Hydrogen and oxygen again (using solar) then the fuel cells are an interesting backup, since it's in principle replenishable and probably easier to do than making more methane and the components either side of the reaction are useful.
There's ammonia to burn too.
I was subconsciously including all combustion in that... but yep. Once you can split your water back into Hydrogen and oxygen again (using solar) then the fuel cells are an interesting backup, since it's in principle replenishable and probably easier to do than making more methane and the components either side of the reaction are useful.
There's ammonia to burn too.
I might have missed the lines in so many posts. But what are the requirements exactly on this discussion? Backup for what phase of the trip, for how long, considering which contingencies? I understand that nuclear power will be prepossitioned. And I'm assuming a certain amount of CH4/LOX is supposed to be on the Mars pad, but the MCT would have some margin on its own storage tanks.Or use independent batteries. The simplest, most reliable solution.Methane/oxygen fuel cellsDo you have an IC in your rebreather to act as a backup to the battery? No? Of course not, because that'd be ridiculous. It'd be dangerous, complicated, and would probably end up adding more failure modes than it'd save. You use backup batteries.
NO ONE is saying not to have redundancies and margin. What we're saying is ICE is the wrong solution.
Yeah I'm struggling to think what other options there are beyond solar/battery power combo and an ICE, unless there is some useful power that can be dragged out of the cooling system (I'm assuming that there won't be nuclear power or RTG on the crew transports)
Hydrogen/oxygen fuel cells
Stirling engine based on catalytic Methane/Oxygen reaction
Turbine based on Methane/Oxygen combustion
I was subconsciously including all combustion in that... but yep. Once you can split your water back into Hydrogen and oxygen again (using solar) then the fuel cells are an interesting backup, since it's in principle replenishable and probably easier to do than making more methane and the components either side of the reaction are useful.
There's ammonia to burn too.
There is methane, and oxygen, for propellant. Methane is currently a byproduct of the oxygen generation of one LSS system in the ISS IIRC.
But then you have to reclaim the exhaust, include an electrolysis kit, a compressor... which adds a bunch of mass. So much that you get better storage capacity per mass with... *dundundun* ...a lithium rechargable battery.
I was subconsciously including all combustion in that... but yep. Once you can split your water back into Hydrogen and oxygen again (using solar) then the fuel cells are an interesting backup, since it's in principle replenishable and probably easier to do than making more methane and the components either side of the reaction are useful.
There's ammonia to burn too.
There is methane, and oxygen, for propellant. Methane is currently a byproduct of the oxygen generation of one LSS system in the ISS IIRC.
sure, but is the methane as easy to replenish as hydrogen/oxygen would be? If something is going to be used in a burning system, then you want to be able to replenish it when you have spare energy again.
But then you have to reclaim the exhaust, include an electrolysis kit, a compressor... which adds a bunch of mass. So much that you get better storage capacity per mass with... *dundundun* ...a lithium rechargable battery.
I was subconsciously including all combustion in that... but yep. Once you can split your water back into Hydrogen and oxygen again (using solar) then the fuel cells are an interesting backup, since it's in principle replenishable and probably easier to do than making more methane and the components either side of the reaction are useful.
There's ammonia to burn too.
There is methane, and oxygen, for propellant. Methane is currently a byproduct of the oxygen generation of one LSS system in the ISS IIRC.
sure, but is the methane as easy to replenish as hydrogen/oxygen would be? If something is going to be used in a burning system, then you want to be able to replenish it when you have spare energy again.
(not to mention much better round-trip efficiency)
If you're using a fuel cell (which is much better than a small IC, though still bad IMO), just vent the exhaust or possibly use the water for crew use (like on Shuttle). Anything else will just result in a lithium battery being a much better option.
As I said, the backup would be batteries. But you'd also build-in redundancy in the solar array. For life support, no doubt you'd actually have a few oxygen candles and chemical scrubbers, or rely on the large volume of the MCT as a buffer.As mentioned previously, the concern is power "generation" for all modes/situations; in space or on Mars. In a life critical context, having two or more independent power generation capabilities is obvious. However, batteries are for power "storage", not generation. Sure you can use batteries when your power generation fails but only to the extend of the amount of power stored in the batteries and the number of physical batteries you have.
What about in a dust storm on mars? Your solar panels probably aren't going to be much use then.
In a life critical context, having two or more independent power generation capabilities is obvious.
Though I'm not for or against any technology, I do hold the opinion that an ICE (or derivative technology) genset is the best, simplest, non-nuclear solution.
Regenerative fuel cells have very low round-trip efficiency. Use frakking batteries, you'll get a lot more power for your trouble and no moving parts (regenerative fuel cells need pumps, etc).
There seems to be some sort of allergy to batteries here.
Batteries just don't have the energy density to even come close to Methane LOX, either by combustion or reaction, mass for mass. That's why a ICE sedan (aka Audi or BMW) similar to Tesla Model S weighs half as much and has twice (give or take) the range.
It's also why Methane LOX is rocket fuel and a battery powered rocket engine will never exist.
A dust storm on Mars (which have covered the entire planet) is possibly one of those "non-deterministic" lapses that could last months.
A dust storm on Mars (which have covered the entire planet) is possibly one of those "non-deterministic" lapses that could last months.
Solar panels still work during dust storms. (Dust is actually easier for modern solar panels than water-droplet clouds on Earth.)
1) Mars surface is reached by a larger version of the Dragon, with engines suitable for precision landing.Neither. Or rather, something in between.
2) Mars surface is reached by a larger version of the F9R, with engines suitable for both orbital launch and precision landing.
You need maybe 6-7km/s to return to Earth from the surface of Mars. You can do that in a single stage with TPS, though it is hard.
The biggest concern is that the "Mars Dust" is more of salty (sticky), abrasive (sharp edges) and much smaller particles (talcum powder sized) which is electrically charged attracted to any man-made elements, not the "Earth dust" we think of ones from the dry deserts.
Needs to be able to withstand entry to Mars, reentry back to Earth, and perhaps aerocapture. Definitely need a thermal protection system for that. But it'll have to be extremely lightweight.1) Mars surface is reached by a larger version of the Dragon, with engines suitable for precision landing.Neither. Or rather, something in between.
2) Mars surface is reached by a larger version of the F9R, with engines suitable for both orbital launch and precision landing.
You need maybe 6-7km/s to return to Earth from the surface of Mars. You can do that in a single stage with TPS, though it is hard.
Thank you! So something in between a squat capsule and a pencil-shaped booster.
Presumably with short, stubby legs to support the weight of a fully tanked craft?
Presumably with engines angling outward (Dragon-style) to direct blast debris away from craft?
I am not sure that I understood the relevance of TPS (temperature protection system)?
I am not sure that I understood the relevance of TPS (temperature protection system)?
Cars on Earth don't have to carry their own oxygen. As RB said, you are letting your experience on Earth mislead you.No I'm not... I'm directly looking at the modes of the entire mission and matching it with existing technology or technology that could be applied. Energy density per kilogram mass and the efficiency of turning that energy into useful work is the factors to determine the viability of using the solution for the mission.
Solar electric propulsion has vastly better "range" than any chemical rocket.Yes, SEP has great ISP but no thrust because the amount of energy required to ionize a larger reaction mass to provide greater thrust is ridiculous. SEP is great for low mass long duration flights measured in years. It's not practicable for high mass "as short as possible" duration flights like a manned Mars mission.
Solar panels still work during dust storms. (Dust is actually easier for modern solar panels than water-droplet clouds on Earth.)So you lose some power... How much power and for how long? No one can answer this question because it's non-deterministic and we have little experience to fall back to. Remember you have power ECLSS and ISRU because your lives depend on it. This is why two primary methods of power generation is required.
The power generated will drop, of course. But most of the power required on Mars will be for ISRU. You simply suspend power hungry ops during the worst of the storm. (Which, judging by MER-Opportunity, is only a few days even in a month long dust storm.)
Question: what is the best material to make MCT out of?The simplicity of your initial question is daunting. For me, it would have to be broken down into a long list, just for the major components, because each one will require drastically different engineering considerations. (Then I would duck and defer to a materials specialist.)
The options that I come up with are:
1) Aluminum - Lithium alloy just like Falcon.
2) Carbon Fiber composite using a honeycomb core.
3) Titanium single shell.
4) Titanium double shell with a core of some kind.
Considerations include heat tolerance, insulation needed, corrosion on Mars and Earth.
Stiffness when not pressurized. Thick core walls are much stiffer.
The material that I like is Titanium with hollow spaces to form a a 2 wall Ti sandwich with a Ti honeycomb core.
Obviously this is very difficult to make as it is hard to weld Ti alloys in air. But if that could be worked out it would seem to be the most durable.
Insulation is needed not only on Earth, but on Mars when accumulating ISRU propellants, so an inner layer of insulation is needed. For Carbon Fiber or Al-Li an outer layer of insulation would also be needed.
@SpaceXTrip
Today I celebrate nine years @SpaceX! It's amazing how much it's changed, and I am excited for what the future holds https://t.co/oNLgPQM9my (http://"https://t.co/oNLgPQM9my")
Picard:"Somehow I doubt this will be the last ship to carry the name Enterprise."
Trip Harris is celebrating 9 with SpaceX today, and is currently is Manager of Falcon landings. The YouTube clip he chose to link to his celebratory tweet is....interesting.
https://twitter.com/SpaceXTrip/status/733869950067036160 (http://"https://twitter.com/SpaceXTrip/status/733869950067036160")Quote@SpaceXTrip
Today I celebrate nine years @SpaceX! It's amazing how much it's changed, and I am excited for what the future holds https://t.co/oNLgPQM9my (http://"https://t.co/oNLgPQM9my")QuotePicard:"Somehow I doubt this will be the last ship to carry the name Enterprise."
I don't know if it's been talked about yet - but what about MCT/BFS and the hoverslam? As F9R is showing, hoverslam can be pretty damn rough/tough on the rocket. So what would MCT/BFS face during descent on Earth or Mars, and would hoverslam be feasible? Or are you just going to need much more deeply throttleable engines?It's not necessarily the hoverslam that is rough on the rocket, it's the reentry. Shield from reentry heating, and hoverslam is simply a more efficient way to land (though maybe a bit sketchy with crew).
I don't know if it's been talked about yet - but what about MCT/BFS and the hoverslam? As F9R is showing, hoverslam can be pretty damn rough/tough on the rocket. So what would MCT/BFS face during descent on Earth or Mars, and would hoverslam be feasible? Or are you just going to need much more deeply throttleable engines?It's not necessarily the hoverslam that is rough on the rocket, it's the reentry. Shield from reentry heating, and hoverslam is simply a more efficient way to land (though maybe a bit sketchy with crew).
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Hopefully (going out on a limb to say certainly) BFS will have a lower thrust-to-weight ratio compared with a F9 stage one.
What do you mean 3.8g and 4.5g, thouse are incredible acceleration rate and likely beyond what crew could tolerate,
Humans can handle 8-9 gees just fine if oriented the correct way.
Humans can handle 8-9 gees just fine if oriented the correct way.
Not for very long, though. NASA limited sustained gees on Apollo flights to 4.5 g and Shuttle flights to 3 g for very good physiological reasons. Spacex should be aiming for a similar sustained g load for manned launches and reentries for the same reasons.
8gees would be fine, too. But I suspect if they limit it, they'll do it for structural reasons, i.e. they could have a lighter structure if they limited hoverslam acceleration to, say, 5 gees vs 8 gees.Humans can handle 8-9 gees just fine if oriented the correct way.
Not for very long, though. NASA limited sustained gees on Apollo flights to 4.5 g and Shuttle flights to 3 g for very good physiological reasons. Spacex should be aiming for a similar sustained g load for manned launches and reentries for the same reasons.
Sure, for long term/sustained... but ~4 g for a hoverslam landing just isn't that big of a deal at all.
I'm still looking for a clarification by envy887, is this a calculation of the vehicles acceleration rate during a hover-slam landing? I had though he meant a launch but as a landing g force that makes a lot more sense.
Still we need to remember that the crew/cargo inside the vehicle gets an additional local g worth of force upon them (sanity check, if the rocket were hovering we would feel 1 g, not weightlessness). So that means a force of 5.5 g at Earth and around 4.1 g on Mars. In practice I'm sure you would see throttling of engines and a gentler deceleration in the 3 g neighborhood.
I'm still looking for a clarification by envy887See my reply above. I was figuring a hover-slam at burnout mass, not a launch.
....
Not sure how you are calculating that.He added local gravity to my calculations above, which were based on my rough estimates of MCT dry mass, engine count, and engine thrust. That is indeed the perceived acceleration during landing.
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Not sure how you are calculating that.He added local gravity to my calculations above, which were based on my rough estimates of MCT dry mass, engine count, and engine thrust. That is indeed the perceived acceleration during landing.
....
What do you mean 3.8g and 4.5g, thouse are incredible acceleration rate and likely beyond what crew could tolerate,
I don't think so... 4 g isn't that big of a deal with correct positioning. The Shuttle was limited to 3, but Apollo astronauts took way more g's than that.
Humans can handle 8-9 gees just fine if oriented the correct way.
Not for very long, though. NASA limited sustained gees on Apollo flights to 4.5 g and Shuttle flights to 3 g for very good physiological reasons. Spacex should be aiming for a similar sustained g load for manned launches and reentries for the same reasons.
Yes. 100t to Mars, 25t to Earth.Not sure how you are calculating that.He added local gravity to my calculations above, which were based on my rough estimates of MCT dry mass, engine count, and engine thrust. That is indeed the perceived acceleration during landing.
....
Different payloads? if the mass and thrust are the same then perceived acceleration will be the same. If the craft is nearly empty (of propellant) has the same payload and can thrust at 4g in free fall, then in a hover slam the passengers experience 4g on Mars or Earth but on Mars they are decelerating at 3.62g and on Earth at 3g.
Yes. 100t to Mars, 25t to Earth.Not sure how you are calculating that.He added local gravity to my calculations above, which were based on my rough estimates of MCT dry mass, engine count, and engine thrust. That is indeed the perceived acceleration during landing.
....
Different payloads? if the mass and thrust are the same then perceived acceleration will be the same. If the craft is nearly empty (of propellant) has the same payload and can thrust at 4g in free fall, then in a hover slam the passengers experience 4g on Mars or Earth but on Mars they are decelerating at 3.62g and on Earth at 3g.
nadreck your forgetting that the passengers are not part of the vehicle, the vehicle acceleration upward towards the occupants but the occupants are still affected by gravity and are accelerated downward into the vehicle too, the sum of the forces are what they feel.
on Earth the landing mass is expected to be a lower 125 mt resulting in a higher 6.7g force, but this is due ONLY to the lower mass and has nothing to do with differences in gravity fields.
And again I reiterate I expect throttle-down of rockets to give a gentler ~3g landing at all locations.
Humans can handle 8-9 gees just fine if oriented the correct way.
Not for very long, though. NASA limited sustained gees on Apollo flights to 4.5 g and Shuttle flights to 3 g for very good physiological reasons. Spacex should be aiming for a similar sustained g load for manned launches and reentries for the same reasons.
Sure, for long term/sustained... but ~4 g for a hoverslam landing just isn't that big of a deal at all.
@CNBCNow
SpaceX's Elon Musk says could launch flight to Mars with people aboard in 2024 and they would arrive in 2025. #CodeCon
...40 tonnes of LH2 needs a 575 m^3 tank kept below 30K... the tankage and cooling are going to weigh a lot more than a few percent.
Since only about 5% of a stochiometric Methane/LOX fuel is hydrogen, they could bring down about 40 tons of molecular H with each BFS, well within its payload capacity (of 100 tons). Even if we add a few percent of that for the mass of the tank, this is still very feasible.
...
After Elon's Mars schedule announcement, I was wondering about this (I don't know if this has been discussed before, I presume so but perhaps not in this specific context) :
A way to speed up the development of the first manned Mars mission (to keep up with the aggressive 2024/5 schedule) would be to bring molecular H to the Martian surface (which could also serve as a radiation shield on the way there). If the BFS must have a dV capacity of 6 km/s and has a dry mass of 200 tons, it will need about 800 tons of fuel for the return journey. Since only about 5% of a stochiometric Methane/LOX fuel is hydrogen, they could bring down about 40 tons of molecular H with each BFS, well within its payload capacity (of 100 tons). Even if we add a few percent of that for the mass of the tank, this is still very feasible.
All that the first SpaceX astronauts would then have to do after landing is react the H with the atmospheric CO2 in the Sabatier reaction - this could virtually been done without any human supervision or involvement. This is also a much simpler and cleaner (and less energy intensive) process than having the crew looking for water ice, digging it out, cleaning it, testing it for any potential chemical agents (e.g., peroxides, salts, etc.) and removing them before electrolysis, etc. This is also a technology that 1) is already in development and 2) would be reasonably easy to demonstrate with a precursor Red Dragon mission. The same cannot be said of actively digging for water ice on Mars.
While I am certain the plan is for a Mars base to eventually provide BFS's with fuel synthesized from martian water ice and atmospheric CO2, the very first landings could / should skip this step and bring the H from Earth instead.
Also, consider that the initial crew would have to dig up 360 tons of water ice to produce the fuel, or 60 tons each if they are 6 astronauts.... Hard even if you bring some landmoving equipment....
@kch: sure, I am aware that the idea of bringing in H is from Mars Direct (and as such I should perhaps have said so), but I mean in the specific context of the BFS and speeding up the schedule to make a 2024/5 manned SpaceX mission to Mars feasible. ISRU is (well, presumably so) an integral part of the SpaceX architecture, but some of it (like digging up water ice) has very low technological readiness and this is unlikely to change before 2024/5. Also, consider that the initial crew would have to dig up 360 tons of water ice to produce the fuel, or 60 tons each if they are 6 astronauts.... Hard even if you bring some landmoving equipment....
Envy887 made an excellent point on the volume and cooling systems this would require. This is the kind of answer and input I was going after. A tank this size would only be 3 m high at 15 diameter for the BFS. Even if cooling etc take up 10 tons or more, at 100 tons total you still have some serious payload capacity left.
Did Musk say humans would LAND on Mars, or just be sent TO Mars in 2024? That later leaves open a purely orbital mission.
Assuming its a landing in 2025, ISRU water from martian ground ice has still extremely low technological readiness,
I wonder if Elon is an investor in Boston Dynamics...
The other option is to land enough propellant on Mars for the ascent of the escape vehicle. This is possible, but not nearly as economical, potentially requiring one MCT to stay in low Mars Orbit with its payload of propellant that came from Earth and 3 or 4 cargo MCTs arriving with just propellant for the emergency ascent vehicle. The complexity of this is that propellant transfer from landed craft is required. The ISRU example might be accomplished with the ISRU MCT refueling itself and it being the emergency return vehicle not just an ascent to orbit vehicle.The emergency ascent spacecraft can be very small, like a Dragon, without the heat-shield and SDs. Launch it to reach the Earth-return MCT in low Mars orbit will not require an enormous amount of fuel. If the capsule + ascent stage is 10 t dry, then 30 t wet is enough. The ascent stage can be derived (shortened) from F9 second stage, so the ascent vechicle need not be developed from scratch. A single MCT can land the fueled ascent vechicle together with extra cargo, or even the hab modul.
So... while SpaceX may get plenty of kudos for landing the first humans on Mars, it will pretty much kill their colonization plans if those first humans end up on a one-way trip with no hope of either survival or return. I'm therefore awfully confident that they will have anticipated most, if not all, of the obvious common-sense precautions we're all coming up with here, and will have plans in place to cope with those situations if they arise. :)
I wonder if Elon is an investor in Boston Dynamics...
That rings a bell - will do some searching and report back if it turns out to be factual.
Why do you think bringing down H to the surface is a weak point of Mars Direct? The motivation is not so much that there is no H on Mars, it's that this allows you to take one big unknown (how youw get from sub-surface dirty ice to CH4 in your tanks) out of the equation.
The other option is to land enough propellant on Mars for the ascent of the escape vehicle. This is possible, but not nearly as economical, potentially requiring one MCT to stay in low Mars Orbit with its payload of propellant that came from Earth and 3 or 4 cargo MCTs arriving with just propellant for the emergency ascent vehicle. The complexity of this is that propellant transfer from landed craft is required. The ISRU example might be accomplished with the ISRU MCT refueling itself and it being the emergency return vehicle not just an ascent to orbit vehicle.The emergency ascent spacecraft can be very small, like a Dragon, without the heat-shield and SDs. Launch it to reach the Earth-return MCT in low Mars orbit will not require an enormous amount of fuel. If the capsule + ascent stage is 10 t dry, then 30 t wet is enough. The ascent stage can be derived (shortened) from F9 second stage, so the ascent vechicle need not be developed from scratch. A single MCT can land the fueled ascent vechicle together with extra cargo, or even the hab modul.
Assuming its a landing in 2025, ISRU water from martian ground ice has still extremely low technological readiness, much less doing it fully automated (for prepositioning)! The energy required is also much more than the one needed for electrolysis, its also the digging itself, warming up the ice to form water (in a -55C environment), cleaning and chemically purifying it (by destillation?), etc. And all that without human supervision? If we take Elons requirement of prepositioned fuel serious, I see no other way than to bring in H with an unmanned mission in 2022/3 and producing the fuel during the two years that follow.
Another interesting tidbit to take away from the RecodeDotNet interview with Musk is that the trip will (initially) take less than 90 days. We should be able to deduct how much propellant would be needed to do such a short trip time with that much payload. He also said that they want to get to shorter trips (less than 30 days) at a later time. I presume that this would require some change to the whole architecture, though.
Assuming its a landing in 2025, ISRU water from martian ground ice has still extremely low technological readiness, much less doing it fully automated (for prepositioning)! The energy required is also much more than the one needed for electrolysis, its also the digging itself, warming up the ice to form water (in a -55C environment), cleaning and chemically purifying it (by destillation?), etc. And all that without human supervision? If we take Elons requirement of prepositioned fuel serious, I see no other way than to bring in H with an unmanned mission in 2022/3 and producing the fuel during the two years that follow.
You cane extract water from the Martian atmosphere. It requires power (you get about 1.5 kg/kW/sol, if I remember correctly), but the equipment is not particularly massive. It has the advantage of simplicity (little need for purification) and may be preferable for initial missions, switching to local mined water once it has been located and characterised etc.
The other option is to land enough propellant on Mars for the ascent of the escape vehicle. This is possible, but not nearly as economical, potentially requiring one MCT to stay in low Mars Orbit with its payload of propellant that came from Earth and 3 or 4 cargo MCTs arriving with just propellant for the emergency ascent vehicle. The complexity of this is that propellant transfer from landed craft is required. The ISRU example might be accomplished with the ISRU MCT refueling itself and it being the emergency return vehicle not just an ascent to orbit vehicle.The emergency ascent spacecraft can be very small, like a Dragon, without the heat-shield and SDs. Launch it to reach the Earth-return MCT in low Mars orbit will not require an enormous amount of fuel. If the capsule + ascent stage is 10 t dry, then 30 t wet is enough. The ascent stage can be derived (shortened) from F9 second stage, so the ascent vechicle need not be developed from scratch. A single MCT can land the fueled ascent vechicle together with extra cargo, or even the hab modul.
Fleshing this out a little more:
First, I am happy to entertain this for the first expedition so that they can leave scale ISRU until people are there.
2nd, for this to be worth more savings than the cost of sending enough propellant for an MCT return, the development has to be truly minimal, as well the development of this becomes critical path for the Mars Mission and either has to fly with the 2024 armada or the window before.
3rd, having a single one of these limits the first expedition crew to 7 (presuming that is still the maximum complement on a D2) 2 limit it to 14. By the time you have 2 the development cost really has to be just about zero because 3 or 4 and you just use MCTs to ferry propellant down for 1 MCT to make orbit again.
technical difficulties that may push the development from minimal to too complex to be worth doing:
1. Modifying MCT to not only accommodate and land this craft without damaging it but to support its launch
2. keeping RP-1 from freezing/gelling (doable but needs design and development work)
3. modifying D2 to have the lifespan needed for this (going from nominally 9 months or so to 4 years)
4. modifying Falcon S2 to short length required
How exactly are they going about scraping up 100s of tons of ice robotically for each return mission, processing it and all - and without any humans supervising the process?
There are always nay-sayers regarding the business of ISRU on Mars, with the whole business of water mining being dismissed as 'too difficult'. I wonder if we're not exploiting the environmental properties of Mars sufficiently, however.
I'm thinking of an almost passive distillation water-mine, where broken-up winter temperature perchlorate-rich ice is shovelled into a large vessel, which is heated in the summer by sunshine. The ice sublimates, and the water-laden air rises into a second, insulated chamber which is never allowed to heat beyond winter temperatures and where it freezes out as rime or snow. You harvest the output, and perhaps distill again.
My starting point for this was the thought of the traditional adiabatic atmospheric 'springs' constructed in various Earth deserts using no resources other than piles of rock and a knowledge of the local winds. Obviously, Mars is different, but low-tech solutions would be a good tradeoff, using scarce imported resources only where needed.
We don't know that any water mining will be done robotically, that is an assumption based on the rather dubious proposition that SpaceX won't send a crew until there is a fully fueled return BFS waiting on the martian surface.The idea (originating from Mars Direct) was to ensure the safety of the crew.
There are always nay-sayers regarding the business of ISRU on Mars, with the whole business of water mining being dismissed as 'too difficult'. I wonder if we're not exploiting the environmental properties of Mars sufficiently, however.
I'm thinking of an almost passive distillation water-mine, where broken-up winter temperature perchlorate-rich ice is shovelled into a large vessel, which is heated in the summer by sunshine. The ice sublimates, and the water-laden air rises into a second, insulated chamber which is never allowed to heat beyond winter temperatures and where it freezes out as rime or snow. You harvest the output, and perhaps distill again.
My starting point for this was the thought of the traditional adiabatic atmospheric 'springs' constructed in various Earth deserts using no resources other than piles of rock and a knowledge of the local winds. Obviously, Mars is different, but low-tech solutions would be a good tradeoff, using scarce imported resources only where needed.
Day/night temperature swings probably would be enough, that would allow a far greater volume of water to be collected.
We don't know that any water mining will be done robotically, that is an assumption based on the rather dubious proposition that SpaceX won't send a crew until there is a fully fueled return BFS waiting on the martian surface.The idea (originating from Mars Direct) was to ensure the safety of the crew.
As Zubrin said in "The Case for Mars" (pg. 70-71):
"To ensure our Mars crew would not be stranded, the ERV would fly one launch opportunity, or twenty-six months, prior to the launch of the astronauts. Thus all the propellant would be made before the crew ever left Earth, and since the propellant plant was flown to Mars integrated with the ERV there was no question about landing 'within a hose length.' The plumbing that would deliver the Mars-manufactured propellant from the chemical synthesis unit into the ERV's fuel tanks would be hardwired, installed on Earth."
"if the ERV [Earth Return Vehicle] is sent first, the crew will know before they even leave Earth that they have a fully functional Mars ascent and Earth return system waiting for them on the Martian surface, one that has already survived the trauma of landing. In contrast, a crew that lands with their ascent system can only guess in what shape their Mars ascent vehicle will be after they've hit the surface."
SpaceX did test fire F9 engines after they've been through supersonic retropropulsion in conditions relevant to Mars EDL systems development, so the latter might not be a problem.
SpaceX is definitely going to do water ISRU. Here are some of the early instruments that are proposed to fly on Red Dragon:
http://digitalvideo.8m.net/SpaceX/RedDragon/karcz-red_dragon-nac-2011-10-29-1.pdf
slide 8
"Water extraction system &
propellant production system.
HEOMD funded KSC & JSC
ISRU activity at TRL 5."
And Musk has said they want robotic mining to produce propellant on Mars. You cannot get more direct than that.
I am pretty sure there is something very wrong with this idea, but I cannot see it. Send an early MCT with most of it's cargo as water. Include some pumps and soft insulated containers that could be set on the ground and pumped full. Design them to withstand freezing and also have the capability to electrically melt the ice. The goal for the arrivers can still be to mine the water locally, with a back up source for ISRU and consumption.
No need for cryogenic storage of H2.
Having 90 tons or so of pure water as a contingency might be worth the price of tying up an expensive space ship for two plus years. The water MCT could also be a test mission without having any expensive gear at risk, with water as a sort of a mass simulator.
Matthew
Edited for a stray 'O' and to add the bit about H2
@Robobeat: if you read my first post on the matter carefully, I am not suggesting they are not persueing eventual ISRU of Martian water. Bringing in H would just be an interim solution until full scale ISRU is capable of providing the necessary fuel. How exactly are they going about scraping up 100s of tons of ice robotically for each return mission, processing it and all - and without any humans supervising the process?That's like a few hundred pounds a day if spread over a year. Not that bad. I could definitely see it done using scaled up versions of some mining rovers I've seen.
I just don't see this happening in 9 years. You'd probably need multiple design iterations for the entire process to work, and there are only 3 transfer windows left before the process *has* to work if the 2024/5 mission date is to be met.....Yeah, and they're going to utilize each one. 2024/5 doesn't leave a lot of margin, but that's the "if everything goes according to plan" date. When Musk gives that kind of date, usually it's the earliest it can be done, not the "we'll definitely be able to do it by then" date. Which means you should probably expect it to slip.
SpaceX is definitely going to do water ISRU. Here are some of the early instruments that are proposed to fly on Red Dragon:
http://digitalvideo.8m.net/SpaceX/RedDragon/karcz-red_dragon-nac-2011-10-29-1.pdf
slide 8
"Water extraction system &
propellant production system.
HEOMD funded KSC & JSC
ISRU activity at TRL 5."
And Musk has said they want robotic mining to produce propellant on Mars. You cannot get more direct than that.
Musk will have enslaved an entirely robotic populated planet. Without their bio-masters present don't be surprised if there isn't a revolution. You've been warned!
QuoteI just don't see this happening in 9 years. You'd probably need multiple design iterations for the entire process to work, and there are only 3 transfer windows left before the process *has* to work if the 2024/5 mission date is to be met.....Yeah, and they're going to utilize each one. 2024/5 doesn't leave a lot of margin, but that's the "if everything goes according to plan" date. When Musk gives that kind of date, usually it's the earliest it can be done, not the "we'll definitely be able to do it by then" date. Which means you should probably expect it to slip.
But again, Musk has repeatedly said they're going to mine water. I really don't think they'll mess around with liquid hydrogen. Instead, they'll double-down on getting water extraction to work. And some of that they can test here on Earth before sending it to Mars.
We don't know that any water mining will be done robotically, that is an assumption based on the rather dubious proposition that SpaceX won't send a crew until there is a fully fueled return BFS waiting on the martian surface.The idea (originating from Mars Direct) was to ensure the safety of the crew.
As Zubrin said in "The Case for Mars" (pg. 70-71):
"To ensure our Mars crew would not be stranded, the ERV would fly one launch opportunity, or twenty-six months, prior to the launch of the astronauts. Thus all the propellant would be made before the crew ever left Earth, and since the propellant plant was flown to Mars integrated with the ERV there was no question about landing 'within a hose length.' The plumbing that would deliver the Mars-manufactured propellant from the chemical synthesis unit into the ERV's fuel tanks would be hardwired, installed on Earth."
"if the ERV [Earth Return Vehicle] is sent first, the crew will know before they even leave Earth that they have a fully functional Mars ascent and Earth return system waiting for them on the Martian surface, one that has already survived the trauma of landing. In contrast, a crew that lands with their ascent system can only guess in what shape their Mars ascent vehicle will be after they've hit the surface."
SpaceX did test fire F9 engines after they've been through supersonic retropropulsion in conditions relevant to Mars EDL systems development, so the latter might not be a problem.
The ERV doesn't have to be fully fuelled before the crew leaves Earth. It merely needs to have landed safely and be producing propellant at a rate that means it will be fully fuelled by the time the crew needs to use it to launch back to Earth. No doubt there will be margins to take account of any possible breakdown after the crew has set off for Mars. Set against that is the possibility of the crew making repairs and the additional output of any further ISRU equipment they bring with them. If additional margins and/or redundancy is required you can always send two ERVs on the first trip, or perhaps an ERV and a dedicated ISRU propellant production lander.
Go back to that Max Fagin (former SpaceX intern) retropropulsion thesis defense video.
Whole vehicle lands, but only an upper portion returns to Earth. If the propulsion is in the upper portion (Dragon 2 heritage) the left behind cargo section can be used for most anything; habitation, cargo, or perhaps a dual purpose.
How about some cargo sections hauling expandable & repurposeable tanks full of distilled water (thinking Thin Red Line's expandable tank tech.) AKA, 'how to ship hydrogen and oxygen without cryocollers', and the cargo bay volume vould be repurposed for colony use later.
Thirdly, the robotics is hard and is special purpose to the initial landing. It is probably at least an order of magnitude (perhaps two orders of magnitude) harder than Curiosity. The robotics would almost certainly be the long pole with a 2022 launch this would be really hard. The chances are that it would not work perfectly the first time. Correcting the robotics could easily take several attempts, each of which are hard to do before launch of the next synod. Worst case SpaceX could take a decade before they got the ISRU robotics correct, each year burning through a $1B or two.I think this is a valid concern but I am expecting Musk thinks he can hack this in time. Asked about driving AI he called it a "solved problem" and said widespread class 4 (handless, not requiring attention) driving was 2 years away technically. So maybe he thinks ISRU robotics won't be that hard? You'd expect a lot more visible precursor work but if there is work going already, SpaceX is very tightlipped about it!
Thirdly, the robotics is hard and is special purpose to the initial landing. It is probably at least an order of magnitude (perhaps two orders of magnitude) harder than Curiosity. The robotics would almost certainly be the long pole with a 2022 launch this would be really hard. The chances are that it would not work perfectly the first time. Correcting the robotics could easily take several attempts, each of which are hard to do before launch of the next synod. Worst case SpaceX could take a decade before they got the ISRU robotics correct, each year burning through a $1B or two.I think this is a valid concern but I am expecting Musk thinks he can hack this in time. Asked about driving AI he called it a "solved problem" and said widespread class 4 (handless, not requiring attention) driving was 2 years away technically. So maybe he thinks ISRU robotics won't be that hard? You'd expect a lot more visible precursor work but if there is work going already, SpaceX is very tightlipped about it!
In line with the philosophy of reuse mfck articulated...
I could see the MCT being "partially strippable"... if you had cabins for 100, but are only returning 20 people why carry all of that back? if the fittings, panels, wiring, plumbing etc was designed to be modular and reusable it could be removed by humans and used in outfitting the base.
If they simply install a large (but very lightweight) crane
This would be pretty easy to automate.
As Zubrin said in "The Case for Mars" (pg. 70-71):
"the ERV would fly one launch opportunity, or twenty-six months, prior to the launch of the astronauts.
[...]
the crew will know before they even leave Earth that they have a fully functional Mars ascent and Earth return system waiting for them on the Martian surface [...]"
Thirdly, the robotics is hard
Remember Mars has a atmosphere and has water cycles (water vapor and water ice, does not have liquid phase). So to mine all the water you need, you compress the air and liquefy the water vapor in the atmosphere. Admittedly you would have to compress a lot of atmosphere for a little bit of water but the advantage is you do not have to go out and dig it up! Once the quantities needed become very high then comes the mining of water. But for the initial methalox manufacture for the return trip this method would be an easy one to implement and would not require complex equipment just plenty of power. Compressing gas takes a great deal of energy. But you need the atmosphere compressed for other parts of the methalox manufacture as well.
The point about driving AI is not driving per se, although that's a useful part of the overall problem... it's that they implemented something that learns.
...
with two years of learning time the units might be better than when they arrived...
Compression requires much too much work. It should be fairly simple to use dessicants to dry out the air, then extract the water by heating the dessicant. This is one of the most common technologies today to dry compressed air or indoor ice rinks (and was part of the original Mars direct proposal). alternatively you can compress a refrigerant, and use that to create a cold surface on which the water in the air will condense out. Any of these solutions will require a minimum of about 1000 btu/lb or 2300 kJ/kg of water, the phase change energy of water.Remember Mars has a atmosphere and has water cycles (water vapor and water ice, does not have liquid phase). So to mine all the water you need, you compress the air and liquefy the water vapor in the atmosphere. Admittedly you would have to compress a lot of atmosphere for a little bit of water but the advantage is you do not have to go out and dig it up! Once the quantities needed become very high then comes the mining of water. But for the initial methalox manufacture for the return trip this method would be an easy one to implement and would not require complex equipment just plenty of power. Compressing gas takes a great deal of energy. But you need the atmosphere compressed for other parts of the methalox manufacture as well.
Not really possible to get water by compression. Water vapour is about 0.03%, so for every kg of water you would need to compress and liquefy 3 tonnes ofwater<air>.
...on Earth. But sure, power-shovel or bucketwheel would work, too.If they simply install a large (but very lightweight) crane
You presumably mean a power-shovel or dragline. A crane isn't mining equipment.
I never said you'd use a bucket-line, you're the one who mentioned that. :)This would be pretty easy to automate.
It really isn't. It takes a lot of finesse to run a bucket at the end of a flexible line. Even a basic bucket excavator would be better. But a bucket-wheel might the best option for robotic operation.
The point about driving AI is not driving per se, although that's a useful part of the overall problem... it's that they implemented something that learns.
...
with two years of learning time the units might be better than when they arrived...
The way that most successful AI learning works is that they use large amounts of data for training. For instance Tesla collections 1 million miles of driving data every 10 hours. This data is used to train the algorithms, but also used to evaluate how the AI does on real world data, it is also used to detect real world corner cases and allows humans to tweek the AI to handle these unusual situations better. Tesla also runs autopilot in shadow mode on customers cars, comparing what new algorithms do with what the human driver does.
For rover driving on Mars we have none of that, no large data sets, no real world data around the landing site, no ability to look at corner cases and no ability to compare AI results with what a human would do. This applies not only to driving but also to many other tasks the mining system would have to perform.
Robotic deployment of thin-solar panels to power an atmospheric water adsorption system is the most practical means to refuel an initial landing vehicle, it requires the least knowledge about the martian subsurface and is the most reliable due to minimal moving parts and minimal contact with abrasive regolith.Absolutely, given what we know now. But a Red Dragon or two could inform more efficient techniques, if conditions supported it.
...on Earth.If they simply install a large (but very lightweight) craneA crane isn't mining equipment.
[...]
I never said you'd use a bucket-line
A dragline is a specific type of machine. You can have a bucket or a scoop without having a dragline. (I accidentally said "bucket line," but we were talking about a dragline... sometimes it's better NOT to trim quotes.)...on Earth.If they simply install a large (but very lightweight) craneA crane isn't mining equipment.
[...]
I never said you'd use a bucket-line
Okay, in all seriousness, how did you envision using a crane to do mining, without a bucket or scoop?
Robotic deployment of thin-solar panels to power an atmospheric water adsorption system is the most practical means to refuel an initial landing vehicle, it requires the least knowledge about the martian subsurface and is the most reliable due to minimal moving parts and minimal contact with abrasive regolith.Absolutely, given what we know now. But a Red Dragon or two could inform more efficient techniques, if conditions supported it.
Raptor with a 4 m nozzle loses about 1% of ISP compared to a 4.8 m nozzle: 376 s vs 380 s.
This is based on sim in RPA lite using: Methane/LOX at:
9.7 MPa chamber pressure (same as Merlin)
2.8 O/F ratio (optimum for methalox at 9.7 MPa)
165 expansion ratio for the 4.8 m nozzle (same as Merlin Vac)
115 expansion ratio for the 4 m nozzle (assuming same throat diameter as the 4.8 m nozzle)
From the FH discussion speculating on Raptor upper stage for FH
Raptor with a 4 m nozzle loses about 1% of ISP compared to a 4.8 m nozzle: 376 s vs 380 s.
This is based on sim in RPA lite using: Methane/LOX at:
9.7 MPa chamber pressure (same as Merlin)
2.8 O/F ratio (optimum for methalox at 9.7 MPa)
165 expansion ratio for the 4.8 m nozzle (same as Merlin Vac)
115 expansion ratio for the 4 m nozzle (assuming same throat diameter as the 4.8 m nozzle)
IF Rvac does have a 4.8m diameter, what does this say about max # of Rvac engines for the hypothetical 2nd stage BFS @ a given stage diameter? Makes the case for >10m with even 15m having issues with # of Rvacs.
10m is out for > 3 engines
12m only fits 4 engines
15m seems OK for a ring of 6 engines, or 5 engines and a center engine. Maybe too little swivel clearance for CE.
My guess is 15m BFS with slightly under 4.8m Rvac diameter bell, say ~4.5m such that a center engine is feasible.
4.5m also allows a 12m BFS to hold 5 engines in a ring.
By more US DV I mean (1) less performance needed from 1st stage BFR making the big boy's re-use environ slightly more benign and lowering cost {My unsubstantiated assumption}, (2) when you re-fuel in LEO you still need the tankage structure mass for propellant for the BFS's DV needed for fast transfer to Mars, say 90-120 days whatever, possible or maybe not propulsive breaking, aerobraking and powered landing.
Using today's F9 as a model, the Rvac would have ~ 110% the thrust of the assumed 230mT Elon referred to.
Honestly, you can't figure it out? Different requirements from Earth mining. Don't need nearly the throughput. You're talking like a factor of 1000-100,000x less than large mining equipment on Earth, like a few tons of soil per day to fuel up an MCT in a synod (even less for a subscale MCT). So if you can pick up and drop some weight, you can mine with it. If it takes 15 minutes to move 100kg of soil, that's fine. That would, of course, be a joke on Earth. And yeah, you could outperform it with a hand shovel. But that's irrelevant.
You still haven't said how you can use a crane to do mining. (Let alone how this could be automated.)
...Not true. An axi-symmetrical "Y" shape with 4 engines would allow it.
4 engines precludes a center engine landing...
Sure, but that takes up as much space as 7 engines. I suppose if you're pointing a central engine in the direction of reentry, you want a very wide, very light stage anyway. Optimize for reentry ballistic coefficient, and just live with the drag on the way up....Not true. An axi-symmetrical "Y" shape with 4 engines would allow it.
4 engines precludes a center engine landing...
The nozzle extensions of vac engines will need protection to survive reentry at least on earth. So they will probably be retractable for that reason. Can the remaining nozzle part formed like first stage engines and work as such in retracted state?
......for a LOT of missions, success isn't entirely binary. I can point to several times when RL-10-based upper stages had a significantly early shutdown for one reason or another but the performance was still made up for, perhaps by the payload. This may lead to lower on-orbit fuel, which is annoying and can shorten the life of the payload, but this isn't LoM.
Also of interest: in a vacuum it gets 92 to 96 % of the performance fully expanded vac nozzle. If the mission requires that last 4-8% of performance, then an extension failure would cause LoM.
On the end of the crane, you'd have a clamshell bucket or orange-peel grabs or something like a blend between the two.
and could be enhanced by dragging it with another cable
I thought you were referring to a dragline. Anyway, I've seen these mechanisms in action. Very little finesse is necessary to grab material if you're not concerned with exactly what spot of dirt you're grabbing or if you're getting a whole bunch of dirt or not. This is not an especially hard problem.On the end of the crane, you'd have a clamshell bucket or orange-peel grabs or something like a blend between the two.
I said a bucket on a cable requires human finesse to control, something that can't be automated yet (certainly not "easily"). You specifically said you weren't talking about that....
...I WASN'T talking about a dragline, I was simply acknowledging there that you COULD use one.and could be enhanced by dragging it with another cable
So now you're describing a dragline, which you specifically said you weren't.
You're shocked at my confusion when you completely contradict yourself?
I've operated a dragline[1]. It's pretty hard to get a good load and raise it successfully. Especially repeatably. So I think it's not exactly something I'd call "very little finesse".A) Getting a good load is not at all a requirement. We're talking about a very small amount of material, relatively speaking.
1 - It was a LEGO model, but it was built by someone whose job it is to spec draglines for coal mines and he said the real thing operates the same way. Very finicky in operation, and taking a very deft hand on the controls with situational awareness of exactly what kinds of lumps and irregularities face you.
A lot will depend on the condition of the soil/regolith. If it is solidly frozen, rather like permafrost, then shovels and draglines will not be much use, it'll be too tough to penetrate. If it's loose with boulders or large 'rocks' of sand and ice, then it's a lot simpler. If it's a cliff face, then it's relatively easy to bore into. A few kg of well placed explosives might do wonders in loosening up the soil. That's a well known procedure!I've operated a dragline[1]. It's pretty hard to get a good load and raise it successfully. Especially repeatably. So I think it's not exactly something I'd call "very little finesse".A) Getting a good load is not at all a requirement. We're talking about a very small amount of material, relatively speaking.
1 - It was a LEGO model, but it was built by someone whose job it is to spec draglines for coal mines and he said the real thing operates the same way. Very finicky in operation, and taking a very deft hand on the controls with situational awareness of exactly what kinds of lumps and irregularities face you.
B) The dragline comment was an aside, an acknowledgement of a possible improvement. Not my actual point. I was talking about this: https://en.wikipedia.org/wiki/Bucket_(machine_part)#/media/File:Coal_loading_shell_grabs,_Cardiff.jpg
Having some trouble with the notion that you can do this "inefficiently" and actually get any useful material before you wear out your equipment.Go back and read the orders of magnitude point I made. We can get just 1/1000th the material that you'd get with typical mining equipment and still have sufficient material. Our intuitions are screaming about stuff that is not relevant at this scale.
Those coal loader scoops only work when dropped pretty much straight down. So now you need a boom of some sort to position the scoop where you want it (which you did with a dragline too).Yeah, that's what the crane is for.
Why are we arguing about this? And why here? I've lost how this is related to the MCT per se. It's related to how to do ISRU, for sure but not MCT design...Because I threw out the fact that something simple, even something as dumb as a crane, could be used instead of a rover. And the throw of a crane would be enough that it wouldn't even need to move its base to capture enough material. It was a throwaway comment that someone couldn't resist challenging with their intuitions honed on Earth but misled because the intuition is trained with orders of magnitude more throughput requirements.
......for a LOT of missions, success isn't entirely binary. I can point to several times when RL-10-based upper stages had a significantly early shutdown for one reason or another but the performance was still made up for, perhaps by the payload. This may lead to lower on-orbit fuel, which is annoying and can shorten the life of the payload, but this isn't LoM.
Also of interest: in a vacuum it gets 92 to 96 % of the performance fully expanded vac nozzle. If the mission requires that last 4-8% of performance, then an extension failure would cause LoM.
Very site specific is my guess.
a few more teasers before the september reveal:Bold mine.
https://www.washingtonpost.com/news/the-switch/wp/2016/06/10/elon-musk-provides-new-details-on-his-mind-blowing-mission-to-mars/
1 red dragon in 2018, 'at least 2' in 2020, then first flight of MCT in 2022...
a few more teasers before the september reveal:Bold mine.
https://www.washingtonpost.com/news/the-switch/wp/2016/06/10/elon-musk-provides-new-details-on-his-mind-blowing-mission-to-mars/
1 red dragon in 2018, 'at least 2' in 2020, then first flight of MCT in 2022...
This thread hasn't been updated for a while (so there may be some public details that have been missed), but here's something new:a few more teasers before the september reveal:
https://www.washingtonpost.com/news/the-switch/wp/2016/06/10/elon-musk-provides-new-details-on-his-mind-blowing-mission-to-mars/
1 red dragon in 2018, 'at least 2' in 2020, then first flight of MCT in 2022...
“the first mission wouldn’t have a huge number of people on it because if something goes wrong, we want to risk the fewest number of lives as possible.”
a few more teasers before the september reveal:Bold mine.
https://www.washingtonpost.com/news/the-switch/wp/2016/06/10/elon-musk-provides-new-details-on-his-mind-blowing-mission-to-mars/
1 red dragon in 2018, 'at least 2' in 2020, then first flight of MCT in 2022...
That is first fight of MCT to Mars in 2022, which likely means that BFR/BFS first flight would need to be in 2021.
Says to me, first launch of MCT in 2022, NOT specifically TO Mars, just launch.
a few more teasers before the september reveal:Bold mine.
https://www.washingtonpost.com/news/the-switch/wp/2016/06/10/elon-musk-provides-new-details-on-his-mind-blowing-mission-to-mars/
1 red dragon in 2018, 'at least 2' in 2020, then first flight of MCT in 2022...
That is first fight of MCT to Mars in 2022, which likely means that BFR/BFS first flight would need to be in 2021.
"Then in 2022, Musk said he hoped to launch what the company now sometimes refers to as the Mars Colonial Transporter, designed to bring a colony to Mars."
Says to me, first launch of MCT in 2022, NOT specifically TO Mars, just launch.
Mars EDL is different enough from Earth EDL that you're going to have to do a test of MCT on Mars before you send people.
Maybe they'll send 2 BFSes the first time, one to return and one to stay.Mars EDL is different enough from Earth EDL that you're going to have to do a test of MCT on Mars before you send people.
Well an landed and not refueled crewless BFS / MCT would make an excellent hab and start of a Mars base.
Imagine the workshops, repair and other facilities, heavy earth movers, cranes, rovers, etc that could be built into and delivered by a crewless BFS that never returned to Earth but became SpX Mars Base 1.
Maybe they'll send 2 BFSes the first time, one to return and one to stay.Mars EDL is different enough from Earth EDL that you're going to have to do a test of MCT on Mars before you send people.
Well an landed and not refueled crewless BFS / MCT would make an excellent hab and start of a Mars base.
Imagine the workshops, repair and other facilities, heavy earth movers, cranes, rovers, etc that could be built into and delivered by a crewless BFS that never returned to Earth but became SpX Mars Base 1.
Launch 1st BFS into LEO, do heaps of checks,If this is the 1st BFS you don't have anything to put in a bit of fuel.put in a bit of fuel, do a lunar loop& propulsive land on Earth.
Maybe next check flight is to land on Luna and rtn Earth.
No need to wait for Mars synods to do these and other pre TMI burn checkouts of the 1st BFS.
Launch 1st BFS into LEO, do heaps of checks,If this is the 1st BFS you don't have anything to put in a bit of fuel.put in a bit of fuel, do a lunar loop& propulsive land on Earth.
Maybe next check flight is to land on Luna and rtn Earth.
No need to wait for Mars synods to do these and other pre TMI burn checkouts of the 1st BFS.
Perhaps Mission 2a & 2b:
Relauch BFS 1; Launch BFS 2
BFS 1 refuels BFS 2, or vice versa.
BFS 1 lands; BFS 2 goes to moon, possibly landing if fuel levels permit return.
BFS 2 returns and lands.
But certainly agree this is not tied to Mars synods, other than to be done beforehand.
Launch 1st BFS into LEO, do heaps of checks,If this is the 1st BFS you don't have anything to put in a bit of fuel.put in a bit of fuel, do a lunar loop& propulsive land on Earth.
Maybe next check flight is to land on Luna and rtn Earth.
No need to wait for Mars synods to do these and other pre TMI burn checkouts of the 1st BFS.
Perhaps Mission 2a & 2b:
Relauch BFS 1; Launch BFS 2
BFS 1 refuels BFS 2, or vice versa.
BFS 1 lands; BFS 2 goes to moon, possibly landing if fuel levels permit return.
BFS 2 returns and lands.
But certainly agree this is not tied to Mars synods, other than to be done beforehand.
BFR, without a BFS load can put a lot of fuel into LEO. So launch BFS on BFR which lands, is refueled and launch into LEO, then transfers fuel to BFS for moon loop around or later Luna landing and return (will need to be a very light BFS or maybe put a BFR tanker in LLO)
For sure SpX will have solved how to maintain methlox in space for a very long time or Mars landing of BFS using methlox Raptors will never happen.
It will need two orbiting vehicles. One that gets refuelled and one that does the fuelling run. An ability that needs to be tested early. The tanker may not qualify as MCT.
I'm extremely curious about what kind of site selection process they'll be going through.
Musk quoted in the Washington Times article (https://www.washingtonpost.com/news/the-switch/wp/2016/06/10/elon-musk-provides-new-details-on-his-mind-blowing-mission-to-mars/):
"At another point he said, “I’m so tempted to talk more about the details of it. But I have to restrain myself.”"
First a bit of irony is that Musk is talking about his plans in a publication that Jeff Bezos owns.
What caught my eye in the article also was this:
"Then in 2022, Musk said he hoped to launch what the company now sometimes refers to as the Mars Colonial Transporter, designed to bring a colony to Mars."
Being a product scheduling professional, my mind started backing off that date all of the physical tasks that needed to happen in just 6 short years. Including:
- Launch and manufacturing site selection, development, and building and staffing of the factory (assumed to be co-located).
- Production of the first test vehicle, which may not go to Mars but is used only for test purposes.
- Production and testing of the first vehicle to go to Mars in 2022, with test flights to validate it's ready.
- Sources of funding, i.e. money, and lots of it. Sure SpaceX and Musk will be supplying some part of that, but I would imagine that Musk is working on getting BOMC's (Believers Of Mars Colonization) to contribute too.
And if we're talking MCT, we have to assume the BFR will also be under development. That is a lot. I would imagine we'll get a sense of the schedule of events when Musk unveils more details in September, and as he says in the article:
"And he acknowledged that the company would have to “get lucky and things go according to plan” to hit a launch window for manned flight in late 2024, with a landing in 2025."
I'm already crossing my fingers for luck...
Can't wait until September when we find out what is really going on.
When I enumerate the myriad serious steps and when I look at the cash needed, I am very skeptical of Musk's admitted "everything goes right" schedule. But unlike myself as a former VP Engineering/CTO, he won't fire himself when he seriously misses the best case schedule.
When you own the whole thing you can set best case schedules and not suffer the consequences of missing them....unless you run out of cash! I also have an MBA in finance and do worry about sufficient cash to fund all this.
On top of everything SX is doing, they gotta fund BFR/BFS development, a BFR/BFS factory, a launch complex, a likely off shore launch complex... the list goes on.
As someone who had to nickel & dime product cost as well as product performance to be competitive, I have industry experience based respect for Elon's product cost management and life cycle cost management. The guy is a true polymath.
Bottom Line: even if he's years late as I believe he will be, it is civilization changing.
Good health, Elon!
It will need two orbiting vehicles. One that gets refuelled and one that does the fuelling run. An ability that needs to be tested early. The tanker may not qualify as MCT.
Well, util the recent article the MCT term had seemed to be in retirement, replaced by the BFR/BFS terminology, and purposefully or not Elon seems to keep the terms 'muddy'.
My point was/is that the first space vehicle boosted to orbit by a raptor powered first stage is highly unlikely to be refueled by anything. It is far more likely that that space vehicle will be the thing expected to do the refueling in the future. If there is a specific tanker version, I would expect it to be the first version launched because none of the others go too far without it. It's also likely to be the simplest version (if in fact there are different versions).
BFS without refueling should able to a free return circumlunar mission with a small payload. No need for a tanker version or even a second spacecraft for that.
I think BFS will only have a payload adapter (passenger hab or cargo hold integrated as needed), so a tanker would only differ in tank size and might not be necessary at all.
BFS would necessarily be able to reach LEO with lots of fuel, just not also with a maximum payload. So the point stands.BFS without refueling should able to a free return circumlunar mission with a small payload. No need for a tanker version or even a second spacecraft for that.
I think BFS will only have a payload adapter (passenger hab or cargo hold integrated as needed), so a tanker would only differ in tank size and might not be necessary at all.
BFS is likely designed to reach LEO with little fuel remaining. Refueling is essential.
This is not one-shot rocketry.
I think a BFS with only main rocket systems(tanks, engines, control, etc),can reach a TMI and land on Mars surface by a single launch.BFS without refueling should able to a free return circumlunar mission with a small payload. No need for a tanker version or even a second spacecraft for that.
I think BFS will only have a payload adapter (passenger hab or cargo hold integrated as needed), so a tanker would only differ in tank size and might not be necessary at all.
BFS is likely designed to reach LEO with little fuel remaining. Refueling is essential.
This is not one-shot rocketry.
BFS would necessarily be able to reach LEO with lots of fuel, just not also with a maximum payload. So the point stands.BFS without refueling should able to a free return circumlunar mission with a small payload. No need for a tanker version or even a second spacecraft for that.
I think BFS will only have a payload adapter (passenger hab or cargo hold integrated as needed), so a tanker would only differ in tank size and might not be necessary at all.
BFS is likely designed to reach LEO with little fuel remaining. Refueling is essential.
This is not one-shot rocketry.
But refueling will be needed early on.
BFS without refueling should able to a free return circumlunar mission with a small payload. No need for a tanker version or even a second spacecraft for that.A BFS with only main rocket systems can reach LLO and back without refueling.But it's not worth.
I think BFS will only have a payload adapter (passenger hab or cargo hold integrated as needed), so a tanker would only differ in tank size and might not be necessary at all.
Dry mass likely to be much less than the payload. ...
As large as its likely to be, and with baked in habitation, could a BFS be considered a lunar "base" if left in place? Guessing it depends on ease of egress/ingress, crew rotation landers and power generation, but still.If Tiangong can be considered a space station, then sure, why the heck not?
I doubt the hab will be baked in, which probably helps. Thermal control during the month long day/night cycle is an issue though.As large as its likely to be, and with baked in habitation, could a BFS be considered a lunar "base" if left in place? Guessing it depends on ease of egress/ingress, crew rotation landers and power generation, but still.If Tiangong can be considered a space station, then sure, why the heck not?
According to Musk, it will be.I doubt the hab will be baked in, which probably helps. Thermal control during the month long day/night cycle is an issue though.As large as its likely to be, and with baked in habitation, could a BFS be considered a lunar "base" if left in place? Guessing it depends on ease of egress/ingress, crew rotation landers and power generation, but still.If Tiangong can be considered a space station, then sure, why the heck not?
Musk quoted in the Washington Times article (https://www.washingtonpost.com/news/the-switch/wp/2016/06/10/elon-musk-provides-new-details-on-his-mind-blowing-mission-to-mars/):
"Being a product scheduling professional, my mind started backing off that date all of the physical tasks that needed to happen in just 6 short years. …...
And if we're talking MCT, we have to assume the BFR will also be under development. That is a lot. I would imagine we'll get a sense of the schedule of events when Musk unveils more details in September, and as he says in the article:
Break ground on launch site by 2018/2019.
I think that lots of guys have shared your skepticism about the timeline. But do you think that slippage of one synod would not be OK for this grand plan? Meaning, that he's shooting for the 2022 synod but expects the next one to be achievable.
From a leadership standpoint, it seems that this plan might be very effective to motivate the "true believers" to put in their 100%. He's got a real strong cadre with world-class skills, but they need a clear goal. If he sets a goal for 2022, they'll bust their a**es to make it happen. If that comes to pass, GREAT, but if it slips for whatever technical reasons, won't this still be a winning situation? All of them will understand why thy didn't *quite* make it for that synod, and they'll be stoked to make it happen for the next one.
Newer refined model for BFR/BFS. Spreadsheet attached.
BFR & BFS MODELS: MCT as 2nd Stage with 100 Metric Tons CARGO
BLUE: Enter parameter variable
S1 Avg ISP Sea L to MECO 335
S2 vac ISP 380
Raptor sea level thrust KLB & mT 518 235
Rvac thrust KLB & mT 559 254
S1 Dry Wt % 4.5%
BFR 1
BFR DIA 15.0 m
MCT BFS Dry Wt & Cargo 225 mT
S2 Dry Mass 125 mT
Propellant for S2 BFS Landing 30 mT
Total Mass to LEO 255 mT
1st Stage Propellant Tank Length 18.0 m
S1 Propellant Volume 3179 m3
Propellant Mass 3371 mT
S1 Dry Wt % 4.5%
S1 DRY Weight 159 mT
S1 Total Weight mT 3529 mT
S! Dry Wt Delta V (No 2nd stage) 10.2 Km/sec Rocket Equation
Stage One Full Load Delta V 3.35 Km/sec Rocket Equation
RTxx Propellant 70 mT
RTxx Delta V @Minimum Load 1.20 Km/sec Rocket Equation
Est S1 Gravity Loss 0.9 Km/sec
Est S1 Velocity @ Burnout 2.45 Km/sec
2nd Stage Propellant Tank Length 7.5 m
Propellant Volume 1,325 m3
Propellant Mass 1,404 mT
S2 Mass w/MCT 1,629 mT
S2 Mass w/MCT 3.6 Million LBS
Calc # Rvac Raptor Eng 5.01 0.78
Stage 2 Thrust mT 1271
Stage 2 Thrust 2.8 Million LBS
Stage 2 Km/sec 6.91 Km/sec Rocket Equation
Stage 2 Wet to Dry Mass Ratio 12.2
S1 + S2 Total Delta V 9.4 Km/sec
TOTAL WT mT 5,159 mT
TOTAL WT LBS 11.4 Million LBS
THRUST Needed LBS 14.0 Million LBS
THRUST Needed mT 6345
THRUST Needed MegaNewtons 62
1st Stage T/W @ Takeoff 1.23
1st STAGE # ENG 27
LEO Mass Fract 4.4% %
LEO Wet to Dry Mass Ratio 23 F9 v1.1 25/1 Musk
MCT Cargo Hold length 10 m
MCT Cargo Vol m3 1766 m3
Eng 16+8+3=27
Outer Ring, Inner Ring, Central Engs
NOTE 1: S1 Km/sec + S2 Km/sec must ~9Km/sec for LEO with grav losses
NOTE 2: Rocket Equation
According to Musk, it will be.I doubt the hab will be baked in, which probably helps. Thermal control during the month long day/night cycle is an issue though.As large as its likely to be, and with baked in habitation, could a BFS be considered a lunar "base" if left in place? Guessing it depends on ease of egress/ingress, crew rotation landers and power generation, but still.If Tiangong can be considered a space station, then sure, why the heck not?
We know that return payload is 25 tonnes. If the hab were integrated into the BFS (and it is still counted as payload) then it is likely yo be too heavy to be returned. But if it isn't returned then how can crew return.
Return from Mars is a strong constraint on the architecture, there are lots of potential architectures that can "land the whole thing" with 100 tonnes of payload, but very few that can return 25 tonnes within one synod. About the only one I can see working is a modular habitat of 25 tonnes (wet fully provisioned), able to carry 25 crew, which may be left on Mars or returned within 1-4 BFS.
We know that return payload is 25 tonnes. If the hab were integrated into the BFS (and it is still counted as payload) then it is likely yo be too heavy to be returned. But if it isn't returned then how can crew return.
Return from Mars is a strong constraint on the architecture, there are lots of potential architectures that can "land the whole thing" with 100 tonnes of payload, but very few that can return 25 tonnes within one synod. About the only one I can see working is a modular habitat of 25 tonnes (wet fully provisioned), able to carry 25 crew, which may be left on Mars or returned within 1-4 BFS.
If we go back to the Max Fagin presentation only a fraction of the BFS returns to Earth, leaving (for lack of a better term) a large logistics module on Mars. Fagin shows an "über-Dragon" style configuration with integrated sidewall engines. If these are only enough get it to LMO (easier on the ISRU needs) a pre-positioned depot could provide the deltaV home. The interstage between "über-Dragon" and the logistics module protects the formers heat shield inbound.
IP protections don't cover obviousness. See rocket landings on barges. See here for another split vehicle config,
We know that return payload is 25 tonnes. If the hab were integrated into the BFS (and it is still counted as payload) then it is likely yo be too heavy to be returned. But if it isn't returned then how can crew return.
Return from Mars is a strong constraint on the architecture, there are lots of potential architectures that can "land the whole thing" with 100 tonnes of payload, but very few that can return 25 tonnes within one synod. About the only one I can see working is a modular habitat of 25 tonnes (wet fully provisioned), able to carry 25 crew, which may be left on Mars or returned within 1-4 BFS.
If we go back to the Max Fagin presentation only a fraction of the BFS returns to Earth, leaving (for lack of a better term) a large logistics module on Mars. Fagin shows an "über-Dragon" style configuration with integrated sidewall engines. If these are only enough get it to LMO (easier on the ISRU needs) a pre-positioned depot could provide the deltaV home. The interstage between "über-Dragon" and the logistics module protects the formers heat shield inbound.
It is very unlikely that the Max Fagin presentation is the MCT architecture, he would have been prevented by non-disclosure agreement from using SpaceX IPR.
Pretty much what I think the the BFS/MCT will be, with a few minor changes:
1. larger (longer) cargo hold.
2. slightly better mass fractions (but maybe not on the first MCT).
3. slightly less delta-v required to reach orbit, difference from your figures gives margin.
philw1776: It seems that your design would require 14 refueling flights to be ready for TMI, and then at Mars a staggering amount of propellant to return. I'm incredibly doubtful of these fast LEO departures and direct returns because the launch count necessary to do a mission will run up costs and the Mars surface refueling will stress ISPP too far.
Also the vehicle dimensions seem incredibly squat, with tanks that are nearly hockey-pucks in shape, whats the total stack height at launch, it seems like it would be shorter then F9 given the numbers your providing. I don't see the motivation for such squatness unless you believe Raptor has terrible thrust density, but Russian staged combustion hydrocarbon engines (our best analogs for Raptor) have great thrust density which should easily support a vehicle of 80-100 m of height at liftoff.
In line with the philosophy of reuse mfck articulated...
I could see the MCT being "partially strippable"... if you had cabins for 100, but are only returning 20 people why carry all of that back? if the fittings, panels, wiring, plumbing etc was designed to be modular and reusable it could be removed by humans and used in outfitting the base.
Compression requires much too much work. It should be fairly simple to use dessicants to dry out the air, then extract the water by heating the dessicant. This is one of the most common technologies today to dry compressed air or indoor ice rinks (and was part of the original Mars direct proposal). alternatively you can compress a refrigerant, and use that to create a cold surface on which the water in the air will condense out. Any of these solutions will require a minimum of about 1000 btu/lb or 2300 kJ/kg of water, the phase change energy of water.Remember Mars has a atmosphere and has water cycles (water vapor and water ice, does not have liquid phase). So to mine all the water you need, you compress the air and liquefy the water vapor in the atmosphere. Admittedly you would have to compress a lot of atmosphere for a little bit of water but the advantage is you do not have to go out and dig it up! Once the quantities needed become very high then comes the mining of water. But for the initial methalox manufacture for the return trip this method would be an easy one to implement and would not require complex equipment just plenty of power. Compressing gas takes a great deal of energy. But you need the atmosphere compressed for other parts of the methalox manufacture as well.
Not really possible to get water by compression. Water vapour is about 0.03%, so for every kg of water you would need to compress and liquefy 3 tonnes ofwater<air>.
Your dry mass is too high.philw1776: It seems that your design would require 14 refueling flights to be ready for TMI, and then at Mars a staggering amount of propellant to return. I'm incredibly doubtful of these fast LEO departures and direct returns because the launch count necessary to do a mission will run up costs and the Mars surface refueling will stress ISPP too far.
Also the vehicle dimensions seem incredibly squat, with tanks that are nearly hockey-pucks in shape, whats the total stack height at launch, it seems like it would be shorter then F9 given the numbers your providing. I don't see the motivation for such squatness unless you believe Raptor has terrible thrust density, but Russian staged combustion hydrocarbon engines (our best analogs for Raptor) have great thrust density which should easily support a vehicle of 80-100 m of height at liftoff.
2 good observations.
First, the tanks. You're on target. What I did was simply compute volume and weight of cylinders to estimate the mass of the rocket, etc. I do NOT mean that the propellant tank is really18m for example, My bad. There is a big O2 tank and a separate methane tank. They have rounded ends in reality, making them longer. A 15m wide rocket does not really have a 15m wide tank. Again the simplification is used to estimate mass, thrust, etc. and not length of rocket which however I believe will still be squat under 100m. Engines, interstage, whatever. I needed to add prose to be clear on that.
Mike A observed that there is excess Km/sec capacity in that the craft can arrive in LEO with extra tons of fuel. I believe there will be an upper stage BFS configuration used as a fuel truck with less dry mass. I get 8-10 refueling trips for one 120 day or less transit. I think that's too many refueling trips.
Rest assured SX has a much better solution as I'm just a EE and not an aerospace engineer.
One final point. This is a brute force all chemical approach. SX will be more imaginative.
Mike A observed that there is excess Km/sec capacity in that the craft can arrive in LEO with extra tons of fuel. I believe there will be an upper stage BFS configuration used as a fuel truck with less dry mass. I get 8-10 refueling trips for one 120 day or less transit. I think that's too many refueling trips.
Rest assured SX has a much better solution as I'm just a EE and not an aerospace engineer.
One final point. This is a brute force all chemical approach. SX will be more imaginative.
Your dry mass is too high.philw1776: It seems that your design would require 14 refueling flights to be ready for TMI, and then at Mars a staggering amount of propellant to return. I'm incredibly doubtful of these fast LEO departures and direct returns because the launch count necessary to do a mission will run up costs and the Mars surface refueling will stress ISPP too far.
Also the vehicle dimensions seem incredibly squat, with tanks that are nearly hockey-pucks in shape, whats the total stack height at launch, it seems like it would be shorter then F9 given the numbers your providing. I don't see the motivation for such squatness unless you believe Raptor has terrible thrust density, but Russian staged combustion hydrocarbon engines (our best analogs for Raptor) have great thrust density which should easily support a vehicle of 80-100 m of height at liftoff.
2 good observations.
First, the tanks. You're on target. What I did was simply compute volume and weight of cylinders to estimate the mass of the rocket, etc. I do NOT mean that the propellant tank is really18m for example, My bad. There is a big O2 tank and a separate methane tank. They have rounded ends in reality, making them longer. A 15m wide rocket does not really have a 15m wide tank. Again the simplification is used to estimate mass, thrust, etc. and not length of rocket which however I believe will still be squat under 100m. Engines, interstage, whatever. I needed to add prose to be clear on that.
Mike A observed that there is excess Km/sec capacity in that the craft can arrive in LEO with extra tons of fuel. I believe there will be an upper stage BFS configuration used as a fuel truck with less dry mass. I get 8-10 refueling trips for one 120 day or less transit. I think that's too many refueling trips.
Rest assured SX has a much better solution as I'm just a EE and not an aerospace engineer.
One final point. This is a brute force all chemical approach. SX will be more imaginative.
a few more teasers before the september reveal:Bold mine.
https://www.washingtonpost.com/news/the-switch/wp/2016/06/10/elon-musk-provides-new-details-on-his-mind-blowing-mission-to-mars/
1 red dragon in 2018, 'at least 2' in 2020, then first flight of MCT in 2022...
Then in 2022, Musk said he hoped to launch what the company now sometimes refers to as the Mars Colonial Transporter, designed to bring a colony to Mars.
We don't know the development status of MCT. It could be anywhere from a few powerpoints which seem to hang together as an architecture, to having passed PDR (or the SpaceX equivalent) several months ago. About the only thing known in public is that the Raptor has had component level tests which have gone quite well.
It is now 8 months since Chris Bergin made that tweet https://forum.nasaspaceflight.com/index.php?topic=38593.0 and things will have moved on a lot since then. Chris has shared some of the information that he received in L2, I cannot say what that is, but just given the fact that things were far enough advanced for Chris to be shown data means that they were far enough advanced for the basic factory and launch site specs to be determined (not the detailed ones, but such things as floor area, overhead crane height, access requirements, thrust levels and landing pad requirements). These are enough to start looking for a factory and launch site, and if they have been at it for 8 months SpaceX probably have a pretty good idea about the possibilities.
And so you have 14 refuelings. That's not going to happen. That is definitely not what SpaceX is planning. Your dry masses are too high.Your dry mass is too high.philw1776: It seems that your design would require 14 refueling flights to be ready for TMI, and then at Mars a staggering amount of propellant to return. I'm incredibly doubtful of these fast LEO departures and direct returns because the launch count necessary to do a mission will run up costs and the Mars surface refueling will stress ISPP too far.
Also the vehicle dimensions seem incredibly squat, with tanks that are nearly hockey-pucks in shape, whats the total stack height at launch, it seems like it would be shorter then F9 given the numbers your providing. I don't see the motivation for such squatness unless you believe Raptor has terrible thrust density, but Russian staged combustion hydrocarbon engines (our best analogs for Raptor) have great thrust density which should easily support a vehicle of 80-100 m of height at liftoff.
2 good observations.
First, the tanks. You're on target. What I did was simply compute volume and weight of cylinders to estimate the mass of the rocket, etc. I do NOT mean that the propellant tank is really18m for example, My bad. There is a big O2 tank and a separate methane tank. They have rounded ends in reality, making them longer. A 15m wide rocket does not really have a 15m wide tank. Again the simplification is used to estimate mass, thrust, etc. and not length of rocket which however I believe will still be squat under 100m. Engines, interstage, whatever. I needed to add prose to be clear on that.
Mike A observed that there is excess Km/sec capacity in that the craft can arrive in LEO with extra tons of fuel. I believe there will be an upper stage BFS configuration used as a fuel truck with less dry mass. I get 8-10 refueling trips for one 120 day or less transit. I think that's too many refueling trips.
Rest assured SX has a much better solution as I'm just a EE and not an aerospace engineer.
One final point. This is a brute force all chemical approach. SX will be more imaginative.
4.5% for 1st stage may be high. I'm in the camp of those who say minimum 6 legs. I hope it's high, but I think the BFR has to be more robust, read heavier, if it really is a quick turn around, only very minor refurbishment vehicle. I think today's F9 has unresolved issues there.
As to the BFS, with all the exotica of engines placed high for Mars landing & takeoff, cargo arrangement complications and robust TPS that lasts many re-entries at interplanetary velocities, I don't buy the dry mass under 100mT thinking for such a complex, lightly serviced vehicle.
We don't know the development status of MCT. It could be anywhere from a few powerpoints which seem to hang together as an architecture, to having passed PDR (or the SpaceX equivalent) several months ago. About the only thing known in public is that the Raptor has had component level tests which have gone quite well.
It is now 8 months since Chris Bergin made that tweet https://forum.nasaspaceflight.com/index.php?topic=38593.0 and things will have moved on a lot since then. Chris has shared some of the information that he received in L2, I cannot say what that is, but just given the fact that things were far enough advanced for Chris to be shown data means that they were far enough advanced for the basic factory and launch site specs to be determined (not the detailed ones, but such things as floor area, overhead crane height, access requirements, thrust levels and landing pad requirements). These are enough to start looking for a factory and launch site, and if they have been at it for 8 months SpaceX probably have a pretty good idea about the possibilities.
And to tie it together to other "signs and indicators," for those who have been following the "Where will BFR launch from?" thread, if you look at SpaceX planning on building and launching BFRs in six years, the only place where they are currently beginning construction on new facilities is Boca Chica.
If y'all are saying that SpaceX needs to be building the BFR factory and launch facilities right now, well -- maybe they are.
Yeah, that's certainly a possibility. Boca Chica isn't perfect for daily launches, but for the first decade or two, when the launch rate is more modest, it could certainly function as the first BFR launch site. SpaceX seems to not have a problem with multiple launch sites nor with solutions that only are going to work for a few years or a decade or so.We don't know the development status of MCT. It could be anywhere from a few powerpoints which seem to hang together as an architecture, to having passed PDR (or the SpaceX equivalent) several months ago. About the only thing known in public is that the Raptor has had component level tests which have gone quite well.
It is now 8 months since Chris Bergin made that tweet https://forum.nasaspaceflight.com/index.php?topic=38593.0 and things will have moved on a lot since then. Chris has shared some of the information that he received in L2, I cannot say what that is, but just given the fact that things were far enough advanced for Chris to be shown data means that they were far enough advanced for the basic factory and launch site specs to be determined (not the detailed ones, but such things as floor area, overhead crane height, access requirements, thrust levels and landing pad requirements). These are enough to start looking for a factory and launch site, and if they have been at it for 8 months SpaceX probably have a pretty good idea about the possibilities.
And to tie it together to other "signs and indicators," for those who have been following the "Where will BFR launch from?" thread, if you look at SpaceX planning on building and launching BFRs in six years, the only place where they are currently beginning construction on new facilities is Boca Chica.
If y'all are saying that SpaceX needs to be building the BFR factory and launch facilities right now, well -- maybe they are.
Boca Chica doesn't have to be the manufacturing site, it just has to be nearby so that transporting a large structure isn't too hard.We don't know the development status of MCT. It could be anywhere from a few powerpoints which seem to hang together as an architecture, to having passed PDR (or the SpaceX equivalent) several months ago. About the only thing known in public is that the Raptor has had component level tests which have gone quite well.
It is now 8 months since Chris Bergin made that tweet https://forum.nasaspaceflight.com/index.php?topic=38593.0 and things will have moved on a lot since then. Chris has shared some of the information that he received in L2, I cannot say what that is, but just given the fact that things were far enough advanced for Chris to be shown data means that they were far enough advanced for the basic factory and launch site specs to be determined (not the detailed ones, but such things as floor area, overhead crane height, access requirements, thrust levels and landing pad requirements). These are enough to start looking for a factory and launch site, and if they have been at it for 8 months SpaceX probably have a pretty good idea about the possibilities.
And to tie it together to other "signs and indicators," for those who have been following the "Where will BFR launch from?" thread, if you look at SpaceX planning on building and launching BFRs in six years, the only place where they are currently beginning construction on new facilities is Boca Chica.
If y'all are saying that SpaceX needs to be building the BFR factory and launch facilities right now, well -- maybe they are.
In my opinion Boca Chica is not suitable for the MCT manufacturing site. Other places in the Brownsville area maybe. That we have not seen an environmental impact statement is perhaps an indication that SpaceX have found a site that does not need one (existing large factory or facility?) or that they think an environmental impact statement will be a formality (contaminated land, brownfield site?).
If Musk thinks 2024 MCT is possible, then he must see a way forward, but I have no clue about what that path is.
Also, the "local manufacturing site" is likely going to be at least partially an assembly site. They likely won't want to truck BFR stages around -- too large -- but I bet the engines and whatever serves as an octaweb (the thrust and plumbing structures) could still be made at one primary site, like Hawthorne, and shipped out to the tank manufacture/stage assembly sites.
It's not like there would be no manufacturing happening near the BFR launch sites, but the only things that, it would seem, are required to be built near the launch site are the tanks/stage structures. A lot of pieces will be sub-assemblies that are manufactured elsewhere and shipped to the BFR sites in by conventional means...
...could happen in late 2010s, too. I've suspected they'd start with a Raptor upper stage before MCT. I'm unsure if they will or not, but it's certainly a possibility, since a Raptor upper stage for Falcon 9 and Heavy is mentioned in the USAF contract for Raptor.Also, the "local manufacturing site" is likely going to be at least partially an assembly site. They likely won't want to truck BFR stages around -- too large -- but I bet the engines and whatever serves as an octaweb (the thrust and plumbing structures) could still be made at one primary site, like Hawthorne, and shipped out to the tank manufacture/stage assembly sites.
It's not like there would be no manufacturing happening near the BFR launch sites, but the only things that, it would seem, are required to be built near the launch site are the tanks/stage structures. A lot of pieces will be sub-assemblies that are manufactured elsewhere and shipped to the BFR sites in by conventional means...
Possibly, but Hawthorne will be at or near capacity with F9/FH, as reusability reduces the need for first stages, an increased flight rate would increase second stage production*. It is much easier to build on a greenfield site than trying to cram production for completely different (and bigger elements). That said the avionics is probably going to be similar so could be produced at Hawthorne with little difficulty.
[*] in future a reusable raptor based second stage might reduce Hawthorne production requirements, but probably not until the early 2020's.
And so you have 14 refuelings. That's not going to happen. That is definitely not what SpaceX is planning. Your dry masses are too high.Your dry mass is too high.philw1776: It seems that your design would require 14 refueling flights to be ready for TMI, and then at Mars a staggering amount of propellant to return. I'm incredibly doubtful of these fast LEO departures and direct returns because the launch count necessary to do a mission will run up costs and the Mars surface refueling will stress ISPP too far.
Also the vehicle dimensions seem incredibly squat, with tanks that are nearly hockey-pucks in shape, whats the total stack height at launch, it seems like it would be shorter then F9 given the numbers your providing. I don't see the motivation for such squatness unless you believe Raptor has terrible thrust density, but Russian staged combustion hydrocarbon engines (our best analogs for Raptor) have great thrust density which should easily support a vehicle of 80-100 m of height at liftoff.
2 good observations.
First, the tanks. You're on target. What I did was simply compute volume and weight of cylinders to estimate the mass of the rocket, etc. I do NOT mean that the propellant tank is really18m for example, My bad. There is a big O2 tank and a separate methane tank. They have rounded ends in reality, making them longer. A 15m wide rocket does not really have a 15m wide tank. Again the simplification is used to estimate mass, thrust, etc. and not length of rocket which however I believe will still be squat under 100m. Engines, interstage, whatever. I needed to add prose to be clear on that.
Mike A observed that there is excess Km/sec capacity in that the craft can arrive in LEO with extra tons of fuel. I believe there will be an upper stage BFS configuration used as a fuel truck with less dry mass. I get 8-10 refueling trips for one 120 day or less transit. I think that's too many refueling trips.
Rest assured SX has a much better solution as I'm just a EE and not an aerospace engineer.
One final point. This is a brute force all chemical approach. SX will be more imaginative.
4.5% for 1st stage may be high. I'm in the camp of those who say minimum 6 legs. I hope it's high, but I think the BFR has to be more robust, read heavier, if it really is a quick turn around, only very minor refurbishment vehicle. I think today's F9 has unresolved issues there.
As to the BFS, with all the exotica of engines placed high for Mars landing & takeoff, cargo arrangement complications and robust TPS that lasts many re-entries at interplanetary velocities, I don't buy the dry mass under 100mT thinking for such a complex, lightly serviced vehicle.
If y'all are saying that SpaceX needs to be building the BFR factory and launch facilities right now, well -- maybe they are.
And so you have 14 refuelings. That's not going to happen. That is definitely not what SpaceX is planning. Your dry masses are too high.Your dry mass is too high.philw1776: It seems that your design would require 14 refueling flights to be ready for TMI, and then at Mars a staggering amount of propellant to return. I'm incredibly doubtful of these fast LEO departures and direct returns because the launch count necessary to do a mission will run up costs and the Mars surface refueling will stress ISPP too far.
Also the vehicle dimensions seem incredibly squat, with tanks that are nearly hockey-pucks in shape, whats the total stack height at launch, it seems like it would be shorter then F9 given the numbers your providing. I don't see the motivation for such squatness unless you believe Raptor has terrible thrust density, but Russian staged combustion hydrocarbon engines (our best analogs for Raptor) have great thrust density which should easily support a vehicle of 80-100 m of height at liftoff.
2 good observations.
First, the tanks. You're on target. What I did was simply compute volume and weight of cylinders to estimate the mass of the rocket, etc. I do NOT mean that the propellant tank is really18m for example, My bad. There is a big O2 tank and a separate methane tank. They have rounded ends in reality, making them longer. A 15m wide rocket does not really have a 15m wide tank. Again the simplification is used to estimate mass, thrust, etc. and not length of rocket which however I believe will still be squat under 100m. Engines, interstage, whatever. I needed to add prose to be clear on that.
Mike A observed that there is excess Km/sec capacity in that the craft can arrive in LEO with extra tons of fuel. I believe there will be an upper stage BFS configuration used as a fuel truck with less dry mass. I get 8-10 refueling trips for one 120 day or less transit. I think that's too many refueling trips.
Rest assured SX has a much better solution as I'm just a EE and not an aerospace engineer.
One final point. This is a brute force all chemical approach. SX will be more imaginative.
4.5% for 1st stage may be high. I'm in the camp of those who say minimum 6 legs. I hope it's high, but I think the BFR has to be more robust, read heavier, if it really is a quick turn around, only very minor refurbishment vehicle. I think today's F9 has unresolved issues there.
As to the BFS, with all the exotica of engines placed high for Mars landing & takeoff, cargo arrangement complications and robust TPS that lasts many re-entries at interplanetary velocities, I don't buy the dry mass under 100mT thinking for such a complex, lightly serviced vehicle.
I don't get 14 refuelings as there is excess capacity in the Km/sec budget which as I responded above to another poster translates into propellant to LEO.
The dry mass listed for the BFS is for the Mars transport vehicle. I assume a lesser dry mass for the stripped down tanker to LEO version. I ran the numbers again and get 6-7 tanker trips which I agree is too many for a cost effective campaign. SX will do better.
Yes, I am saying that SpaceX would have to be building the production facilities and launch site right now, and much much more. I don't think there is any evidence of that, and it would all be much too large to keep secret. The factory will have to have a welding tool that is twice the diameter of the vertical welding facility for SLS, and probably much taller. The launch site will have to be qualified for launching a vehicle with twice the thrust of the Saturn-V. The transporter/tilt-up gantry will have to be the length of a football field. The MCT itself will dwarf anything launched to space except for ISS, how could it possibly be far enough along (in total secrecy) to be ready to launch in six years?
I'm not saying all of this is impossible. I'm saying that I do think it would be impossible for it all to be ready for launch in six years, with no one hearing a peep of any such activity already under way.
I would love to be proven wrong, though. :)
Any attempt to throw 'computing power' and 'manufacturing' into the same improvement over time comparison is insane and shreds your credibility, the differences are astronomical. Concrete cures at the same speed now as it did in 1950, if anything launch pads and related construction is slower now.
Can I get the numbers for your Tanker dry mass and the expected propellant delivery per trip, I don't see how it can get the BFS vehicle full in the 7 trips your proposing as it would need to offload ~200 mt per flight. Between landing propellant reserves and the dry mass of something that is basically just a 2nd stage I don't think you have that much propellant offload capacity, 10 flights seems more reasonable.
All that said I agree even 6-7 refueling flights is not viable, and no amount of Dry mass shaving on RB's part is going to make it work. This is why I belive the only viable architecture is a SEP tug based system in which propellant is pre-placed into high Earth orbit and Mars orbit. The BFS is radically smaller and goes from LEO to EML-2, refuels, goes to Mars, refuels, launches to LMO, refuels and then returns to Earth, all the DeltaV legs are ~4 km/s and allow fast transit times and only requires a vehicle carry around 300 mt of propellant.
So I expect that if the resources and will are there, MCT is **barely** doable to have a first launch in 2022. If you claim it's impossible, you're wrong, it's not impossible. (show your work if you disagree!!). If you claim it's improbable, you're probably right.It is good to discuss the optimal, albeit improbable, case. We can trust Elon that he has a timeline in his mind. I am more troubled with the payload for this launch. Developing surface systems is not considered trivial. Neither mining operaations. Especially troubled if robotic fuel production for MCT return is assumed before the first crewed lauch in '24. You need to understand soil properties at the chosen location to design mining and oxigen extraction. The '18 Dragon lander is said to be only a landing test. Then the Dragons launched in '20 have to study the soil and to do mining and O2 extraction experiments. Then, you have a little more than a single year to develop the equipment. Maybe, the '18 lander should be more than a landing experiment. Maybe, it will include drilling and analysis of the samples. (Some of you say that it would be impossible to develop such payload for '18.) Then, the '20 landers can validate the mining/extraction methods developed on an informed basis. Maybe, the '20 landers include sample return...
a few more teasers before the september reveal:Bold mine.
https://www.washingtonpost.com/news/the-switch/wp/2016/06/10/elon-musk-provides-new-details-on-his-mind-blowing-mission-to-mars/
1 red dragon in 2018, 'at least 2' in 2020, then first flight of MCT in 2022...
From the article:QuoteThen in 2022, Musk said he hoped to launch what the company now sometimes refers to as the Mars Colonial Transporter, designed to bring a colony to Mars.
I'm sorry, but this isnutssomewhat optimistic. You all realize that 2022 is only six years away, right? Regardless of the fact that Dragon v2 hasn't flown yet, and regardless of the fact that FH hasn't flown yet; NOTHING concrete about BFR/MCT has even been released, and Musk is talking about launching one in six years. Six. Years.
Six years to get BFR off the ground, literally. To build a factory on the scale of Michoud (only bigger) for fabrication and assembly of BFR and MCT. To build a huge HIF to handle the 12.5m or 15m cores, or heck even to lease one of the VAB high bays and get it fitted out for BFR. To build all of the ground support infrastructure and ground transportation. To get the entire Raptor engine (not just components) off of the drawing board and into the test stands and validated.
Heck, you guys are still arguing over where the thing will be built and launched from. Do you think that would really be the case if they were going to be rolling off the assembly line in less than six years?
I like SpaceX and they are doing amazing things. But come on. Please apply a little common sense when these kinds of pronouncements are made. Two days after the article was published and three pages on in this thread, I would expect to see some kind of discussion about how that would even be possible.
Cheers!
Your only allowing 900 m/s for all drag and gravity losses which looks to be too low by around 300 m/s as I can't find any vehicle with total losses of less then around 1200 m/s. That would drop your mass in LEO by about 25 mt and put the flight total at an even 8 which is roughly meeting in the middle of our earlier estimates. Alternatively stretching your vehicle will likely make up the difference.
I'm willing to accept 8 as the best estimate for refueling flights needed to perform this all-chemical brute-force mission architecture when performing a fast crew transfer. Do you have an estimate for what could be sent on a slow cargo flight?
P.S. Oops, 300 m/s is what you get from Earth rotation so looks good.
Your only allowing 900 m/s for all drag and gravity losses which looks to be too low by around 300 m/s as I can't find any vehicle with total losses of less then around 1200 m/s. That would drop your mass in LEO by about 25 mt and put the flight total at an even 8 which is roughly meeting in the middle of our earlier estimates. Alternatively stretching your vehicle will likely make up the difference.8 is too many.
I'm willing to accept 8 as the best estimate for refueling flights needed to perform this all-chemical brute-force mission architecture when performing a fast crew transfer. Do you have an estimate for what could be sent on a slow cargo flight?
P.S. Oops, 300 m/s is what you get from Earth rotation so looks good.
1.2-1.5km/s is on the high end for EDL except if you decide to do a braking burn before entry.
Right. And you'll have enough thrust to do a high-thrust landing, which means low gravity losses.1.2-1.5km/s is on the high end for EDL except if you decide to do a braking burn before entry.
That would indicate 30t of fuel is excessive for Earth EDL, as that's 0.85 to 1.0 km/s for a 100t tanker depending on the altitude/ISP. I don't anticipate anything with an orbital re-entry heatshield will do an entry burn (it's rather pointless in Earth's dense atmosphere), and terminal velocity on Earth will be under 200 m/s for a very large, relatively light vehicle.
Right. And you'll have enough thrust to do a high-thrust landing, which means low gravity losses.1.2-1.5km/s is on the high end for EDL except if you decide to do a braking burn before entry.
That would indicate 30t of fuel is excessive for Earth EDL, as that's 0.85 to 1.0 km/s for a 100t tanker depending on the altitude/ISP. I don't anticipate anything with an orbital re-entry heatshield will do an entry burn (it's rather pointless in Earth's dense atmosphere), and terminal velocity on Earth will be under 200 m/s for a very large, relatively light vehicle.
Also, 100t tanker is much too small.
100t dry is too high for a tanker, I think. And maybe try 1000t instead. And greater payload.Right. And you'll have enough thrust to do a high-thrust landing, which means low gravity losses.1.2-1.5km/s is on the high end for EDL except if you decide to do a braking burn before entry.
That would indicate 30t of fuel is excessive for Earth EDL, as that's 0.85 to 1.0 km/s for a 100t tanker depending on the altitude/ISP. I don't anticipate anything with an orbital re-entry heatshield will do an entry burn (it's rather pointless in Earth's dense atmosphere), and terminal velocity on Earth will be under 200 m/s for a very large, relatively light vehicle.
Also, 100t tanker is much too small.
That's 100t dry. Would be 1200t or 1500t at launch.
100t dry is too high for a tanker, I think. And maybe try 1000t instead. And greater payload.For a dedicated tanker which does not share the outer mould line of the BFS then 100 tonnes is too high (I reckon something like 70 tonnes, but SpaceX might be able to reduce that even more). However, the first tanker flights may not be a dedicated design, instead just standard BFS with extra fuel in their tanks and perhaps long duration equipment removed. Then 100 tonnes dry is probably a bit too little.
Phil, if I might ask, why do you have so much performance spec'ed into the upper stage?
A fast (80-120 day) transfer to Mars rarely requires more than 4.5 to 5 km/s from LEO, but your numbers (1604t wet, 225t dry, 380 ISP) give 7.32 km/s of total performance with a 100t payload. Mars EDL will add somewhere in the 1.2 to 1.5 km/s range (I don't see an estimate in your spreadsheet), but even with that requirement your margins run from 12.6% to 28.4%. Since it can do a fast transfer and Mars EDL while only partly fueled, I get 4 to 5.5 refueling launches per fast Mars transfer.
Earth return definitely can use that performance, but you don't include any numbers for that. By my estimates the poor alignments in the 2020's prevent even a 1604t wet, 125t dry vehicle from returning before the next synod's optimal launch window.
I think the initial BFSes will be on the order of 80-120 tons dry, even with crew equipment.100t dry is too high for a tanker, I think. And maybe try 1000t instead. And greater payload.For a dedicated tanker which does not share the outer mould line of the BFS then 100 tonnes is too high (I reckon something like 70 tonnes, but SpaceX might be able to reduce that even more). However, the first tanker flights may not be a dedicated design, instead just standard BFS with extra fuel in their tanks and perhaps long duration equipment removed. Then 100 tonnes dry is probably a bit too little.
Although a dedicated tanker design with all the excess mass removed would be cheaper in the long run, during BFR/BFS development it would be another craft competing for development funds and effort.
This other transaction agreement requires shared cost investment with SpaceX for the development of a prototype of the Raptor engine for the upper stage of the Falcon 9 and Falcon Heavy launch vehicles.
...Thanks. Terminal velocity should be in the range of 500 to 1,250 m/s at Mars, depending on atmospheric density, drag coefficient, etc. 2 km/s for EDL is very conservative.
"LEO esc 3.2Km/s + Fast transit ~1.7 Km/s + aerocapture + Mars landing 2Km/s ~ 7 Km/sec Delta V for 2nd stage MCT" Since I was sloppy and forgot to include the URL on my spreadsheet I'm not sure of the source. I was trying for the 7Km/sec to Mars and propellant tank volume for 8 something Km/sec for return from Mars with reduced payload, 25t, not 100t.
...
Or, this stage is a SpaceX long duration stage for bidding on the same contracts as Vulcan ACES.Right, but that could be done with their current kerolox stage as well, you just need more batteries.
2km/s is ridiculously high for EDL unless you're doing a large braking burn.
Sandbagging stuff is not helpful and will mislead you.
7km/s is reasonable for like 90-100 day transits and /especially/ bad transit opportunities. But I think 6km/s is enough for nominal mission.2km/s is ridiculously high for EDL unless you're doing a large braking burn.
Sandbagging stuff is not helpful and will mislead you.
...and reasonable Km/sec budget ranges from LEO to Mars surface are...?
7km/s is reasonable for like 90-100 day transits and /especially/ bad transit opportunities. But I think 6km/s is enough for nominal mission.If you budget 6.3 km/s (5.0 km/s for TMI and 1.3 km/s for EDL), you will typically be able to make a ~100-day transit with plenty of margin. Fast TMI is typically 4.5 to 4.9 km/s:
...a big question is Mars to Earth.
Right.7km/s is reasonable for like 90-100 day transits and /especially/ bad transit opportunities. But I think 6km/s is enough for nominal mission.If you budget 6.3 km/s (5.0 km/s for TMI and 1.3 km/s for EDL), you will typically be able to make a ~100-day transit with plenty of margin. Fast TMI is typically 4.5 to 4.9 km/s:
...a big question is Mars to Earth.
Edit: fuel loads and payloads can always be tweaked for faster or slower transits, so tank size isn't really a limiting design factor. Design for a standard case, then optimize for other factors.
7km/s is reasonable for like 90-100 day transits and /especially/ bad transit opportunities. But I think 6km/s is enough for nominal mission.2km/s is ridiculously high for EDL unless you're doing a large braking burn.
Sandbagging stuff is not helpful and will mislead you.
...and reasonable Km/sec budget ranges from LEO to Mars surface are...?
...a big question is Mars to Earth.
Upto now SpaceX modus operandi has been to under deliver initially and over deliver after a few iterations.
Dragon is not a scaled down MCT. So they will probably start with something small to validate it and learn and only later do the huge stuff when everything else closes.
But there's tons of room for mass optimization. How lightweight can things really get? Incredibly lightweight. Think backpacking, but with materials and manufacturing capabilities available in the 2030s.
2km/s is ridiculously high for EDL unless you're doing a large braking burn.
Sandbagging stuff is not helpful and will mislead you.
...and reasonable Km/sec budget ranges from LEO to Mars surface are...?
I think he's talking about starting at EML1, dropping to ~LEO and doing an Oberth burnIndeed. And if you had a full propellant load at EML1/2, you could do even faster transits, though you'd have to reserve some of the propellant for slowing down. You're hitting diminishing returns if you have to use a lot of propellant to slow down, but it's still interesting to consider.
Given Musk's suggestions of 2-3MN thrust to weight optimum for SC engines, I think that the design starts to look quite constrained. with perhaps the big choice being 2 rings (7-9 engines, 2000-3000 tonne GTOW) or 3 rings, 21-25 engines and 6-9000 tonne GTOW, and given development costs, launch site costs, and market it will almost certainly be the former.
MCT booster design height:
An optimally over-expanded rocket (eg RD193) produces 56 tonnes thrust per square meter of nozzle area at sea level, won't be much difference between various hydrocarbon rockets as all have similar pressures and velocities at nozzle exit.
Packing of multiple nozzles will only cover about 60% of the rocket base area at best, probably more like 50% given allowances for gimballing.
Need about 1.3g acceleration at liftoff (or more for higher ISP)
So 56*.5/1.3= ~20-25 tonnes of mass per square meter of base area that can be supported by a rocket. For LOX/CH4 bulk density of ~800kg/m that means that rockets can only support a column of fuel about 25-30m high in a prismatic form factor.
There are of course advantages to being thin for reduced aero losses - so maximum fuel column height within this constraint is likely to be the design chosen.
There will of course be substantial height taken by domed tank ends, 1st and 2nd stage engines, and finally cargo bay, but assuming a common bulkhead between LOX and CH4 and a single stick design I think could safely assume that will only be 30m+~1.4 stage diameter (4x domed ends for 2 stacked stages) + engine lengths (perhaps 4m per stage).
So overall I think 1st+2nd stage very unlikely to be taller than 50-60m. Though payload might add 30m in some cases. Ie rocket portion is pretty damned similar heights to Falcon9. It will just be substantially bigger diameter.
100M+ tall rockets only make sense for LH2.
Given Musk's suggestions of 2-3MN thrust to weight optimum for SC engines, I think that the design starts to look quite constrained. with perhaps the big choice being 2 rings (7-9 engines, 2000-3000 tonne GTOW) or 3 rings, 21-25 engines and 6-9000 tonne GTOW, and given development costs, launch site costs, and market it will almost certainly be the former.
Assuming 9x2.5MN engines each 4m long , that is about 9m diameter and total 50m tall (without cargo)
Assuming 25x2.5MN engines each 4m long, that is about 14m diameter and 58m tall (without cargo)
How about the reusable RaptorVac S2 first, likely 5+ meters, with a later Falcon-Raptor high performance S1 to replace the F9 Merlin cores? That eliminates duplicious ground systems and possibly having 2 launchers.
Unless they've already started building it.How about the reusable RaptorVac S2 first, likely 5+ meters, with a later Falcon-Raptor high performance S1 to replace the F9 Merlin cores? That eliminates duplicious ground systems and possibly having 2 launchers.
Musk is saying MCT will launch for Mars in 2022 if everything works right. I don't see any time to develop a ~5 meter Raptor booster before that...
I think it's pretty clear that they have started building Raptor. They might have started building a ~5m upper stage for it since that could go on FH pretty easily and maybe even on F9.
But building a Falcon Heavy replacement booster, before FH even flies? I don't think so. MCT has a pretty clearly defined set of goals including super heavy lift, and building another heavy lift booster doesn't help achieve those in the very tight timeframe Musk is shooting for.
IAC 2016 PLENARIES AND HIGHLIGHT LECTURES (http://www.iafastro.org/events/iac/iac2016/plenary-programme/)
Tuesday September 27
13:30-14:30 CDT = 14:30-15:30 EDT
Colonizing Mars -- A deep technical presentation on the space transport architecture needed to colonize Mars (SpaceX late breaking)
Is this in addition to the previously known Friday 8:30-10:30 CDT = 9:30-11:30 EDT session?
I'm not so sure the engines can't be protruding through a heatshield (more or less like a Falcon 9 first stage), assuming they have sufficient propellent to run at least some of the engines with enough power to use the supersonic retropropulsion effect to push the shock wave far enough out to reduce aero heating. For Mars decent that is probably easier since the atmosphere is thinner in the first place (less need for a heat shield), for Earth this may be more difficult.
Even if you cant the engines around the heatshield (more like Crew Dragon than Falcon 9 first stage), if you drop off the cargo as one giant module, you still have a problem, since you just dropped your (Mars) heat shield you need a second, more massive (at least you could save mass on the Martian one) Earth heat shield, which is above your engine nacelles. Alternatively you might put your engines far up the side of the vehicle to be above the Earth return heat shield, but now you need to shield the sides of the cargo from the engine exhaust and so on.
I think it's pretty clear that they have started building Raptor. They might have started building a ~5m upper stage for it since that could go on FH pretty easily and maybe even on F9.
But building a Falcon Heavy replacement booster, before FH even flies? I don't think so. MCT has a pretty clearly defined set of goals including super heavy lift, and building another heavy lift booster doesn't help achieve those in the very tight timeframe Musk is shooting for.
It's quite common to be designing and test building your next version of <whatever> before the current version hits the market...Especially true of things that have a long development time.
This would be developing and testing TWO versions ahead, since they have clearly been working on BFR for a while also. I don't see any reason to build a booster that will at best marginally outperform Heavy and will be superseded by BFR within ~2 years unless they plan to fly it concurrently with BFR... which seems unlikely.
On MCT speculation, I'm sure I read quotes by Musk saying he advocates fast transit times 102 or 100 days or something. Anyone know a source of such statement(s) or do I have a false memory?
They may want to fly BFR only to equatorial destinations. Which is most but not all of Falcon payloads.
They will want to go fully reusable on all their flights.Which would be possible with a Raptor upper stage on F9 or FH, although Raptor might be easier to reuse than Merlin since methane burns cleaner. The only real case for a ~5m booster that I can see is if kerolox turns out to be a significant pain to reuse compared to methalox, or if its a USEFUL part of the MCT architecture.
They may want to switch to all methane and Raptor. That would need a new first stage replacing Falcon. Not a priority but in the pipeline, I suspect. Especially if they build a wider upper stage.Flying a mix of kerolox and methalox is entirely viable long term IMO. Almost every launch provider has used higher energy fuels for upper stages for years, so these aren't new problems. Handling RP-1 is probably easier than LCH4.
They may want to switch to all methane and Raptor. That would need a new first stage replacing Falcon. Not a priority but in the pipeline, I suspect. Especially if they build a wider upper stage.Flying a mix of kerolox and methalox is entirely viable long term IMO. Almost every launch provider has used higher energy fuels for upper stages for years, so these aren't new problems. Handling RP-1 is probably easier than LCH4.
A switch to only producing Raptor is possible if Falcon is switched to using 5 Raptor engines on first stage AND the Raptor has a very deep throttle capability allowing it to land on the center engine.But you really might as well use a wider tank and go for more engines, like 7 or 9, to get at least FH performance. Still substantially less work than BFR to turnaround, but also allows full RTLS (no droneship needed) for almost all commercial payloads.
But you really might as well use a wider tank and go for more engines, like 7 or 9, to get at least FH performance. Still substantially less work than BFR to turnaround, but also allows full RTLS (no droneship needed) for almost all commercial payloads.
Just wondering, how hard is methalox propellant transfer in microgravity compared to hypergolic or hydrolox refueling?
Just wondering, how hard is methalox propellant transfer in microgravity compared to hypergolic or hydrolox refueling?
Unknown as nobody has done any.
Hydrolox would be the toughest mainly because of hydrogen's tiny atoms finding leaks really well and its need for really low temperature.
Hypergolic might be the easiest...until you accidentally started it up.
ULA did a LOX transfer test on a Centaur several years ago. The result was that cryo transfer was simple. Just need a pressure differential between tanks and a constant ulage motor firing to keep the cryo settled. The rest is just isolation valves opening and closing and pressures of the tanks control.Just wondering, how hard is methalox propellant transfer in microgravity compared to hypergolic or hydrolox refueling?
Unknown as nobody has done any.
Hydrolox would be the toughest mainly because of hydrogen's tiny atoms finding leaks really well and its need for really low temperature.
Hypergolic might be the easiest...until you accidentally started it up.
ISS is refueled by Progress.
I'm saying that I do think it would be impossible for it all to be ready for launch in six years, with no one hearing a peep of any such activity already under way.
I would love to be proven wrong, though. :)
Actually, retractable nozzles usually have an expansion ratio of ~50:1, but an extendable nozzle with the retraction split pulled back to about 15% of the nozzle length gets the retracted expansion ratio down to ~20:1 so the SL ISP goes up to ~300s and the thrust to ~2000 kN if they run the chamber at 12MPa. That ISP increase is enough to reduce fuel requirements to ~650t so that a very lightweight (~42t) 4-engine vehicle could put itself in a low parking orbit for refueling.They may want an engine like that for the center landing engine in any case.
I don't think anyone has flown a engine like that, and it has it own set of engineering difficulties. But that certainly doesn't mean SpaceX wouldn't try.
Actually, retractable nozzles usually have an expansion ratio of ~50:1, but an extendable nozzle with the retraction split pulled back to about 15% of the nozzle length gets the retracted expansion ratio down to ~20:1 so the SL ISP goes up to ~300s and the thrust to ~2000 kN if they run the chamber at 12MPa. That ISP increase is enough to reduce fuel requirements to ~650t so that a very lightweight (~42t) 4-engine vehicle could put itself in a low parking orbit for refueling.They may want an engine like that for the center landing engine in any case.
I don't think anyone has flown a engine like that, and it has it own set of engineering difficulties. But that certainly doesn't mean SpaceX wouldn't try.
Pulling this over from the BFR launch pad thread, since it's really BFS/MCT speculation.
Some estimates at what a Raptor flying on Falcon Heavy would look like (in the FH thread) put a different light on BFS as a potential self-SSTO. The ISP numbers Tom Mueller quoted for the SL Raptor engine (321s SL, 363s vacuum) are only achievable at much higher chamber pressures than Merlin, around 20-25 MPa. At those chamber pressures a RL10-B style engine becomes considerably more feasible for SSTO operation, with a retracted average ISP of 330 to 340s. Whether rapid reusability and long life are possible at those chamber pressures is questionable, but SpaceX clearly intends to try.
Specifically at 25 MPa, 3.45 O/F, 55:1 expansion (retracted), it would get 220 tonnes thrust at a 316s SL and 359s vac ISP, increasing to 265t and 380s when the vac bell is extended. 4 of those could theoretically put a 51t ship in low parking orbit, with a liftoff mass of 700t.
The other interesting factor is that a 4 to 5 meter vac bell is unnecessary at these chamber pressures. A 25 Mpa engine can realistically get 265t of thrust at 380s ISP with only a 3.15m nozzle - only slightly larger than Mvac.
That's interesting. So the best way to increase the payload of the Falcon Heavy is a bigger upper stage. If such a need existed, I would guess space X would choose to stretch the second stage, and add two more vaccum Merlins to the stage. Easier than integating a second fuel type into the launch infrastructure to allow for the use of the Raptor engine.No need to add an engine, the Merlin has plenty of thrust for a larger upper stage, especially if that stage is only used with the FH so it will be further along than the current F9 upper stage is at ignition.
The reason to go Raptor is to get improved ISP and more balanced fuel/oxidizer temperatures (which may make long life easier without adding extra mass).
Plus the Merlin Vacuum engine is gigantic, there is no way to fit 3 of them on any remotely Falcon-sized stage. You'd need a 8m diameter interstage at minimum.
To re-iterate that point.
(http://i.imgur.com/FRszulq.png)
MVac pushes 95 tonnes. By the time the fuel runs out, the engine has to throttle below 40% so as to not kill the payload with high g-forces.
...or have an extendible nozzle like the part that you trimmed from my post.The feasible way is to have a wider upper stage and interstage.This probably is what they will do if they go wider. But I think it will take some careful thought, since the widest part of the nozzle is at the bottom. And the bottom is where the narrowest part of the interstage would be, since that is where it starts flaring out... so there may need to be a gap of some size to get it to work.
The last bit of the conversation is great.
Extendible nozzle helps with interstage length, doesn't nothing for width.
A 4.8 meter nozzle is 4.8 meters whether in 1, 2 or more pieces. As we've seen from on board footage the separation events are not that smooth.
A rocket with a 500,000 lbf upper stage is going to need good clearances and some creative problem solving.
Of course maybe there is a smaller Raptor in the works and this discussion changes.
Raptor with a 4 m nozzle loses about 1% of ISP compared to a 4.8 m nozzle: 376 s vs 380 s.How much ISP would Raptor lose when it has a nozzle diameter of only 4m?Does it have to make the same thrust as the 4.8m nozzle? Because ISP isn't solely dictated by nozzle size - for the same fuels it is a function of pressure ratios... and there are other ways to change pressure ratios (e.g. change mass flow and/or throat diameter) at the expense of thrust.
Obviously I don't even know enough to ask the right question. For arguments sake assume the same engine, just a smaller nozzle.
Though what we know does not rule out that a dedicated smaller version of Raptor might be built.
This is based on sim in RPA lite using: Methane/LOX at:
9.7 MPa chamber pressure (same as Merlin)
2.8 O/F ratio (optimum for methalox at 9.7 MPa)
165 expansion ratio for the 4.8 m nozzle (same as Merlin Vac)
115 expansion ratio for the 4 m nozzle (assuming same throat diameter as the 4.8 m nozzle)
O/F ratio was supposed to be 3.8. Chamber pressure seems like the big unknown - if the M1D gas generator produces 9.7MPa, what should we expect in a FFSC Raptor?
Musk: "The critical elements of the solution are rocket reusability and low cost propellant (CH4 and O2 at an O/F ratio of ~ 3.8 ). And, of course, making the return propellant on Mars, which has a handy CO2 atmosphere and lots of H2O frozen in the soil." - http://waitbutwhy.com/2015/08/how-and-why-spacex-will-colonize-mars.html/4
Neither of those particularly affect the relative differences between the 4 meter and 4.8 meter nozzle; the benefit of having extra expansion is about the same for a higher pressure engine.
I just grabbed that 2.8:1 O/F number off a chart, and according to RPA it is actually a bit low for a 165:1 expansion vacuum engine. At 10 MPa the O/F for optimal ISP is 3.4:1, and at 15 MPa it optimizes at 3.45:1. However, 3.8:1 is rather high and actually loses a second or two of ISP (although you get a bit better thrust due to higher massflow).How are real numbers likely to come in, in relation to the figures provided by RPA Lite? Would the RPA numbers represent a theoretical maximum or perhaps a plausible best-guess at the exact values?
RPA calculates both the theoretical best possible performance, and a best guess at the actual performance after accounting for combustion and nozzle inefficiencies. I have no idea how accurate it is in reality. And we don't have a ton of known parameters to put in.From the O/F and the two isp points I had estimated 20.5MPa as Pc. Can't recall the expansion.The four datapoints we have are 321s (sealevel small-nozzle?) , 363s (vacuum small-nozzle?), and 380s (vacuum big-nozzle?), with an O:F ratio of 3.8. I had 14.5MPa working for theoretical numbers ("Express thermodynamic analysis") at expansion ratios of 32.5 and 75.
Changing it to "Extended analysis" and examining chamber performance given the estimated reaction and nozzle efficiencies, I end up losing 14s off of each number.I used actual vs theoretical. And modified a bit the freeze chemistry since the ch4/lox combustion, while appears simple, has a lot of steps.
Punch it up to the RD-180 chamber pressure - 26.7MPa
Using the listed estimates for 3.8:1 inefficiencies rather than theoretical maxima in RPA Lite:
With an expansion ratio of 59, you get 321s at sealevel and 363s in vacuum.
With an expansion ratio of 164, you get 380s in vacuum.
Sidenote: Do we have a higher chamber pressure liquid rocket engine in existence than the RD-180?
That's very interesting. Do you really think they will try to run at 26.7 MPa? For comparison, the RD-191 and RD-180 run at 26.7 MPa, and the SSME ran up to 21 MPa. I think that's the highest pressure ever flown in a reusable engine. A 100% length bell nozzle gets the same performance at 25 MPa and 55:1 expansion (155:1 for the vac nozzle).
The nice thing about higher chamber pressures is the engine gets much smaller (but not lighter) for the same thrust. At 25 MPa the SL Raptor only needs a 1.88m diameter nozzle to get Musk's estimated 500klbf of thrust, and the vac version gets 591klbf with only a 3.16m dia nozzle.
To bring this entire conversation back to Falcon Heavy (we were wandering OT for a bit there), that would mean it's possible a 25 MPa Raptor Vac could fly in the current Falcon interstage. Even with the nozzle trimmed back to 70% (which would put it very nearly in the Mvac envelope), it could realistically get 265 tonnes of thrust at 377s ISP. That's a healthy upgrade over Mvac's 95 tonnes thrust and 348 ISP.
I have no idea. I was about to ask you, and everybody else.
What chamber pressures should we expect out of Raptor? Do we have any evidence or reasoning one way or the other? (other than the previously mentioned 321s, 363s, 380s, and 3.8:1, and "FFSC should permit higher chamber pressures" notion)
Sidenote: Do we have a higher chamber pressure liquid rocket engine in existence than the RD-180?RD-191, 262.6kg/cm² vs 261.7kg/cm². RD-270 was 266.1. I know of nothing bigger than this. But some military RCS engines (like the one in MIRVs) probably have higher Pc.
I had always though the estimates of nozzle diameter for the Raptor were WAY too high, if were looking at vac nozzles only around 3 m across then a hexagonal 6 engine arrangement can easily fit in an upper-stage of around 12.5 m in diameter.
Your payload is too small at 100 mt, I'm thinking payload to LEO of around 200 mt even after engine loss, so that would require 4 functional engines and 6 total.
How much dV are you estimating for return, and how fast a transit? A typical 6-month return is only about 6.5 to 7 km/s from Mars surface to Earth surface (5.25 for launch to escape, 1 to 1.5 for transfer injection, 0.35 for EDL).6.5-7km/s sounds about right. I expect return trips to be much longer than 100 days.
Increasing return DV reduces the total mission duration by allowing a earlier launch, but I don't see any solutions in the trajectory browser for reducing the actual return transit time below about 5 months. If you have computed such solutions I'd love to see them.
How much dV are you estimating for return, and how fast a transit? A typical 6-month return is only about 6.5 to 7 km/s from Mars surface to Earth surface (5.25 for launch to escape, 1 to 1.5 for transfer injection, 0.35 for EDL).6.5-7km/s sounds about right. I expect return trips to be much longer than 100 days.
Increasing return DV reduces the total mission duration by allowing a earlier launch, but I don't see any solutions in the trajectory browser for reducing the actual return transit time below about 5 months. If you have computed such solutions I'd love to see them.
A typical 6-month return is only about 6.5 to 7 km/s from Mars surface to Earth surface (5.25 for launch to escape, 1 to 1.5 for transfer injection, 0.35 for EDL).We had already established that the trip would be 3 months?
Who said the return trip must be the same length of time as the trip to Mars?A typical 6-month return is only about 6.5 to 7 km/s from Mars surface to Earth surface (5.25 for launch to escape, 1 to 1.5 for transfer injection, 0.35 for EDL).We had already established that the trip would be 3 months?
The return flight will not likely be in the exact optimum window. Also, it'll likely be less than 6 months.
Remember, SpaceX wants the vehicle back each time.
Why rush faster than 6 months on return? Take 3-4 months Earth to Mars with humans. One month or less on Mars. 6 months return, some little cargo and a few possible humans. Plenty of time to get the BFS ready for the next synod which is the goal.
LMO refueling isn't required.The return flight will not likely be in the exact optimum window. Also, it'll likely be less than 6 months.
Remember, SpaceX wants the vehicle back each time.
I don't think it's feasible or necessary for early missions to return the vehicle for launch the next synod. IMO that won't happen until LMO refueling is possible and all the kinks/upgrades are worked out of the system. Probably 2030's. That requires a lot of pre-positioned assets to support fast ground unloading/refueling and LMO refueling.
...That changes if you're using massive amounts of ISRU propellants.
Orbital mechanics generally preclude having both short stays and fast transits.
...That changes if you're using massive amounts of ISRU propellants.
Orbital mechanics generally preclude having both short stays and fast transits.
Or you just develop a lightweight vehicle with lots of room for propellant....That changes if you're using massive amounts of ISRU propellants.
Orbital mechanics generally preclude having both short stays and fast transits.
You are right that in orbit refuelling is not necessary. It would make using massive amounts of propellants much easier though. Especially if there were in orbit ressources for fuel available.
...That changes if you're using massive amounts of ISRU propellants.
Orbital mechanics generally preclude having both short stays and fast transits.
You are right that in orbit refuelling is not necessary. It would make using massive amounts of propellants much easier though. Especially if there were in orbit ressources for fuel available.
Your payload is too small at 100 mt, I'm thinking payload to LEO of around 200 mt even after engine loss, so that would require 4 functional engines and 6 total.
But why would you want that? Staging to LEO and LEO to Mars surface have almost exactly the same DV requirements, so the vehicle should be optimized to deliver the same payload to both. A typical fast transit and EDL only requires ~6 kms. Launching 200t payload and 100t ship into LEO requires about 1400t of propellant in the US, but transit and EDL of 100t payload and 100t ship only require about 800t. You're shipping an extra 20% to 30% of useless dry mass (engines and tankage) to Mars and back.
For LEO tanker runs the 200t payload is useful, but that can be accomplished with larger tanks that the Mars transit ship doesn't need. And it still doesn't really need 6 engines for high-thrust engine out capability, because a tanker won't be carrying people and it can always use some of it's extra fuel load for margin. An off-nominal LEO launch has a lot of options for abort, particularly if a rapid launch cadence, on-orbit refueling, and LEO rendezvous are SOP.
... The ideal Earth launch vehicle is thus a simple 2nd stage with a payload fairing which can launch any cargo imaginable, and the ideal mars lander is a modest size capsule that can fit on top of it as a 3rd stage.
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Richard Heidmann (Snecma and Ariane) about the MCT (http://planete-mars.com/mars-colonization-transport-rumeurs-avant-revelation-du-projet/).
A winged belly-landing super-shuttle?
A winged belly-landing super-shuttle?
In the latest article the "super-shuttle" is not really a shuttle. It's more about maximizing the surface available for aerobraking. This is at least how I interpret the renders.
A winged belly-landing super-shuttle?
In the latest article the "super-shuttle" is not really a shuttle. It's more about maximizing the surface available for aerobraking. This is at least how I interpret the renders.
It's maximizing the surface available for aerobraking by entering with the ventral side forward, like the shuttle. That's great for increasing ballistic drag, but not great for optimizing structural mass. And the horizontal take-off/landing is interesting.
That image is from the older article. The article from 07/09/2016 shows a different BFS: just small winglets. IMHO this one is meant to actually land vertically.
EDIT: Of course it will enter "with the ventral side forward", but later on reorient for a vertical landing. I don't know whether this is actually possible.
Where does it show a vertical landing? He seems rather stuck on the horizontal landing, and there are no engines in the rear of that thing in the illustrations.
Which makes me wonder... how does it get to orbit? Horizontally?
The booster doesn't go to orbit, and there's no second stage so the shuttle must be it's own 2nd stage (confirmed by the trajectory graphic). At staging it's clearly pointed nose-downrange, but there are apparently no engines in the tail... so does it point it's dorsal side downrange and use belly engines? That seems questionable as there are still considerable aero loads at staging. Or do the engines pivot out and point back enough to continue vertically? There's no mechanism described for that.
The booster clearly lands vertically, but I don't see any indication that the shuttle does.
Vertical landing after a ventral entry certainly seems feasible, that's not what I was questioning. I don't see any indication that Heidmann considered a vertical landing at all for this spaceship: all his articles show horizontal landings.Do we need a new thread for Heidmann's design? Because this thread is supposed to be about MCT, and it may get confusing if we're talking about a /specific/ (and fairly fleshed-out) speculation of someone's own imagining in a thread that's supposed to be about SpaceX's actual MCT.
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Vertical landing after a ventral entry certainly seems feasible, that's not what I was questioning. I don't see any indication that Heidmann considered a vertical landing at all for this spaceship: all his articles show horizontal landings.
I was really wondering how it goes up, since all the visible engines are pointed to the side during launch.
Vertical landing after a ventral entry certainly seems feasible, that's not what I was questioning. I don't see any indication that Heidmann considered a vertical landing at all for this spaceship: all his articles show horizontal landings.
I was really wondering how it goes up, since all the visible engines are pointed to the side during launch.
This is something that I have wondered about. In most phases of the mission a horizontal orientation is better. EDL at both mars and earth, and surface operations. In transit does not care about the orientation. The only part that "wants" to be vertical is the few minutes it takes to leave earth's atmosphere. If you are going to have side mounted engines for super sonic retro propulsion you might as well use them for ascent.
The trade off is that you move your flip maneuver from just before landing to just after booster separation. That gives you one flip event rather than two. Though I am not sure it is inherently risky enough to matter. I know that as a passenger I would rather perform it once at altitude.
Who said you'd have side mounted engines for supersonic retropropulsion?Vertical landing after a ventral entry certainly seems feasible, that's not what I was questioning. I don't see any indication that Heidmann considered a vertical landing at all for this spaceship: all his articles show horizontal landings.
I was really wondering how it goes up, since all the visible engines are pointed to the side during launch.
This is something that I have wondered about. In most phases of the mission a horizontal orientation is better. EDL at both mars and earth, and surface operations. In transit does not care about the orientation. The only part that "wants" to be vertical is the few minutes it takes to leave earth's atmosphere. If you are going to have side mounted engines for super sonic retro propulsion you might as well use them for ascent.
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Who said you'd have side mounted engines for supersonic retropropulsion?Vertical landing after a ventral entry certainly seems feasible, that's not what I was questioning. I don't see any indication that Heidmann considered a vertical landing at all for this spaceship: all his articles show horizontal landings.
I was really wondering how it goes up, since all the visible engines are pointed to the side during launch.
This is something that I have wondered about. In most phases of the mission a horizontal orientation is better. EDL at both mars and earth, and surface operations. In transit does not care about the orientation. The only part that "wants" to be vertical is the few minutes it takes to leave earth's atmosphere. If you are going to have side mounted engines for super sonic retro propulsion you might as well use them for ascent.
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In that video from Max Fagin, he strongly advocated side-mounted engines to increase the drag. OFC, he was only an intern at SpX, so that doesn't mean that his opinion represents corporate plans.Who said you'd have side mounted engines for supersonic retropropulsion?Vertical landing after a ventral entry certainly seems feasible, that's not what I was questioning. I don't see any indication that Heidmann considered a vertical landing at all for this spaceship: all his articles show horizontal landings.
I was really wondering how it goes up, since all the visible engines are pointed to the side during launch.
This is something that I have wondered about. In most phases of the mission a horizontal orientation is better. EDL at both mars and earth, and surface operations. In transit does not care about the orientation. The only part that "wants" to be vertical is the few minutes it takes to leave earth's atmosphere. If you are going to have side mounted engines for super sonic retro propulsion you might as well use them for ascent.
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It seems very unlikely that SpaceX would launch horizontally from Mars. Nor does horizontal engines on an upper stage make sense.
Except there's still a lot of drag for a second stage, especially one like the BFS (which probably stages fairly early).It seems very unlikely that SpaceX would launch horizontally from Mars. Nor does horizontal engines on an upper stage make sense.
You will have to help me on the basis for those assertions. With a thin martian atmosphere there would not be much penalty. Throttling would certainly have to be more responsive than vertical launching but that is not a known show stopper.
You don't have horizontal engines on an upper stage. You have an upper stage that you are launching sideways. You just light the engines on one side of it slightly before the other and it rotates 90 degs. It makes as much sense as a swan dive maneuver on a landing craft.
In that video from Max Fagin, he strongly advocated side-mounted engines to increase the drag. OFC, he was only an intern at SpX, so that doesn't mean that his opinion represents corporate plans.Who said you'd have side mounted engines for supersonic retropropulsion?Vertical landing after a ventral entry certainly seems feasible, that's not what I was questioning. I don't see any indication that Heidmann considered a vertical landing at all for this spaceship: all his articles show horizontal landings.
I was really wondering how it goes up, since all the visible engines are pointed to the side during launch.
This is something that I have wondered about. In most phases of the mission a horizontal orientation is better. EDL at both mars and earth, and surface operations. In transit does not care about the orientation. The only part that "wants" to be vertical is the few minutes it takes to leave earth's atmosphere. If you are going to have side mounted engines for super sonic retro propulsion you might as well use them for ascent.
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Except there's still a lot of drag for a second stage, especially one like the BFS (which probably stages fairly early).It seems very unlikely that SpaceX would launch horizontally from Mars. Nor does horizontal engines on an upper stage make sense.
You will have to help me on the basis for those assertions. With a thin martian atmosphere there would not be much penalty. Throttling would certainly have to be more responsive than vertical launching but that is not a known show stopper.
You don't have horizontal engines on an upper stage. You have an upper stage that you are launching sideways. You just light the engines on one side of it slightly before the other and it rotates 90 degs. It makes as much sense as a swan dive maneuver on a landing craft.
Except there's still a lot of drag for a second stage, especially one like the BFS (which probably stages fairly early).It seems very unlikely that SpaceX would launch horizontally from Mars. Nor does horizontal engines on an upper stage make sense.
You will have to help me on the basis for those assertions. With a thin martian atmosphere there would not be much penalty. Throttling would certainly have to be more responsive than vertical launching but that is not a known show stopper.
You don't have horizontal engines on an upper stage. You have an upper stage that you are launching sideways. You just light the engines on one side of it slightly before the other and it rotates 90 degs. It makes as much sense as a swan dive maneuver on a landing craft.
Except there's still a lot of drag for a second stage, especially one like the BFS (which probably stages fairly early).It seems very unlikely that SpaceX would launch horizontally from Mars. Nor does horizontal engines on an upper stage make sense.
You will have to help me on the basis for those assertions. With a thin martian atmosphere there would not be much penalty. Throttling would certainly have to be more responsive than vertical launching but that is not a known show stopper.
You don't have horizontal engines on an upper stage. You have an upper stage that you are launching sideways. You just light the engines on one side of it slightly before the other and it rotates 90 degs. It makes as much sense as a swan dive maneuver on a landing craft.
The shuttle shape is going to be most aerodynamically stable in the reentry configuration, so launching from Mars (or after staging) horizontally would be like trying to fly a jet airplane backwards. Possible, perhaps, but definitely not ideal. Especially with those winglets.
Landing horizontally seems like a terrible idea, not at all SpaceX-y.
DC-X demonstrated swan dive maneuver already. We don't need this horizontal landing nonsense.
The booster doesn't go to orbit, and there's no second stage so the shuttle must be it's own 2nd stage (confirmed by the trajectory graphic). At staging it's clearly pointed nose-downrange, but there are apparently no engines in the tail... so does it point it's dorsal side downrange and use belly engines? That seems questionable as there are still considerable aero loads at staging. Or do the engines pivot out and point back enough to continue vertically? There's no mechanism described for that.
The booster clearly lands vertically, but I don't see any indication that the shuttle does.
I've looked up the original article of version II of Heidmann's design. The ship flies up after separation in the horizontal position using the belly landing nozzles. The back doors stay closed. There are no additional thrusters.The booster doesn't go to orbit, and there's no second stage so the shuttle must be it's own 2nd stage (confirmed by the trajectory graphic). At staging it's clearly pointed nose-downrange, but there are apparently no engines in the tail... so does it point it's dorsal side downrange and use belly engines? That seems questionable as there are still considerable aero loads at staging. Or do the engines pivot out and point back enough to continue vertically? There's no mechanism described for that.
The booster clearly lands vertically, but I don't see any indication that the shuttle does.
Yes, at booster staging, the ship would need to pitch 90 degrees, open it's belly engine doors (which would need to be closed during launch for aerodynamics) and then begin thrusting up to LEO. Would probably need to be high enough so that any aerodynamic issues would be minimal for that. At that point it can keep the doors open for TMI. It would need to close them for Mars atmospheric entry, and then open them up again at terminal velocity to propulsively brake and then land. . They could stay open on the surface until launch, and stay open for TEI. They'd need to close again for Earth atmospheric entry, and then open again at terminal velocity for propulsive landing and braking.
It also provides a pretty un-aerodynamic surface for Mars Launch like that. But maybe that's not too much of a problem in the thin Mars atmosphere? Not sure.
Vertical landing after a ventral entry certainly seems feasible, that's not what I was questioning. I don't see any indication that Heidmann considered a vertical landing at all for this spaceship: all his articles show horizontal landings.Do we need a new thread for Heidmann's design? Because this thread is supposed to be about MCT, and it may get confusing if we're talking about a /specific/ (and fairly fleshed-out) speculation of someone's own imagining in a thread that's supposed to be about SpaceX's actual MCT.
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Have you read the title of the thread, it is indeed a SPECULATION thread,
Have you read the title of the thread, it is indeed a SPECULATION thread,
It is supposed to be a MCT speculation thread. Not a generic speculation thread on just any Mars architecture. We have the Mars section for that. Things here should be based on what was said by SpaceX about MCT.
Vertical landing after a ventral entry certainly seems feasible, that's not what I was questioning. I don't see any indication that Heidmann considered a vertical landing at all for this spaceship: all his articles show horizontal landings.
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Do we need a new thread for Heidmann's design? Because this thread is supposed to be about MCT, and it may get confusing if we're talking about a /specific/ (and fairly fleshed-out) speculation of someone's own imagining in a thread that's supposed to be about SpaceX's actual MCT.
ISTM that accommodating two main load axis (for a 100 mT to Mars surface vehicle) that are perpendicular one to another is in severe conflict with SX known optimizations for PMF and cost. That alone seems to make the Heidmann (HCT) design off-topic for any MCT discussion, before even counting L2 info available (which further invalidates HCT as possible SX design). YMMV
I was simply suggesting that if we're starting to talk for several pages about one person's specific vision for what MCT is like, and if it's highly fleshed out, then it's probably best to have a dedicated thread. That doesn't mean it's off-topic here, but that it might be better in its own thread. Geez y'all.
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I don't see anyway that a super-dragon style capsule will get enough drag to slow down enough.
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I was simply suggesting that if we're starting to talk for several pages about one person's specific vision for what MCT is like, and if it's highly fleshed out, then it's probably best to have a dedicated thread. That doesn't mean it's off-topic here, but that it might be better in its own thread.
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I don't see anyway that a super-dragon style capsule will get enough drag to slow down enough.
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If Red Dragon works, a larger vehicle with the same entry velocity, lift/drag ratio, ballistic coefficient, propellant mass fraction, and engine efficiency should work just as well. Because of the cube-square law it will have to be less dense, but since it's largely made of nearly empty propellant tanks that shouldn't be a major issue.
Trading propellant mass for engine efficiency (Raptor 380s vs. SuperDraco 240s) also helps hit the required density reduction.
I was simply suggesting that if we're starting to talk for several pages about one person's specific vision for what MCT is like, and if it's highly fleshed out, then it's probably best to have a dedicated thread. That doesn't mean it's off-topic here, but that it might be better in its own thread. Geez y'all.Yes.
Yes, but -- how wide would such a scaled-up Dragon capsule be at the base, if it's got to be designed to carry at least 100 people, along with all the stuff (like food, water and life support) they will need for three to six months?
I think simply scaling up the dimensions of a Dragon to where it has the interior space available for the actual stated mission requirements would make the base diameter in the hundreds of meters. How ya gonna fit that on any booster, much less one with the most-commonly-speculated (see L2) width of the BFR?
And if you change the basic shape, you inevitably have to go through a lot of design work to achieve similar lift/drag characteristics, etc. In other words, BFS is gonna have to be designed from scratch, it won't take advantage of Dragon's shape or flying characteristics as a starting point.
Mars Colonisation Transportation: Project Revelation Before Rumours
But a noise is that SpaceX would work on a Raptor 700 Tf !
One interesting tidbit, though:QuoteBut a noise is that SpaceX would work on a Raptor 700 Tf !
Not sure of rumor source for an F-1 class engine, but certainly would simplify first stage design if nine 1.5-1.6Mlbf engines were used on booster.
One interesting tidbit, though:QuoteBut a noise is that SpaceX would work on a Raptor 700 Tf !
Not sure of rumor source for an F-1 class engine, but certainly would simplify first stage design if nine 1.5-1.6Mlbf engines were used on booster.
Not sure how this jives with the news that the AirForce is paying for a raptor based upper for F9 and FH. 2 engines or rumor wrong?
So, it wouldn't surprise me if September reveal shows more than one engine size. (Three is my guess.)
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Landing engines on Mars would require deep throttling of 550klbf engine.
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Yes, but -- how wide would such a scaled-up Dragon capsule be at the base, if it's got to be designed to carry at least 100 people, along with all the stuff (like food, water and life support) they will need for three to six months?
I think simply scaling up the dimensions of a Dragon to where it has the interior space available for the actual stated mission requirements would make the base diameter in the hundreds of meters. How ya gonna fit that on any booster, much less one with the most-commonly-speculated (see L2) width of the BFR?
And if you change the basic shape, you inevitably have to go through a lot of design work to achieve similar lift/drag characteristics, etc. In other words, BFS is gonna have to be designed from scratch, it won't take advantage of Dragon's shape or flying characteristics as a starting point.
Let's do at least a first order pass at actually calculating this. If BFR's diameter is 15m, BFS can easily be at least 15 to 16m. Dragon's OML including the nose but not the trunk encloses about 25 m^3 with a major radius of 1.85 meters. Volume increases with size as r^3, so a 15 meter Dragon would enclose about 1,875 m^3 and a 16m version about 2,050 m^2
Estimates of necessary volume per person and propellant mass vary, but usually range somewhere from 5 to 15 m^2 per person and 300-1200 tonnes of propellant. Averaging those gets 10 m^3 per person and 750t propellant, or 1000 m^2 habitable volume and 915 m^3 tank volume (100 people, 820 kg/m^3 methalox). That totals 1915 m^3, or pretty close to what a 15 to 16m Dragon would enclose.
Obviously, that's only enclosed volume and not all the space can be utilized efficiently, and engines etc. all take up volume. But volume shouldn't be a show-stopper considering that it's probably feasible to launch up to a 20m diameter (4,500 m^3 volume) object on top of BFR.
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Landing engines on Mars would require deep throttling of 550klbf engine.
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Just to address this point: Merlin throttles to 40%, and MVac to 30%. That's more than deep enough to land 100t payload on Mars even using two 550 klbf engines.
Even if spacecraft dry weight is same as payload, T/W is still greater than 1 with a single engine.
It is clear Spacex is mastering first stage usability, but they indicated they are not finding practical bringing the second stages down to earth. ULA on their part are toying with ACES, a second stage spending extended time in orbit, but they have not been explaining how they plan to bring supplies to orbit without creating a glut of second stages up there... If it is not practical to bring back to earth the second stage, the only remaining solution for a reusable second stage is to park it in orbit and perform a suborbital load transfer from the first stage carrying the load and a second stage descending from orbit to pick up the load. Unquestionable the maneuver is tricky and must be performed quickly... Have you seen discussions on this topic somewhere?SpaceX didn't think reuse of the second stage was practical when both these conditions are true:
... If it is not practical to bring back to earth the second stage, the only remaining solution for a reusable second stage is to park it in orbit and perform a suborbital load transfer from the first stage carrying the load and a second stage descending from orbit to pick up the load. Unquestionable the maneuver is tricky and must be performed quickly... Have you seen discussions on this topic somewhere?
Even if spacecraft dry weight is same as payload, T/W is still greater than 1 with a single engine.
There is the all important difference between weight and mass. It is the weight that would need to be countered for a hover. But they don't want to hover. It is the mass that needs to be decelerated. That's mostly the same effort as on earth for landing, just with less gravity losses.
My expectation is that the main propulsion will 'hover slam' to a full stop at a height of a few dozen meters above the ground and then cut out. Then the vehicle will touch down on orbital maneuvering engines located higher on the vehicle and canted outward, where they will not impinge on the ground, this will only require the equivalent of a dozen Super Draco engines.
My expectation is that the main propulsion will 'hover slam' to a full stop at a height of a few dozen meters above the ground and then cut out. Then the vehicle will touch down on orbital maneuvering engines located higher on the vehicle and canted outward, where they will not impinge on the ground, this will only require the equivalent of a dozen Super Draco engines.
Some form of orbital maneuvering engine is a must, you can't use the main propulsion system for thouse kinds of maneuvers. Draco (not super) might do the very fine maneuvers for docking and such but you will still want something stronger for a de-orbit burn.
Lets compare to shuttle, it massed around 75 mt on orbit and had OMS with 53 kN thrust. Scale up to a likely 200+ mt mass in orbit and your looking at the thrust of 2 Super Draco engines for a de-orbit burn. A Raptor engine would need to throttle down to 6% to give that thrust.
Still we would be looking at about a 6 fold increase over a reasonable orbital maneuvering system so it's not free. Their may be additional benefits to these engines, their position high on the vehicle gives them huge leverage on the vehicle useful in EDL if a flip over is needed.
Some form of orbital maneuvering engine is a must, you can't use the main propulsion system for thouse kinds of maneuvers. Draco (not super) might do the very fine maneuvers for docking and such but you will still want something stronger for a de-orbit burn.
Lets compare to shuttle, it massed around 75 mt on orbit and had OMS with 53 kN thrust. Scale up to a likely 200+ mt mass in orbit and your looking at the thrust of 2 Super Draco engines for a de-orbit burn. A Raptor engine would need to throttle down to 6% to give that thrust.
Still we would be looking at about a 6 fold increase over a reasonable orbital maneuvering system so it's not free. Their may be additional benefits to these engines, their position high on the vehicle gives them huge leverage on the vehicle useful in EDL if a flip over is needed.
Couldn't you just use a single Raptor and fire it for a shorter time period? What type of delta V is needed for a de-orbit burn typically?
Some form of orbital maneuvering engine is a must, you can't use the main propulsion system for thouse kinds of maneuvers. Draco (not super) might do the very fine maneuvers for docking and such but you will still want something stronger for a de-orbit burn.
Lets compare to shuttle, it massed around 75 mt on orbit and had OMS with 53 kN thrust. Scale up to a likely 200+ mt mass in orbit and your looking at the thrust of 2 Super Draco engines for a de-orbit burn. A Raptor engine would need to throttle down to 6% to give that thrust.
Still we would be looking at about a 6 fold increase over a reasonable orbital maneuvering system so it's not free. Their may be additional benefits to these engines, their position high on the vehicle gives them huge leverage on the vehicle useful in EDL if a flip over is needed.
Couldn't you just use a single Raptor and fire it for a shorter time period? What type of delta V is needed for a de-orbit burn typically?
Another concept I just considered, with a small BFS and a larger second stage without using SEP. Launch with just cargo in the BFS and all propellant in the 2nd stage, the two stay together and reach orbit much like a Dragon capsule and it's trunk.
Then both are refueled to full, TMI is conducted by firing the 2nd stage for ~2 km/s of acceleration then the BFS separates and performs the remaining boost. This leaves the 2nd stage far short of Earth escape and it will be in an elliptical orbit which can easily allow it to land again.
Another 2nd stage without a BFS attached is also placed in orbit and refueled to make a TMI on it's own with a propulsive insertion at Mars, by utilizing a slow hohoman transfer the propellant delivery is much more efficient. BFS then rendezvouses with it in orbit and takes on the necessary propellant for TEI. The 2nd stage now very light now returns to Earth via a slow transfer and aero-captures at Earth.
The 2nd stage would be capable of this kind of total DeltaV because it is almost nothing but tank so a 4-5% dry mass fraction is reasonable for it (far lower then what the BFS could achieve). The main challenge is endurance and maintaining propellants against boil-off for that length of time, but any vehicle that waits in LEO while it is being filled up will need considerable insulation so this 2nd stage can perform the job of tanker to LEO, depot in LEO, tanker to Mars and heavy lifter to LEO for the BFS and other payloads.
I see this stage being around 60 mt in dry mass with 1200 mt of propellant capacity and equipped with solar and radiators on the surface as on the Dragon 2 capsule.
It is not "considered dangerous", it elevates the risks for mission success. It's a statistics and systems analysis thing and prior performance of the general principle of docking is probably only marginally relevant to a specific system, in this context. Three docking events present more risk than two identical docking events, even if you eventually manage to pull both flawlessly, 100 times each.
mfck you had it right, I am saying their is an irrational fear of rendezvouses in mission planning.
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NASA assesses the chance of a crewed Orion capsule with a contingency EVA option failing to dock with a lunar assent vehicle at just 1 in 546. https://www.nasa.gov/pdf/140639main_ESAS_08.pdf (Page 13)
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So I would estimate a failure chance of the BFS to achieve refueling in Mars orbit at around 1:1000 and not at all something worth of concern.
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mfck you had it right, I am saying their is an irrational fear of rendezvouses in mission planning.
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NASA assesses the chance of a crewed Orion capsule with a contingency EVA option failing to dock with a lunar assent vehicle at just 1 in 546. https://www.nasa.gov/pdf/140639main_ESAS_08.pdf (Page 13)
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So I would estimate a failure chance of the BFS to achieve refueling in Mars orbit at around 1:1000 and not at all something worth of concern.
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I am not an engineer, but those risk figures seem odd. 1:546 might be considered OK for Orion, a craft that will hardly (imo) fly 100 missions combined, but a 1:1000, for a mass transportation architecture that MCT is, seems unacceptable. What am I missing? What are the comparable risks in, say, aviation?
We don't know isn't a valid reason. Same line of reasoning would give us Apollo-like hypergols.
SpaceX is going to find out. So make a first-try estimate. You'll see that a ton of water per day isn't unreasonable to harvest since you need to harvest at least hundreds of kilograms per day anyway. Better than tripling the required number of launches and the complexity of the architecture.
While 936 might seem prohibitive, it is substituting for 600 mt of propellant MADE on the surface of Mars, so the ratio of substitution is ~3:2 when propellant is delivered to a high Mars orbit. This is quite different then the traditional view that Mars ISPP will have a IMLEO reduction factor of 10x - 20x but thouse estimates are based on taking all propellant down to the Martian surface which is an additional 10 km/s DeltaV, this scenario puts the refueling at the optimum point for efficient delivery.
A stage with all of the capabilities I've described for the 2nd stage is a necessity under any architecture, others have simply been calling it a 'BFS-Tanker' as it is clear that the cargo carrying BFS vehicle would not be an efficient tanker to LEO and something specialized for that purpose is necessary.
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A stage with all of the capabilities I've described for the 2nd stage is a necessity under any architecture, others have simply been calling it a 'BFS-Tanker' as it is clear that the cargo carrying BFS vehicle would not be an efficient tanker to LEO and something specialized for that purpose is necessary.
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Musk said that (paraphrasing) "he's tempted to pursue stage 2 reuse on FH, but it's probably better to focus on Mars architecture".
Since a reusable upper stage on FH is a natural iterative development towards any potential reusable S2 on BFR, I think it's highly probable based on this statement that the Mars architecture won't include a S2 at all. If BFR did have a S2 on it, any focus on developing a smaller version would help (not detract from) focus on the larger one.
So I think the Falcon upper stage is an architectural dead end, and BFR/BFS will derive from the vehicles SpaceX is currently reusing: Falcon Stage 1, and Dragon. That wouldn't mean that specialized (e.g. Tanker) version wouldn't exist, just that they would share heritage with Dragon and not with the F9 S2.
Most of the specializations (lightweight legs, insulation, cryocoolers, extra PV arrays, bigger tanks) will fit just as well in a similar outer mold line to BFS, even though internal and structural components would differ. It would be more like the relation between FH center and side boosters than between F9 S2 and Dragon: looks very much the same, but a bit different inside.
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But we now know that a Raptor upper stage is being made, which IS exactly what would make sense as a for runner to a reusable BFS upper stage, the new F9 upper stage can do controlled re-entry burn tests while still being disposed of, just as the F9 first stage was splashed into the ocean several times before even trying to put legs on it. The new upper stage lets SpaceX do it's tests at customer expense, which is how they like to do all their testing.
Err, how exactly do we now know that a Raptor upper stage is being made? I thought the only things we know are (1) The Air Force has thrown some money SpaceX's way to develop a Raptor vac engine, and (2) Musk has recently said that doing a new, reusable upper stage for Falcon is tempting, but SpaceX will be concentrating on developing the 'Mars rocket' instead. This would imply to me that (sadly) there is no effort underway to do a new upper stage for Falcon, and likely won't be in the future.
This other transaction agreement requires shared cost investment with SpaceX for the development of a prototype of the Raptor engine for the upper stage of the Falcon 9 and Falcon Heavy launch vehicles.
Outer mold line similarities are n8ot en8ough t8o c8onsider 8one vehic2le derivedI listed several other similarities; whether "derived" is a good descriptor of their relation is semantic.
...the new F9 upper stage can do controlled re-entry burn tests while still being disposed of...I don't think it can at all. It will break up from heat load on reentry unless heavily shielded - at which point it's basically a reusable upper stage, which Musk specifically said they aren't working on. The velocity change for a orbital reentry is about 4 to 5x greater, and the max heating rate 20 to 25x greater than what the F9 S1 sees on entry. Even if it's possible to orbit enough fuel retroburn through the peak heating phase (and I STRONGLY doubt that it is), why would they? The stage has a low ballistic coefficient, lots of area to dissipate heat, and slow terminal velocity: shield the front and part of one side, enter nose-first, and keep all the sensitive, expensive parts out of the hypersonic flow and well away from the bow shock. This is exactly how SpaceX envisioned S2 reuse... and they aren't pursuing it.
Err, how exactly do we now know that a Raptor upper stage is being made? I thought the only things we know are (1) The Air Force has thrown some money SpaceX's way to develop a Raptor vac engine, and (2) Musk has recently said that doing a new, reusable upper stage for Falcon is tempting, but SpaceX will be concentrating on developing the 'Mars rocket' instead. This would imply to me that (sadly) there is no effort underway to do a new upper stage for Falcon, and likely won't be in the future.
The Air Force said specifically that the engine is for F9.QuoteThis other transaction agreement requires shared cost investment with SpaceX for the development of a prototype of the Raptor engine for the upper stage of the Falcon 9 and Falcon Heavy launch vehicles.
http://www.defense.gov/News/Contracts/Contract-View/Article/642983
... Find out whether a Rapter VAC could be use on an F9 U/S
... Find out whether a Rapter VAC could be use on an F9 U/S
Almost certainly yes. Read the last few pages of the Raptor on F9/FH discussion: http://forum.nasaspaceflight.com/index.php?topic=39314
"Almost certainly" is "almost certainly" not good enough for the US DoD!
It could be, but it's not going to. At least not before MCT flies.
I started at least one thread speculating about a Raptor-based reusable upper stage for Falcon 9/FH, given that hint from the Air Force (which is not new, we've known about that for a quite long while). But now we know from Musk that they're not going to pursue that right now.
Sure, the future is wide open. But a post-MCT world is going to be weird and difficult to analyze before we see MCT and how well it flies.It could be, but it's not going to. At least not before MCT flies.
I started at least one thread speculating about a Raptor-based reusable upper stage for Falcon 9/FH, given that hint from the Air Force (which is not new, we've known about that for a quite long while). But now we know from Musk that they're not going to pursue that right now.
That leaves the questions of if and when wide open...
I don't think it can at all. It will break up from heat load on reentry unless heavily shielded - at which point it's basically a reusable upper stage, which Musk specifically said they aren't working on. The velocity change for a orbital reentry is about 4 to 5x greater, and the max heating rate 20 to 25x greater than what the F9 S1 sees on entry. Even if it's possible to orbit enough fuel retroburn through the peak heating phase (and I STRONGLY doubt that it is), why would they? The stage has a low ballistic coefficient, lots of area to dissipate heat, and slow terminal velocity: shield the front and part of one side, enter nose-first, and keep all the sensitive, expensive parts out of the hypersonic flow and well away from the bow shock. This is exactly how SpaceX envisioned S2 reuse... and they aren't pursuing it.
The same goes for BFR. I'm extremely doubtful that orbital reentry using primarily retropropulsion for shock standoff and cooling is possible and more optimal than nose-first entry for a S2 type vehicle. If there's any evidence to the contrary please point it out.
Isn't the mVac nozzle much too flimsy to withstand an engine first re-rentry? Or even a relatively low altitude/low speed (100's kmh) environment?Yes, it would be torn to shreds. You'd either have to stow it, or have most of it be disposable.
Isn't the mVac nozzle much too flimsy to withstand an engine first re-rentry? Or even a relatively low altitude/low speed (100's kmh) environment?Yes, it would be torn to shreds. You'd either have to stow it, or have most of it be disposable.
There are two feasible reentry positions, not just one. Side and engines-first. Both have been tested to varying degrees by SpaceX, Shuttle, DC-X, Blue Origin, as well as simulated.
There are two feasible reentry positions, not just one. Side and engines-first. Both have been tested to varying degrees by SpaceX, Shuttle, DC-X, Blue Origin, as well as simulated.
If second stages are going to be reused and deliver worthwhile payloads, I think most of the braking will need to be done with friction and not propulsion.
Matthew
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Second even at the time of the release of that infamous video it was obvious that the 2nd stage recovery shown was Holly Wood nonsense and nothing more then a placeholder, nose first re-entry is impossible because it is totally unstable, the engine is most massive part of a 2nd stage and this will dictate an engine first entry. Second it was clear that the nose of a second stage can't do the basic job of attaching a payload if it is a smooth heat shield.
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it is totally unstable, the engine is most massive part of a 2nd stage and this will dictate an engine first entry.
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...it was clear that the nose of a second stage can't do the basic job of attaching a payload if it is a smooth heat shield.
The only feasible entry is engine first, which means either a shield that moves to cover the engine...
Seems like a reusable Raptor 2nd stage could be done if it entered and landed nose first. Add upside down landing legs near the top of the stage, TPS around the 'nose' (perhaps even a bit oversized diameter, to protect the body of the stage), some Super Dracos for landing assist, and grid fins near the engine.
A reusable second stage would be actively guided it its landing zone, just like the first stage. No need to drop it in the ocean. It could land back at the launch site or on a barge.
BTW, this is the MCT speculation thread, so why are we having what seems to be a F9/FH discussion?
A reusable second stage would be actively guided it its landing zone, just like the first stage. No need to drop it in the ocean. It could land back at the launch site or on a barge.
BTW, this is the MCT speculation thread, so why are we having what seems to be a F9/FH discussion?
Agreed, it's in the wrong thread, I suspect it spun off the Raptor powered upper stage discussion. It should probably be moved.
Whilst the stage could be actively guided in the landing zone, for a GEO launch, the zone itself would be determined by the highly elliptical orbit after launching the satellite, and how quickly that orbit decayed. It can take years for that to happen, so there probably wouldn't be any fuel left for landing.
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it is totally unstable, the engine is most massive part of a 2nd stage and this will dictate an engine first entry.
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False. On a Dragon to ISS mission the residual propellent outweighs the entire second stage, and it only takes 1% of the initial 100 tonne propellent load to outweigh a Merlin.
At atmospheric drag will settle the remaining propellent in the nose long before it's significant enough to overpower the RCA thrusters, resulting in a quasi-stable configuration that cold be controlled by either RCS or small active aerodynamic surfaces during re-entry.
Does the mass of the heat shield exceed the mass of the fuel required to create a sufficiently large "bubble" ?
Please show your working :D
Does the mass of the heat shield exceed the mass of the fuel required to create a sufficiently large "bubble" ?Not but it will technologically and operation less complicated.
Please show your working :D
Is this a SSTO? How do you get enough Km/sec to reach LEO, etc?The MCT is currently assumed to be the second stage of a two stage LV. It will (probably) launch from mars as an SSTO, though.
How do humans get from the living area to/from escape pod with heat shield in between?
If second stages are going to be reused and deliver worthwhile payloads, I think most of the braking will need to be done with friction and not propulsion.
Matthew
How about like this?
Is this a SSTO? How do you get enough Km/sec to reach LEO, etc?Just showing MCT. BFR is not part of this picture, but BFR still need it. By the way BFR will be use together with MCT just 0.0000019% MCT Mars flight.
How do humans get from the living area to/from escape pod with heat shield in between?
If second stages are going to be reused and deliver worthwhile payloads, I think most of the braking will need to be done with friction and not propulsion.
Matthew
How about like this?
It has to turn around for landing, I think retro propulsive braking test give them numbers that propulsion is good enough to build shield.
Don't assume the EDL takes as long as a manned capsules entry, deceleration can be VERY fast if your coming in steep it just makes for very high g-forces and a tank can tolerate that.
Don't assume the EDL takes as long as a manned capsules entry, deceleration can be VERY fast if your coming in steep it just makes for very high g-forces and a tank can tolerate that.
Don't assume the EDL takes as long as a manned capsules entry, deceleration can be VERY fast if your coming in steep it just makes for very high g-forces and a tank can tolerate that.
The crew can tolerate a 20G entry.
Don't assume the EDL takes as long as a manned capsules entry, deceleration can be VERY fast if your coming in steep it just makes for very high g-forces and a tank can tolerate that.
No, frankly. They can't. Check your sources. We're talking about sustained G-loading that lasts tens of seconds at near-peak intensity, presumably with considerable additional vibration from the turbulence. Per Wikipedia, "Only the most motivated volunteers were capable" of reaching this 20G level in testing for 10s duration with a steady centrifuge, and more than that is suggested to result in injury and/or LOC.Don't assume the EDL takes as long as a manned capsules entry, deceleration can be VERY fast if your coming in steep it just makes for very high g-forces and a tank can tolerate that.
The crew can tolerate a 20G entry.
No, frankly. They can't. Check your sources. We're talking about sustained G-loading that lasts tens of seconds at near-peak intensity, presumably with considerable additional vibration from the turbulence. Per Wikipedia, "Only the most motivated volunteers were capable" of reaching this 20G level in testing for 10s duration with a steady centrifuge, and more than that is suggested to result in injury and/or LOC.Don't assume the EDL takes as long as a manned capsules entry, deceleration can be VERY fast if your coming in steep it just makes for very high g-forces and a tank can tolerate that.
The crew can tolerate a 20G entry.
Call it 10G peak eyeballs-in and assume some of the crew will black out or come down with concussions. For existing spacecraft like Soyuz, 8.2G is an emergency condition encountered only during a ballistic reentry, after something goes wrong with separation, and the Shuttle gets more like 3G.
Eyeballs-in may not always be perfectly achievable either. A reusable lander has to decelerate at a variety of orientations, and it may not be weight-feasible to have a pivoting chair. 20G's applied for a very short time at a slight angle to eyeballs-in will cause a red-out or a black-out by pushing enough G's at a person in the wrong direction, where they have low tolerance.
I'm not even trying to take into account deconditioning; A known unknown. But there's too many reasons to be conservative with G-loading already. If your mission needs 20G reentry, redesign your mission.
The forums HMXHMX may not agree with you.
Eyeballs-in may not always be perfectly achievable either. A reusable lander has to decelerate at a variety of orientations, and it may not be weight-feasible to have a pivoting chair. 20G's applied for a very short time at a slight angle to eyeballs-in will cause a red-out or a black-out by pushing enough G's at a person in the wrong direction, where they have low tolerance.
10G is marginal but probably safely achievable assuming you screen/test passengers on the ground before missions. If there's lots of vibration or not a lot of weight to spend on anti-G accommodations or deconditioning is a major stressor, even that's chancy.
In cases where we're reentering from high orbit (definitely Mars fast transits, maybe also Mars slow transits and GTO) without advanced techniques like MAC, it may be advantageous (the additional radiation and uncertainty may be outweighed by the advantages in spacecraft design tradespace) to do a two-stage entry, with an aerocapture to an intermediate elliptical orbit near LEO (or a suborbit with one end below LEO), then a control burn to target the landing zone, then a final entry the rest of the way.
My understanding is, that proper rocket thrust could cause that heated and compressed air is moved from engine compartment in front of the vehicle and plasma slide around vehicle, this what I try describe on picture.Don't assume the EDL takes as long as a manned capsules entry, deceleration can be VERY fast if your coming in steep it just makes for very high g-forces and a tank can tolerate that.
It still has to be shielded; it will be exposed to at least 60 seconds of plasma blast hotter than Raptor exhaust temps. It's going to be WAY hotter than a Falcon 9 S1 entry, which is already pretty toasty.
A separate cargo module attached to the main MCT craft that would have a heavy, "single use" TPS shield for the aerocapture, which is afterwards detached and either automatically landed at the primary landing site, or remote piloted down...
The main craft would then descend and land nearby the cargo module. The advantage here is that a cargo module could take higher G loads and more punishment than the fragile human cargo.
The alternative idea I had was to Aerocapture, then reenter with both craft and cargo module through the "7 minutes of hell" then detach the two, cargo module coming down via parachute with a Russian style retrorocket cushioning in the final few meters, and the main craft descending via retropropulsion. This latter idea keeps both craft and cargo near to each other, while protecting the main craft's TPS for the reentry at Earth.
Yes, but... it's still really hot and still heats up the vehicle through radiative transfer. Compression heating goes up with the velocity^3, so orbital entry at Mach 25 has the heating rate is 30 times more than a hot F9 S1 at Mach 8. It also takes a lot longer to slow down to Mach 3 from Mach 25 than it does from Mach 8, so the vehicle is soaking heat for at least 60 seconds (and can be 10x that) instead of around 20 seconds for an F9 S1.My understanding is, that proper rocket thrust could cause that heated and compressed air is moved from engine compartment in front of the vehicle and plasma slide around vehicle, this what I try describe on picture.Don't assume the EDL takes as long as a manned capsules entry, deceleration can be VERY fast if your coming in steep it just makes for very high g-forces and a tank can tolerate that.
It still has to be shielded; it will be exposed to at least 60 seconds of plasma blast hotter than Raptor exhaust temps. It's going to be WAY hotter than a Falcon 9 S1 entry, which is already pretty toasty.
I think, this is what Spacex found out during his retropropulsion events.
Don't assume the EDL takes as long as a manned capsules entry, deceleration can be VERY fast if your coming in steep it just makes for very high g-forces and a tank can tolerate that.
It still has to be shielded; it will be exposed to at least 60 seconds of plasma blast hotter than Raptor exhaust temps. It's going to be WAY hotter than a Falcon 9 S1 entry, which is already pretty toasty.
I'd imagine MCT would land automatically rather than human-piloted, so blackouts from G-forces may not be relevant. You'd need to keep below the level where any lasting harm is done, of course, but brief unconsciousness itself might not matter.
OK, while computers and software have advanced a lot since Apollo 11, I'm still not quite sure I'd want to put my complete faith and my life in the trust of a machine that could easily glitch out because of a stray cosmic ray.
Interesting idea. Is there any evidence that they run the F9 S1 fuel rich during entry?
Are you proposing that this would work for manned interplanetary entries as well? How much fuel would have to be burned for a 13 km/s entry, and what decelerations are reasonable for manned entries? How many engines are what throttle settings?
and propulsion will only be able to take out 1 km/sec of that.
and propulsion will only be able to take out 1 km/sec of that.
I have been wondering what that exactly means. Is that 1km/s the landing burn or do they need to brake 1km/s before they can enter the atmosphere and the landing burn will be separate? I had anticipated at the time he means the latter but may be completely wrong.
Propellent cost in orbit is significantly different than propellent cost at stage 1 separation; and propellent cost to GTO or escape is even more. Even the largest estimates for BFR only have a GTO payload in the 50t range, and on a second stage the entry propellent trades for payload at a 1:1 ratio.
4 Raptors will burn through over 1000 kg/s of propellent, even when throttled to 40%. A 10G orbital entry needs about 60 seconds to lose 6000 m/s and thus needs about 60 tonnes of fuel for that maneuver; an entry from GTO needs to lose almost 9000 m/s over 90 second, and needs nearly 90 tonnes of fuel; and entry from 12+ km/s interplanetary velocities needs over 120t of fuel for a burn of more than 120 seconds.
When an ablative heatshield for return from GTO or interplanetary velocities is only about 15% of dry mass, why would it possibly make sense to carry more than 100% of the vehicle's dry mass (and more than its payload capacity to GTO) in fuel for the same purpose?
When an ablative heatshield for return from GTO or interplanetary velocities is only about 15% of dry mass, why would it possibly make sense to carry more than 100% of the vehicle's dry mass (and more than its payload capacity to GTO) in fuel for the same purpose?
Same reason the F9 first stage uses propulsion rather then heat-shields to re-enter, orientation demands base first entry and it is simpler to both engineer and operate which means minimal cost.
Also how do you calculate mass flow rate for Raptor?
Your asking me to prove it is more EFFICIENT, that is completely different from saying it is SIMPLER. I have NEVER claimed it is more efficient, but you keep confusing the thouse concepts.
A heat shield has many benefits, and will be employed to maximize those obvious benefits (without adding complexity and risk to the design such as an inflatable HAID) and retro-propulsion will do the rest*.
A heat shield has many benefits, and will be employed to maximize those obvious benefits (without adding complexity and risk to the design such as an inflatable HAID) and retro-propulsion will do the rest*.
A HIAD is just a more efficient heatshield. It is a simple concept and it has been shown to work. This NASA video shows the application of a HIAD for the safe return of a cargo spaceship from the ISS. It is not a great leap to imagine the same technology being used for the return of an F9S2.
https://www.nasa.gov/directorates/spacetech/game_changing_development/HIAD/HEART-Desig-Concept.html
Adding a HAID to the flamey end of a F9 is not simple. It has to deploy from the engine compartment/octaweb over hot engines between burns and then get out of the way for the landing burn.
Efficient, maybe. Simple, NOT.
Adding a HAID to the flamey end of a F9 is not simple. It has to deploy from the engine compartment/octaweb over hot engines between burns and then get out of the way for the landing burn.
Efficient, maybe. Simple, NOT.
Who said to add it to the flamey end?
https://forum.nasaspaceflight.com/index.php?topic=37808.msg1565000#msg1565000
Just pack it in the interstage with some nitrogen bottles.
A HIAD is just a more efficient heatshield. It is a simple concept and it has been shown to work. This NASA video shows the application of a HIAD for the safe return of a cargo spaceship from the ISS. It is not a great leap to imagine the same technology being used for the return of an F9S2.
https://www.nasa.gov/directorates/spacetech/game_changing_development/HIAD/HEART-Desig-Concept.html
How many of the cargo return proposals/designs are using HIADs?
NASA PowerPoint/video does not an operating system make.
How many of the cargo return proposals/designs are using HIADs?
NASA PowerPoint/video does not an operating system make.
HIAD is way beyond PowerPoint presentations. Google is of course your friend, but there have been successful tests of the technology, and it's potential applications also extend to super heavy Mars landers.
How many of the cargo return proposals/designs are using HIADs?
NASA PowerPoint/video does not an operating system make.
HIAD is way beyond PowerPoint presentations. Google is of course your friend, but there have been successful tests of the technology, and it's potential applications also extend to super heavy Mars landers.
How many super heavy Mars landers have used HIADs? Why is that?
Google "If a hammer is the only tool"
2020 and exomars aren't going to be significantly larger than MSL.
This is a dumb way to argue. Please bring better game.
There's some indirect evidence that it may be one of several options they'll consider.2020 and exomars aren't going to be significantly larger than MSL.
This is a dumb way to argue. Please bring better game.
Fine.
What evidence is there that MCT will use a HIAD?
There's some indirect evidence that it may be one of several options they'll consider.2020 and exomars aren't going to be significantly larger than MSL.
This is a dumb way to argue. Please bring better game.
Fine.
What evidence is there that MCT will use a HIAD?
Your asking me to prove it is more EFFICIENT, that is completely different from saying it is SIMPLER. I have NEVER claimed it is more efficient, but you keep confusing the thouse concepts.
Your asking me to prove it is more EFFICIENT, that is completely different from saying it is SIMPLER. I have NEVER claimed it is more efficient, but you keep confusing the thouse concepts.
I'm merely asking for evidence that's it's physically possible, not proof that it's optimal. It does appear possible for a stage that only enters from LEO or suborbital trajectories.
GTO and BLEO orbits would require a 3rd stage with a lot of Delta V, but I suppose that's not an issue is you assume BFS is a third stage and SEP tugs will fill some of those roles eventually.
It's nearly certain that some sort of deployable drag enhancement device will be used for MCT. None will be used on Red Dragon, which is not at all a reason to think it won't be used on MCT.There's some indirect evidence that it may be one of several options they'll consider.2020 and exomars aren't going to be significantly larger than MSL.
This is a dumb way to argue. Please bring better game.
Fine.
What evidence is there that MCT will use a HIAD?
I don't disagree that it will be or already was considered; there are many possibilities that will be discarded as not optimum or unworkable.
Certainly no evidence that it will be used on Red Dragon, which would be the obvious opportunity to test it...
I think there's a reasonable chance no such device will be used for MCT, for the sole reason that reusable controlled reliable powered landing is neither fully compatible nor fully benefitted by HIAD;
I think there's a reasonable chance no such device will be used for MCT, for the sole reason that reusable controlled reliable powered landing is neither fully compatible nor fully benefitted by HIAD;
I thought HIAD was a technology for relatively small scale EDL (NASA@Mars)? Here we are talking about ships that are as large as projected HIAD brake area, so why waste mass?
HIAD is scalable to enormous sizes. It's not just for small scale EDL.I think there's a reasonable chance no such device will be used for MCT, for the sole reason that reusable controlled reliable powered landing is neither fully compatible nor fully benefitted by HIAD;
I thought HIAD was a technology for relatively small scale EDL (NASA@Mars)? Here we are talking about ships that are as large as projected HIAD brake area, so why waste mass?
HIAD is scalable to enormous sizes. It's not just for small scale EDL.
HIAD is scalable to enormous sizes. It's not just for small scale EDL.
Developed by NASA in that way? I only see them developing HIAD as a stage before using parachutes and parachutes limiting the downmass to Mars severely. Combining HIAD with SRP seems a serious headache.
Yes, I am fully aware of that. But how much more than the 1t limit they have reached with Curiosity? No word on that in there. Will it support mannedERVMAV?
It's nearly certain that some sort of deployable drag enhancement device will be used for MCT.
It's nearly certain that some sort of deployable drag enhancement device will be used for MCT.
IMO deployables will be limited to control surfaces much smaller than the heatshield. I tend to doubt that they will use inflatables, just because it's unnecessary complexity if they are willing to spend 1 km/s of delta-v on EDL.
HIAD is scalable to enormous sizes. It's not just for small scale EDL.I think there's a reasonable chance no such device will be used for MCT, for the sole reason that reusable controlled reliable powered landing is neither fully compatible nor fully benefitted by HIAD;
I thought HIAD was a technology for relatively small scale EDL (NASA@Mars)? Here we are talking about ships that are as large as projected HIAD brake area, so why waste mass?
It's nearly certain that some sort of deployable drag enhancement device will be used for MCT.
IMO deployables will be limited to control surfaces much smaller than the heatshield. I tend to doubt that they will use inflatables, just because it's unnecessary complexity if they are willing to spend 1 km/s of delta-v on EDL.
The trade is that to reduce the maximum g forces, you need to increase the available negative lift. If you have more negative lift, you can decelerate more gradually, higher in the Martian atmosphere. It will be interesting to see what SpaceX deem the maximum acceptable g forces. I suspect that will determine the size of any enhancements.
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I would guess retrorockets is the far easier, cheaper method to scale tough.
How far can NASA's current HIAD scale?
..Perhaps the trade may include the mass of PICA-X running down one side of the vehicle versus the mass of HIAD to perform the same EDL.
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Yes, I can't see a company invest in added mass and (fragile) complexity unless it is really necessary. By the way, wouldn't the side of a large craft offer a lot of negative lift? [Not an aerodynamist ... obviously.]
A very thorough presentation of ideas for BFR/MCT from "coborop" at Reddit:
https://www.reddit.com/r/spacex/comments/4wks2h/fanmade_mct_and_bfr_architecture_cad_and_math/
Beautiful renders!:
https://imgur.com/a/2k10I
It's nearly certain that some sort of deployable drag enhancement device will be used for MCT.
IMO deployables will be limited to control surfaces much smaller than the heatshield. I tend to doubt that they will use inflatables, just because it's unnecessary complexity if they are willing to spend 1 km/s of delta-v on EDL.
The trade is that to reduce the maximum g forces, you need to increase the available negative lift. If you have more negative lift, you can decelerate more gradually, higher in the Martian atmosphere. It will be interesting to see what SpaceX deem the maximum acceptable g forces. I suspect that will determine the size of any enhancements.
C. G. Niederstrasser @RocketScient1st
Shotwell - just shipped first Raptor engine to Texas last night. #SpaceX #smallsat
FYI - This may interest people here - first MCT related hardware now exists in pre-production form: Raptor :DAs I mentioned on another thread this is 1 year earlier than we have assumed for such engine and tests. This makes our estimates for NET BFR first launch test conservative by as much as 1 year moving the possibility of a BFR to exist as early as 2019/2020.
https://twitter.com/RocketScient1st/status/763063393745940481QuoteC. G. Niederstrasser @RocketScient1st
Shotwell - just shipped first Raptor engine to Texas last night. #SpaceX #smallsat
No pictures are available yet, but this probably means that Musk will show pictures of real hardware in his September presentation.
FYI - This may interest people here - first MCT related hardware now exists in pre-production form: Raptor :DAs I mentioned on another thread this is 1 year earlier than we have assumed for such engine and tests. This makes our estimates for NET BFR first launch test conservative by as much as 1 year moving the possibility of a BFR to exist as early as 2019/2020.
https://twitter.com/RocketScient1st/status/763063393745940481QuoteC. G. Niederstrasser @RocketScient1st
Shotwell - just shipped first Raptor engine to Texas last night. #SpaceX #smallsat
No pictures are available yet, but this probably means that Musk will show pictures of real hardware in his September presentation.
SpaceX is moving faster than we expect. This is surprise number 2 for this year in relation to Mars after the RD mission date.
A very thorough presentation of ideas for BFR/MCT from "coborop" at Reddit:
https://www.reddit.com/r/spacex/comments/4wks2h/fanmade_mct_and_bfr_architecture_cad_and_math/
Beautiful renders!:
https://imgur.com/a/2k10I
Interesting. Has the upper stage co-fire @ takeoff to increase T/W. Problem with Rvacs there.
What I like is someone not showing some BFR/BFS that is some ridiculous height like past Reddit posts I've seen elsewhere. A >10m wide BFR gets very heavy quickly with propellant. I'm confident that the reveal will be short and stout.
It's nearly certain that some sort of deployable drag enhancement device will be used for MCT. None will be used on Red Dragon, which is not at all a reason to think it won't be used on MCT.There's some indirect evidence that it may be one of several options they'll consider.2020 and exomars aren't going to be significantly larger than MSL.
This is a dumb way to argue. Please bring better game.
Fine.
What evidence is there that MCT will use a HIAD?
I don't disagree that it will be or already was considered; there are many possibilities that will be discarded as not optimum or unworkable.
Certainly no evidence that it will be used on Red Dragon, which would be the obvious opportunity to test it...
I don't have any particular belief that SpaceX will use something like HIAD. I don't envision they will. But you dismiss the possibility far too easily.
Figuring out how to do this on Mars is going to be a challenge, part of why we need retro-propulsion on Mars, more predictable than aerodynamics.
Beautiful renders!:
https://imgur.com/a/2k10I
Sea Level ISP 350 seconds seems too high so the BFR will not perform to that level assumed.
On the other hand you don't need the same Km/sec to RTLS as you launched because some of the launch velocity vector is vertical. It's not pure horizontal velocity. Yes, you need some additional -X delta V to actually fly back.
Sea Level ISP 350 seconds seems too high so the BFR will not perform to that level assumed.
On the other hand you don't need the same Km/sec to RTLS as you launched because some of the launch velocity vector is vertical. It's not pure horizontal velocity. Yes, you need some additional -X delta V to actually fly back.
I think that's 350s for the SL engine operating in vacuum (e.g. booster stage near MECO). SpaceX is proposing 321s at SL and 363s in vacuum for the SL engine, so 350s is probably conservative.
Ugh, not looking forward to half a dozen pages of naming speculation. THE most boring aspect of this whole endeavor.
It's nearly certain that some sort of deployable drag enhancement device will be used for MCT. None will be used on Red Dragon, which is not at all a reason to think it won't be used on MCT.There's some indirect evidence that it may be one of several options they'll consider.2020 and exomars aren't going to be significantly larger than MSL.
This is a dumb way to argue. Please bring better game.
Fine.
What evidence is there that MCT will use a HIAD?
I don't disagree that it will be or already was considered; there are many possibilities that will be discarded as not optimum or unworkable.
Certainly no evidence that it will be used on Red Dragon, which would be the obvious opportunity to test it...
I don't have any particular belief that SpaceX will use something like HIAD. I don't envision they will. But you dismiss the possibility far too easily.
Are HIADs considered to be a ballute?
Are HIADs considered to be a ballute?
They probably shouldn't be. Although ballutes are capable of generating drag, they are not capable of generating negative lift. Also they, parachutes and aerofoils only appear to be intended for supersonic and subsonic flight regimes, rather than hypersonic.
SpaceX is switching to carbon fibers from aluminum as it develops heavy rockets for carrying people and large quantities of material. A lighter body would allow more cargo to be loaded, which would cut transport costs.http://asia.nikkei.com/Business/Deals/Toray-carbon-fiber-to-carry-SpaceX-s-Mars-ambitions
Carbon fiber bodies/stages instead of aluminum... Advances the state-of-the-art in rocketry.
Helps mass reduction and secondary radiation effects simultaneously.QuoteSpaceX is switching to carbon fibers from aluminum as it develops heavy rockets for carrying people and large quantities of material. A lighter body would allow more cargo to be loaded, which would cut transport costs.http://asia.nikkei.com/Business/Deals/Toray-carbon-fiber-to-carry-SpaceX-s-Mars-ambitions
Combined with fast transits, rad problem could be largely mitigated.
PMF could (will) be best ever by far...
Update: On Tuesday evening SpaceX would not confirm that a large deal had been reached. "Toray is one of a number of suppliers we work with to meet our carbon fiber needs for Falcon rocket and Dragon spacecraft production, and we haven’t announced any new agreements at this time," a company spokesman told Ars. "As our business continues to grow, the amount of carbon fiber we use may continue to grow."
Does all-composite construction favor any particular architecture choices? SpaceX seems to put quite a bit of effort into TPS on the composite Falcon interstage compared to the Al-Li tanks. I wonder if they will need TPS basically everywhere on a composite booster and especially on an orbiter.Al-Li hates heat basically just as much, the difference is the interstage gets full blast from the upper stage.
Does all-composite construction favor any particular architecture choices? SpaceX seems to put quite a bit of effort into TPS on the composite Falcon interstage compared to the Al-Li tanks. I wonder if they will need TPS basically everywhere on a composite booster and especially on an orbiter.
Does all-composite construction favor any particular architecture choices? SpaceX seems to put quite a bit of effort into TPS on the composite Falcon interstage compared to the Al-Li tanks. I wonder if they will need TPS basically everywhere on a composite booster and especially on an orbiter.
Does all-composite construction favor any particular architecture choices? SpaceX seems to put quite a bit of effort into TPS on the composite Falcon interstage compared to the Al-Li tanks. I wonder if they will need TPS basically everywhere on a composite booster and especially on an orbiter.
Al-Li favours bi-conic fairing like shapes while composites could enable a much wider range of shapes. I've long thought that the BFS could have a semi-lifting body shape, one without winglets, but a non-circular cross section.
It's not going to have wings or land horizontally like that. I'd put a lot of money on that.
Simply entering an atmosphere sideways as opposed to vertically is only going to approximately double the area at best and it requires the vehicle to be load bearing in a second direction, which is a significant mass penalty for a vehicle this large.That's plausible as well. But a doubling or tripling of the area is not insignificant.
I suspect we will see the vehicle equipped with some form of radial expansion mechanism that give it a much larger cross sectional area and entry will be in a vertical orientation either base first or nose first to keep the forces of entry on the single strong vertical axis that is used during launch.
Simply entering an atmosphere sideways as opposed to vertically is only going to approximately double the area at best and it requires the vehicle to be load bearing in a second direction, which is a significant mass penalty for a vehicle this large.
I suspect we will see the vehicle equipped with some form of radial expansion mechanism that give it a much larger cross sectional area and entry will be in a vertical orientation either base first or nose first to keep the forces of entry on the single strong vertical axis that is used during launch.
I suspect we will see the vehicle equipped with some form of radial expansion mechanism that give it a much larger cross sectional area and entry will be in a vertical orientation either base first or nose first to keep the forces of entry on the single strong vertical axis that is used during launch.
First, frictional dissipation of velocity using heat-shields, parachutes, HIAD, you name it, Always comes out more mass efficient then propellant, we would always choose to trade landing propellant for structural mass in the vehicle because the ratio is likely to be in excess of 10 to 1 in favor of the structural element over propellant. Complexity is a downside, yes but if it means not having to be loaded with large amounts of propellant on atmospheric entry that is a big deal because that saves mass twice in two high delta v maneuvers, departure from Earth and departure from Mars.It's not entirely necessary to increase drag or ballistic coefficient for Earth return, so un-deploying a deployable drag device on Mars and then dragging it back to Earth is actually less efficient. The extra propellant capacity is actually useful for the return transit burns.
Second SRP is only usable in the later parts of EDL after most velocity has already been lost through friction. The problem is that the vehicle is going to have so high of a ballistic coefficient that it would strike the ground on mars before even reaching the speed with sufficient remaining altitude that SRP can handle, their must be something to lower the coefficient.
Simply entering an atmosphere sideways as opposed to vertically is only going to approximately double the area at best ...
Second SRP is only usable in the later parts of EDL after most velocity has already been lost through friction. The problem is that the vehicle is going to have so high of a ballistic coefficient that it would strike the ground on mars before even reaching the speed with sufficient remaining altitude that SRP can handle, their must be something to lower the coefficient.
Red Dragon will (try) to baseline a low Lift/Drag, high ballistic coefficient Mars entry. Assuming that succeeds in getting to the SRP regime with an appreciable amount of payload (i.e. it doesn't go "splat" first), than a larger vehicle with similar aerodynamic characteristics should be feasible.
Red Dragon's gee forces are not too high. Compare a Soyuz ballistic entry after 6 months in orbit.
I suspect we will see the vehicle equipped with some form of radial expansion mechanism that give it a much larger cross sectional area and entry will be in a vertical orientation either base first or nose first to keep the forces of entry on the single strong vertical axis that is used during launch.
Red Dragon's gee forces are not too high. Compare a Soyuz ballistic entry after 6 months in orbit.
Those forces are not nominal and are routinely cause the cosmonauts to black out, your always siting superlatives and extremes as if they are usable as nominal conditions but that is flawed because their would be no margin left for any off nominal event.
A radial area expansion system dose not rule out lift generation, in fact it makes it easier to control the asymmetry of the vehicle and control the amount of lift generated which is highly advantageous given the variable density of mars atmosphere. A lifting body or bi-conic will generated high lift but at the cost of reducing drag as these bodies to not enter fully horizontally and while the lift is needed for g-force reduction from what I've read you lose too much drag which raises the ballistic coefficient back into dangerous territory. A fairly simple radial expansion in the form of fold down petals and connecting carbon fiber webbing between them could easily triple the vehicle radius to 40 meters which would give a comparable ballistic coefficient to the MSL.
Spacecraft | Diameter m | Area m^2 | Volume m^3 | Mass mT | Ballistic coefficient | Sidewall angle ° | Peak gs | Entry velocity km/s |
MSL | 4.5 | 15.21 | ? | 3.3 | 148 | 40 | 15 | 5.8 |
Soyuz | 2.2 | 3.8 | 4 | 3 | 789 | 7 | 4 | 7.6 |
Cargo dragon | 3.7 | 10.75 | 10 | 7.2 | 515 | 15 | 4.5 | 7.6 |
Red dragon | 3.7 | 10.75 | 10 | 10 | 715 | 15 | 7.4 | 9.6 |
BFS | 22 | 380.13 | 4000 | 357 | 715 | 30? | 4 | 12 |
...easily triple the vehicle radius to 40 meters which would give a comparable ballistic coefficient to the MSL.
MSL had 3300 kg at EI behind a 10.75 m2 heatshield, or 307 kg/m2, and it slowed to below 450 m/s before deploying chutes.
MSL had 3300 kg at EI behind a 10.75 m2 heatshield, or 307 kg/m2, and it slowed to below 450 m/s before deploying chutes.
Are you sure? Wikipedia quotes 2,401 kg for the EDL system, and a 4.5m diameter heatshield, giving an area of 15.21 m^2. That's a β of 158 kg/m^2, although I've seen 110 kg/m^2 quoted elsewhere.
Can you list the assumptions behind these values? The volume and mass seem rather high to me.
Spacecraft Diameter m Area m^2 Volume m^3 Mass mT Ballistic coefficient Sidewall angle ° Peak gs Entry velocity km/s BFS 22 380.13 4000 357 930 30? 4 12
Can you list the assumptions behind these values? The volume and mass seem rather high to me.
MSL had 3300 kg at EI behind a 10.75 m2 heatshield, or 307 kg/m2, and it slowed to below 450 m/s before deploying chutes.
Are you sure? Wikipedia quotes 2,401 kg for the EDL system, and a 4.5m diameter heatshield, giving an area of 15.21 m^2. That's a β of 158 kg/m^2, although I've seen 110 kg/m^2 quoted elsewhere.
You're right about the 4.5m MSL shield, I was thinking of the Viking shields.
But the 2.4t mass is for the EDL system ONLY (skycrane, fuel, chutes, shield, backshell), and doesn't include the 900 kg rover itself. 3300 kg over 15.2 m2 gives a beta of 217 kg/m2
Edit: here's the link to JPL's listed masses for MSL... http://mars.jpl.nasa.gov/msl/mission/spacecraft/
I don't find 22 m diameter to be plausible, it's far wider then even the widest speculated size of the BFR, but a vehicle 15 m in diameter with fold down flaps the same size as the current F9 landing legs which are 18 m long would give a 51 m diameter, and the legs only mass 2,500 kg on the F9R so this looks eminently practical and it far exceeds the diameter that a monolithic vehicle could be launched as.
I'm also operating under the assumption of a 200 mt entry mass and a much more modest propellant quantity sufficient for 800 m/s.
Regarding ballistic coefficients, I've attached a table that compares several designs, including a hypothetical BFS with the same ballistic coefficient as Red Dragon.
Spacecraft Diameter m Area m^2 Volume m^3 Mass mT Ballistic coefficient Sidewall angle ° Peak gs Entry velocity km/s MSL 4.5 15.21 ? 3.3 148 40 15 5.8 Soyuz 2.2 3.8 4 3 789 7 4 7.6 Cargo dragon 3.7 10.75 10 7.2 515 15 4.5 7.6 Red dragon 3.7 10.75 10 10 715 15 7.4 9.6 BFS 22 380.13 4000 357 715 30? 4 12
Regarding ballistic coefficients, I've attached a table that compares several designs, including a hypothetical BFS with the same ballistic coefficient as Red Dragon.
Spacecraft Diameter m Area m^2 Volume m^3 Mass mT Ballistic coefficient Sidewall angle ° Peak gs Entry velocity km/s MSL 4.5 15.21 ? 3.3 148 40 15 5.8 Soyuz 2.2 3.8 4 3 789 7 4 7.6 Cargo dragon 3.7 10.75 10 7.2 515 15 4.5 7.6 Red dragon 3.7 10.75 10 10 715 15 7.4 9.6 BFS 22 380.13 4000 357 715 30? 4 12
Is there any particular reason why Mars atmospheric entry velocities would be as high as 9-12km/s for the Red Dragon and BFS? I would expect them to be on the same order of past Mars missions (that is, 5-7km/s)...
Regarding ballistic coefficients, I've attached a table that compares several designs, including a hypothetical BFS with the same ballistic coefficient as Red Dragon.
Spacecraft Diameter m Area m^2 Volume m^3 Mass mT Ballistic coefficient Sidewall angle ° Peak gs Entry velocity km/s MSL 4.5 15.21 ? 3.3 148 40 15 5.8 Soyuz 2.2 3.8 4 3 789 7 4 7.6 Cargo dragon 3.7 10.75 10 7.2 515 15 4.5 7.6 Red dragon 3.7 10.75 10 10 715 15 7.4 9.6 BFS 22 380.13 4000 357 715 30? 4 12
Is there any particular reason why Mars atmospheric entry velocities would be as high as 9-12km/s for the Red Dragon and BFS? I would expect them to be on the same order of past Mars missions (that is, 5-7km/s)...
As to BFS Musk has said that he wants shorter trip times, so shorter trip times equates to higher Km/sec entry velocity
I suspect we will see the vehicle equipped with some form of radial expansion mechanism that give it a much larger cross sectional area and entry will be in a vertical orientation either base first or nose first to keep the forces of entry on the single strong vertical axis that is used during launch.I agree with you and when I saw the "Battlestar Galactica" french design... :o Should be much more simple.
- Landing a 9 motor configuration. Or at least vertical landing similar to what they are mastering.How does this monster get to LEO? 9 Raptors are VERY underpowered for Earth launch at sea level, but rather overpowered for anything else.
- During launch, the abort unit is a "Super Dragon" with similar shape to the V2. The trunk include the living quarters and services.How much mass are you putting through abort? What engines would be used for abort? IFAIK SpaceX is not working on a engine that would be suitable to abort that large a capsule from Earth launch.
- In case of main stage issues during EDL, the "Super Dragon" could detach and follow similar trajectory of Red Dragon. Survival in the capsule until new ascent spaceship arrives.I can see where EDL abort capability would be nice, but I can't see any way it's practical. No Earth return craft has ever had that, for good reasons. Any EDL failure that's not survivable on a spacecraft probably isn't survivable in the abort capsule either. There's really no reason to separate the Super Dragon from it's Trunk. Land them together, launch them together, and abort them together if needed.
- The legs have double function: semi expanded work as a frame for a deceleration fabric that increase the surface. (If active hydraulics can also provide control).You don't need grid fins for Mars, and behind a heatshield they won't do anything useful on Earth, so move teh heatshield up to the Trunk. Where does the fabric go when the legs are folded up? Is it stowed? Can it be re-stowed automatically on Mars for relaunch?
- The legs do provide enough clearance for relaunch: no need to move the rocket to vertical position. Just land on top of a leveled clean pad.A separate "cargo bay" isn't practical; the cargo bay and hab should be the same size and shape to fit in the same mold lines. Almost everything will need to be shipped pressurized, and the pressure vessel should to be removed at Mars to serve as storage/habitat. There's no point in returning your shipping container to Earth.
Regarding ballistic coefficients, I've attached a table that compares several designs, including a hypothetical BFS with the same ballistic coefficient as Red Dragon.
Spacecraft Diameter m Area m^2 Volume m^3 Mass mT Ballistic coefficient Sidewall angle ° Peak gs Entry velocity km/s MSL 4.5 15.21 ? 3.3 148 40 15 5.8 Soyuz 2.2 3.8 4 3 789 7 4 7.6 Cargo dragon 3.7 10.75 10 7.2 515 15 4.5 7.6 Red dragon 3.7 10.75 10 10 715 15 7.4 9.6 BFS 22 380.13 4000 357 715 30? 4 12
Is there any particular reason why Mars atmospheric entry velocities would be as high as 9-12km/s for the Red Dragon and BFS? I would expect them to be on the same order of past Mars missions (that is, 5-7km/s)...
As to BFS Musk has said that he wants shorter trip times, so shorter trip times equates to higher Km/sec entry velocity
Not necessarily. You just need more fuel to slow down before you get to entry.
Overall, some really cool ideas, but you could work on how they are configured. Do you have any masses for any of the vehicles, or required fuel masses calculated?The main objective of the post was to get feedback about the concept of decelerator deployed with legs. Done all the design after work. Many artistic licenses. If you want to feed with data, I´m happy to modify and improve. Everything is parametric.
How does this monster get to LEO? 9 Raptors are VERY underpowered for Earth launch at sea level, but rather overpowered for anything else.It´s a second stage BFS. And one stage to orbit form Mars + earth injection.
Are you planning on having those 9 engines facing into the orbital reentry blast? They won't like that, unless they are firing the whole time... and firing the whole time uses too much fuel. ´Was thinking on extensions on the low part of the legs that cover the external engines when deployed. Depends on how many engines are needed in total and how many only for the EDL? 9 were random to visualize the concept (an octaweb would simplify their design).
That lower section should have more engines, and just be a stage 1 booster. It only does sub-orbital EDL and doesn't need much shielding. Move the deployable heat shield up the the Super Dragon Trunk. The Trunk and Super Dragon are the unit that then would go to orbit, and with the deployable shield for reentry at Mars and Earth.Need to analyze this input
Move the "SuperDraco" pods down to the trunk sidewall and fire them nearly straight back, between the landing legs. They have to be able to fire with the shield folded or deployed. That's how it will get from staging to orbit/TMI (with the shield folded) and that's how it will do supersonic retro-propulsion (with the shield open).
How much mass are you putting through abort? What engines would be used for abort? IFAIK SpaceX is not working on a engine that would be suitable to abort that large a capsule from Earth launch.Those huge super dracos are artistic. Could be a much smaller capsule powered by SD like your 6m Dragon 3. But I expect there will be an abort function during launch and EDL.
Also, why do you need to abort the whole thing? MCT can easily carry a 7 man Dragon 2 (or later a hypothetical ~25 passenger, ~6 meter Dragon 3) on top, which can abort on SuperDracos and shuttle back to Earth immediately. Since ~4 refueling launches are needed per Mars injection, just send 1/4 of the people up on every launch.
The main objective of the post was to get feedback about the concept of decelerator deployed with legs. Done all the design after work. Many artistic licenses. If you want to feed with data, I´m happy to modify and improve. Everything is parametric.I can certainly provide some calculations. It would be helpful to know some of the dimensions you have modeled, such the main diameter of the stage.
It´s a second stage BFS. And one stage to orbit form Mars + earth injection.That good, but it's quite over sized for that. A 15m diameter subcooled methalox tank holds about 180 tonnes of propellant per meter of length, so for the 1000t or so of propellant it takes to put 100t of payload through TMI and EDL, you only need some 5 to 6 meter long tanks. The tanks on BFS are going to be pretty close to spherical and 10 to 12m diameter, not a long cylinder as you show. The long cylinder is exactly what the booster stage will look like.
Was thinking on extensions on the low part of the legs that cover the external engines when deployed. Depends on how many engines are needed in total and how many only for the EDL? 9 were random to visualize the concept (an octaweb would simplify their design).Depending on how massive it is, it will need somewhere between 2 to 6 vacuum Raptors on the second stage. Retropropulsion requires that the engines be as far outboard as possible, and if they are set back from the heatshield around the edges they don't need a cover - the bow shock will carry the hot plasma out and around them. Look at the SuperDracos on Dragon 2. That's exactly where you want to put your engines for EDL... or pretty much like this:
Move the "SuperDraco" pods down to the trunk sidewall and fire them nearly straight back, between the landing legs. They have to be able to fire with the shield folded or deployed. That's how it will get to orbit/TMI (with the shield folded) and that's how it will do supersonic retro-propulsion (with the shield open).
But I expect there will be an abort function during launch and EDL.Launch? Yes. EDL? I don't think there are a lot of scenarios where that's survivable with any form of abort. Maybe on early missions where you can fit the crew in a Dragon on top, but certainly not at the colonization level. Some risks are just inherent with no practical means for abort, and EDL has always been one of those. By the time they get past the "3.7m or 6m Dragon on top" type of missions, the architecture will have flown thousands of entries and have a very high confidence in reliability, 99.9% or greater.
Regarding ballistic coefficients, I've attached a table that compares several designs, including a hypothetical BFS with the same ballistic coefficient as Red Dragon.
Spacecraft Diameter m Area m^2 Volume m^3 Mass mT Ballistic coefficient Sidewall angle ° Peak gs Entry velocity km/s MSL 4.5 15.21 ? 3.3 148 40 15 5.8 Soyuz 2.2 3.8 4 3 789 7 4 7.6 Cargo dragon 3.7 10.75 10 7.2 515 15 4.5 7.6 Red dragon 3.7 10.75 10 10 715 15 7.4 9.6 BFS 22 380.13 4000 357 715 30? 4 12
Is there any particular reason why Mars atmospheric entry velocities would be as high as 9-12km/s for the Red Dragon and BFS? I would expect them to be on the same order of past Mars missions (that is, 5-7km/s)...
As to BFS Musk has said that he wants shorter trip times, so shorter trip times equates to higher Km/sec entry velocity
Not necessarily. You just need more fuel to slow down before you get to entry.
Not a great mission profile wasting valuable rocket equation exponential propellant when the atmosphere gives the delta V to you if you are very clever in your spacecraft design and guidance.
Lastly some estimates on mass, as noted earlier the leg system on F9 masses 2500 kg, carbon fiber fabric 5 layers thick totaling 1 kg/m^2 and covering the space between a 15 m base and a 51 m total diameter would be 1800 m and kg bringing total mass to 4.3 mt, a very modest amount that would be far lighter then carrying extra propellant. The propellant needed for just the last bit of deceleration and up to landing is going to need to do around 800 m/s based on extrapolation from Red Dragon, and that's going to come out to around 20 percent propellant fraction minimum. I'm estimating 200 mt at entry and 40 mt of landing propellant.
I have a totally crazy idea, feel free to shoot it down.You'd actually get much more performance by filling the tanks with methane than with hydrogen.
The idea is a Tri-propellant (but not really) MCT.
Assuming that:
1. Methalox engines are required for Mars-side ISRU.
2. Most payload mass will only go one way: Earth to Mars.
How much sense would it make to use Hydrolox engines for the Mars-bound half of the trip?
Basically, you use highly efficient Hydrolox engines to get to Mars, use ISRU there to fill the hydrogen tank with methane, and use Raptors to go back home (and I know, the tank volume ratios are problematic).
You may say, but SpaceX doesn't have Hydrolox engines. True, but Blue Origin does. :-)
You'd actually get much more performance by filling the tanks with methane than with hydrogen.
Hydrogen has terrible density, so you could fit a LOT more methane in there than hydrogen.
You'd actually get much more performance by filling the tanks with methane than with hydrogen.
Hydrogen has terrible density, so you could fit a LOT more methane in there than hydrogen.
This doesn't make sense. There is a reason Saturn 5, STS, SLS and the Centaur all use hydrolox...
Hydrogen already has a lot of negative design trades, and if you add in the requirement that the system also has to burn methane equally well, there's no good reason to use hydrogen at all.
And the tank volume ratios are more than slightly problematic. The hydrogen tank to put 180 tonnes through TMI and EDL would be more than 7 times larger in volume than the methane tank needed to return a 80 tonne vehicle from Mars surface to Earth's surface. It would be larger in volume than most estimates for the entire BFR booster.
And if you switched those stages over to methane with the same total tank volume, the stages would weigh more, but they'd give you a much higher performance. More payload to higher delta-v, GIVEN the same starting point.You'd actually get much more performance by filling the tanks with methane than with hydrogen.
Hydrogen has terrible density, so you could fit a LOT more methane in there than hydrogen.
This doesn't make sense. There is a reason Saturn 5, STS, SLS and the Centaur all use hydrolox...
Assuming that:False assumption.
1. Methalox engines are required for utilizing Mars-side ISRU.
2. Most payload mass will only go one way: Earth to Mars.
Assuming that:False assumption.
1. Methalox engines are required for utilizing Mars-side ISRU.
2. Most payload mass will only go one way: Earth to Mars.
Mars ISRU is not capturing existing methane as we do on earth. Mars ISRU envisions building the methane from water derived H2 and atmospheric CO2. If the system had H2 engines and tanks you'd refill them with H2 and skip the methane altogether.
In other words, if the trades worked in hydrogen's favor, there's nothing preventing them from building a hydrogen rocket without the methane. Since they have chosen to build a methane rocket, one might presume that the trades did not work in hydrogen's favor. Quite possibly for some of the reasons mentioned above.
You'd actually get much more performance by filling the tanks with methane than with hydrogen.
Hydrogen has terrible density, so you could fit a LOT more methane in there than hydrogen.
This doesn't make sense. There is a reason Saturn 5, STS, SLS and the Centaur all use hydrolox...
Thanks for demolishing my proposal. :-)It depends. If your initial mass isn't a constraint, then for the same tanks, you're better off with methane.
So, I am correct to understand that the increased mass of the tanks would eat all the performance gains from using Hydrolox?
Lastly some estimates on mass, as noted earlier the leg system on F9 masses 2500 kg, carbon fiber fabric 5 layers thick totaling 1 kg/m^2 and covering the space between a 15 m base and a 51 m total diameter would be 1800 m and kg bringing total mass to 4.3 mt, a very modest amount that would be far lighter then carrying extra propellant. The propellant needed for just the last bit of deceleration and up to landing is going to need to do around 800 m/s based on extrapolation from Red Dragon, and that's going to come out to around 20 percent propellant fraction minimum. I'm estimating 200 mt at entry and 40 mt of landing propellant.
Reducing dry mass will have priority since the architecture is constrained by the return leg. The main benefit of something like HIAD would be its lower mass compared to a mid L/D aeroshell.
Regarding decelerator devices...
Isn't it somewhat counter-intuitive to put them at the front of the vehicle? Wouldn't such a vehicle normally flip around to have the area of maximum resistance at the back?
I am thinking that any large-scale decelerators should be deployed at the rear of the vehicle, as a form of drogue chute. That would give a heatshield of vehicle diameter at the front and a large-diameter deployable decelerator trailing behind, to create a stable configuration. And no, I am not just talking about a parachute, because I think of a much more robust device, be it with fins or petals or whatever the aerodynamicists can come up with.
Approaching 30 days from what (we hope) is the big reveal, I thought it a good time to revisit and post revised BFR/MCT speculation before any info leaks out. Trying to stay within the parameters of what Musk has said as I best understand. A TSTO vehicle launched by a re-useable, single core BFR that puts the BFS a.k.a. the MCT into LEO where it is re-fueled, travels to and lands on Mars where it is again refueled for the journey back to Earth carrying a quarter of the outbound “cargo” mass. The outbound cargo masses 100 tonnes which I assume means either cargo or people or a combination thereof. BFS/MCT mass not included in the 100T.
Myriad unknowns led by the dry mass of the BFS. Rocket equation dictates various mass assumptions here can produce wildly different answers.
My predictions, metric unless otherwise stated:
1.Entire launch vehicle BFR+BFS masses under 5,000T. Guestimate ~4,500T.
2.BFS dry mass < 100T, my pick is 85T carbon composites BUT heavier than some predictions because ruggedized to allow for minimal maintenance.
3.BFR absolutely > 10m diameter to fit enough engines. Likely between 12.5 and 15m. My guess 15m. Allows addition of more engines in the future.
4.My guestimate BFR+BFS stack <100m height. Certainly <125m.
5.Sticking with the “over 230T” Raptor thrust Elon mentioned, I get 25-27 engines. My guestimate is 26 with “over 230T” as 235T in my spreadsheet. Around 13.5 million Lbs force.
Engine # most likely wrong because…
6.Predict that Raptor engine design goal thrust changed to higher than 230T previously stated, but only by several 10s of tonnes, not hundreds.
7.BFS with 5 Rvac engines
8.RTLS minimizes cost, turnaround time, effort. Changed my opinion from max payload ASDS for those reasons. Just make the BFR bigger. Stages low and slow ~2.2 Km/sec. “Easy” recovery & re-flight vs F9 GTO flights.
9.Initial BFR test flights likely equipped with less engines and less payload.
10.Large crew volume design >2,000m3. Initial flights with less people & people space but more cargo space.
11.Initial crewed Mars mission will carry 6-12 people. 10 is my latest #. Why?
NASA & other nations will buy seats.
http://forum.nasaspaceflight.com/index.php?topic=40683.msg1557261#msg1557261
12.SEP still under development awaits later opposition cargo transits
13.BFS will have “exotic” upper mounted engines for rough terrain Mars landing &takeoff (just echoing others’ analysis here)
14.BFS will be a lifting body for EDL, but not a scaled up Dragon capsule shape. It will look badass.
You know we’re totally screwed trying to predict Musk because he already warned us,
“When it looks more like an alien dreadnought, that’s when you know you’ve won.”
I’ve attached a spreadsheet showing different assumptions, BFS mass, etc.
Anyone else want to update their speculations?
Regarding decelerator devices...That's an option, but would have lower drag and wouldn't protect the rest of the vehicle.
Isn't it somewhat counter-intuitive to put them at the front of the vehicle? Wouldn't such a vehicle normally flip around to have the area of maximum resistance at the back?
I am thinking that any large-scale decelerators should be deployed at the rear of the vehicle, as a form of drogue chute. That would give a heatshield of vehicle diameter at the front and a large-diameter deployable decelerator trailing behind, to create a stable configuration. And no, I am not just talking about a parachute, because I think of a much more robust device, be it with fins or petals or whatever the aerodynamicists can come up with.
As for exotic upper mounted engines, I used to think they were required, but now I think terminal landing will be done on raptors alone (again to save dry mass). For this to be possible from the first BFS landing will require scouting an appropriate flat solid rock surface and a dragon placed rover to sweep away any rocks/pebbles and place beacons to allow for guidance to the landing pad within a few meters.
Whilst I agree that Raptors alone would save dry mass, as well as removing the dependency on two successfully serialised miracles, if a Red Dragon could safely land on an unprepared Martian surface, why couldn't a BFS, especially if it resembles a scaled up Dragon?
Whilst I agree that Raptors alone would save dry mass, as well as removing the dependency on two successfully serialised miracles, if a Red Dragon could safely land on an unprepared Martian surface, why couldn't a BFS, especially if it resembles a scaled up Dragon?
A Red Dragon's heat shield and motors don't need to survive past touchdown. Safe for it is no destructive motor failures. Gouging up the heatshield and throwing rocks up into motors is fine as long as it can land once. A vehicle that has to take off again and do another re-entry needs an intact heatshield and motors for Earth return.
If the Raptors were sufficiently canted, wouldn't they direct ejecta away from the spaceship?
Approaching 30 days from what (we hope) is the big reveal, I thought it a good time to revisit and post revised BFR/MCT speculation before any info leaks out. [...]
Anyone else want to update their speculations?
I would hardly consider a set of pressure fed engines firing on touch-down to be a 'miracle', it's a normal and standard part of the Soyuz capsule landing after all.
If the Raptors were sufficiently canted, wouldn't they direct ejecta away from the spaceship?
Launching directly back to Earth is the highest demanding part of the mission¹, those first moments of take-off from the Martian surface will have the highest fuel load and hence require maximum thrust. You can't afford cosine losses of canted engines. Otherwise the stage is overpowered for every other part of the missions, and hence wasting mass.²
Shotwell dropped the name Falcon 20 for the BFR. >
Regarding decelerator devices...
Isn't it somewhat counter-intuitive to put them at the front of the vehicle? Wouldn't such a vehicle normally flip around to have the area of maximum resistance at the back?
Regarding decelerator devices...
Isn't it somewhat counter-intuitive to put them at the front of the vehicle? Wouldn't such a vehicle normally flip around to have the area of maximum resistance at the back?
With HIAD the center of gravity must be sufficiently low and offset to create lift. It's really no different from a capsule in that respect.
Some of the concepts here which show a long cylinder on a small inflatable aren't realistic IMO.
I would hardly consider a set of pressure fed engines firing on touch-down to be a 'miracle', it's a normal and standard part of the Soyuz capsule landing after all.
Every time a mission critical event in deep space succeeds, it is a (tongue firmly in cheek) 'miracle'. To paraphrase Elon, part of good spaceship design is minimising the number of serialised miracles that need to occur.
Regarding decelerator devices...
Isn't it somewhat counter-intuitive to put them at the front of the vehicle? Wouldn't such a vehicle normally flip around to have the area of maximum resistance at the back?
With HIAD the center of gravity must be sufficiently low and offset to create lift. It's really no different from a capsule in that respect.
Some of the concepts here which show a long cylinder on a small inflatable aren't realistic IMO.
For high velocity Mars EDL, yes, I agree. However, for Earth entry, especially from LEO, lift is less of a requirement. For example IRVE-3, which is quite a long cylinder, has been successfully demonstrated.
There must be sufficient distance between center of gravity and center of pressure for stability.
Please bear in mind, I don't expect the engines to be canted for the entire mission.
Whilst I agree that Raptors alone would save dry mass, as well as removing the dependency on two successfully serialised miracles, if a Red Dragon could safely land on an unprepared Martian surface, why couldn't a BFS, especially if it resembles a scaled up Dragon?
A Red Dragon's heat shield and motors don't need to survive past touchdown. Safe for it is no destructive motor failures. Gouging up the heatshield and throwing rocks up into motors is fine as long as it can land once. A vehicle that has to take off again and do another re-entry needs an intact heatshield and motors for Earth return.
Sure, the BFS has to take off again, but why are you so sure rocks would be thrown up into the motors? If the Raptors were sufficiently canted, wouldn't they direct ejecta away from the spaceship?
I'm wondering about a rover with a solar concentrator to melt a thin layer of surface material in place as a solidified layer. Presuming this layer didn't crack too badly it should make a good surface for a landing pad or road.Don't know but intuition says anything less than one m thick under MCT will be flying FOD.
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ADDED: Your question deserves a better answer so I looked up what NASA scientists have found in their studies. The following statement is a verbatim copy of a summary that addresses the problem. It was a section of the Mars Design Reference Architecture, Addendum A, published in 2009:
"5.9.1 Summary and recommendations
The predictions and recommendations for a 40-t spacecraft on Mars are described in summary in this section. The next section of the report will then explain in detail how these predictions were obtained.
The engine exhaust plume from a 40-t lander on Mars will blow dust, sand, gravel, and even rocks up to about 7 cm in diameter at high velocity. These ejecta will cause significant damage to any hardware that is already placed on the martian surface within the blast radius. However, the blast radius is modest, extending out to approximately 1 km. The largest debris is accelerated by the plume to lower velocities and, thus, falls closer to the landing site; and the smallest particles are attenuated by the martian atmosphere, also falling closer to the landing. Thus, maintaining the distance of about 1 km between the landing site and any existing surface assets will completely solve this problem for all sizes of debris.
A second concern arises because the exhaust from the large engines will form deep, narrow craters that are directly beneath each of their nozzles, and these craters will redirect the supersonic jet of gas with sand and rocks up toward the landing spacecraft. This has been demonstrated in large-scale engine tests in sand and clay (Alexander, et al, 1966) 25, small-scale experiments (Metzger, 2007) 26, numerical simulations (Liever, et al, 2007) 27, and soil dynamics analysis (see section 5.10.2.3), so there is no question that this will occur. It did not occur in the Apollo and Viking missions because the thrust was lower and/or because the lunar regolith had higher shear strength and less permeability than martian soil. These variables have been taken into consideration in this report. An example of a small-scale test is provided in figure 5-55. The impact of debris striking the lander will be sufficient to cause damage to the lander, possibly resulting in LOM and LOC, and therefore must be prevented. Of special concern is damage to the engine nozzles, because with a multiple-engine lander the debris that is ejected by one engine will be aimed directly at the other engines. One mitigation approach is to add shielding to the spacecraft to block the debris. This will increase the mass of the lander and, therefore, reduce the mass of the payload by approximately 1 t."
I'm pretty sure there's a link to the article around here somewhere - I do vaguely remember reading it, But I can't find it right now. Very short version is high-thrust motors close to the surface dig really deep holes that throw junk back at the motors. Either have a prepared pad, keep the motors far away, or build something that can take the beating. The Curiosity rover ended up with quite a lot of big gravel all over the top and visible gouges in the ground from the skycrane motors. And those were at the top end of 25 feet of cable and canted somewhat.
... but why are you so sure rocks would be thrown up into the motors? If the Raptors were sufficiently canted, wouldn't they direct ejecta away from the spaceship?
Quote"5.9.1 Summary and recommendations
[...] It did not occur in the Apollo and Viking missions because the thrust was lower and/or because the lunar regolith had higher shear strength and less permeability than martian soil. [...] "
Staging requires more thrust than any other point in the BFS mission. It's in a bigger gravity well, with a full fuel load and a full payload.
Someone suggested the possibility that MCT/BFS is just a massively scaled-up Crew Dragon. While I'm sure other things are happening than that, it did get me to thinking that arranging BFS's Raptors in the same "covered dugout" configuration that Crew Dragon uses for its Super Draco's might be the winning guess.
Staging requires more thrust than any other point in the BFS mission. It's in a bigger gravity well, with a full fuel load and a full payload.
We don't know the final delta-v of the first stage at separation. If it works anything like conventional launches, then the delta-v for a second stage to LEO is much less than the delta-v from Mars surface to direct Earth return.
Actually we do. Elon said in an interview that the BFR is just there to compensate for the difference in the two longest poles of the mission. The ΔV to LEO is about 9.5 km/s, including gravity and aerodynamic losses. Mars-Earth is about 7km/s. BFR will provide about 3.5km/s, so if the BFS can generate 7.5 km/s, the remainders are the payloads.
However, that just reiterates my point. If the BFR is designed to exactly balance the difference between Mars-Earth and Earth-LEO, then if you make the Mars launch less efficient, you need more fuel, bigger tanks, different volume trades, then you are also changing every other part of the equation.
I expect that BFR will stage even slower than Falcon 9 for efficient RTLS. Which requires BFS to do even more than the F9 upper stage. I don't see the engines of BFS fire at any angle. They do everything to improve efficiency, light weight, highly optimized engines with high ISP. They won't give any of that away for canted engines.
Which of course makes the methods of EDL and avoiding debris hurled up more interesting. If not prepared too much in advance, find a hard flat surface. Such was mentioned as a requirement for a landing site in the NASA workshop on potential landing sites on Mars.
There are only cosine losses for the first few seconds of the Mars launch, and never for launch to LEO. The canting is adjustable.
There are only cosine losses for the first few seconds of the Mars launch, and never for launch to LEO. The canting is adjustable.
CITATION NEEDED!!! What is your epistemology?
Rocket engine gimbal is only on the order of 10 degrees, nothing like the kind of 45 degree plus that's being proposed here has ever been done and would complicate the engines thrust structure and plumping hugely.
A secondary set of touch down engines would be ignited BEFORE shutting down the Raptor engines and in the event that they don't start you would ride the Raptor thrust all the way to the ground in a hover-slam and risk what ever damage may come as that's preferable to crashing. To suggest that one engines is just shut off before any validation of the next engines is to practice in straw-man engineering, Musk even went into detail about how Dragon v2 will test it's engines first before diverting from a splash-down trajectory so a parachute landing can be executed in case of any engine failure and I am sure they will not forget this concept on Mars.
The decelerator designs are generally inherently stable in either direction once they have commenced their descent into the atmosphere. The hypersonic heating rate is proportional to velocity cubed, and to the inverse of the square root of the atmospheric density divided by the leading edge radius.
HeatingRate = 1.83e-4 * Math.Pow(speed, 3) * Math.Sqrt(atmosphericDensity / HeatingRadius());
So, the larger the leading edge radius (i.e. the blunter the body), the lower the heating rate. Although perhaps counter-intuitive, this is why the decelerator is best positioned at the front of the vehicle. I've attached a few design alternatives below.
I would hardly consider a set of pressure fed engines firing on touch-down to be a 'miracle', it's a normal and standard part of the Soyuz capsule landing after all.
I expect that BFR will stage even slower than Falcon 9 for efficient RTLS. Which requires BFS to do even more than the F9 upper stage. I don't see the engines of BFS fire at any angle. They do everything to improve efficiency, light weight, highly optimized engines with high ISP. They won't give any of that away for canted engines.I think you are right BFS will give same or less staging speed then Falcon 9.
Which of course makes the methods of EDL and avoiding debris hurled up more interesting. If not prepared too much in advance, find a hard flat surface. Such was mentioned as a requirement for a landing site in the NASA workshop on potential landing sites on Mars.
There are only cosine losses for the first few seconds of the Mars launch, and never for launch to LEO. The canting is adjustable.
CITATION NEEDED!!! What is your epistemology?
Rocket engine gimbal is only on the order of 10 degrees, nothing like the kind of 45 degree plus that's being proposed here has ever been done ...
... and would complicate the engines thrust structure and plumping hugely.
Ok lets make that comparison, touch down engines would be as simple as a dozen super draco engines around the nose of the vehicle. [...]
That is sufficient thrust to hover the likely landing mass of around 100 tons after the main propulsion system brings the vehicle to zero ground velocity at some safe height.
These engines would mass less then a ton and consume just slightly more propellant then the Raptor engines would to do the same job, the lower touch down speed should allow for a much lighter landing gear system as well which is nothing to sneeze at as landing gear is historically 10 percent of landed mass.
Making an engine gimbal a huge amount will basically require a completely second thrust structure to keep the inward thrust vectors from crushing the vehicle, thrust structures are usually more massive then the actual engines they hold. Propellant lines need to use bellows to allow them to flex but they are like springs, the bend radius is large so to get a greater deflection the engine needs to be at the end of a longer and wider thrust structure, the mass of all this would quickly become prohibitive.
And as for the shuttle main engine achieving 10 degrees of movement please look at some cut-aways and see how deeply all the mess of plumbing goes in the vehicle just to do that much. The shuttle engine had to be designed with that level of gimbaling, so unless SpaceX had the whole mars EDL system figured out before starting engine development it's very unlikely that Raptor would ever be capable of more then the typical single digit gimbaling of a standard rocket engine.
Ok lets make that comparison, touch down engines would be as simple as a dozen super draco engines around the nose of the vehicle. [...]
None of which helps for the re-launch.
...The idea is to allow the canting of the entire engine mount in a single degree of freedom.
...
Canting is required anyway by SRP, so there is no extra engineering required for landing.
Legs need to be very strong, don't they. What's the weight on Mars of the fuel needed to launch direct to Earth?
Would deployable thrust deflectors make sense? Rather than gimballing the engines, just redirect the exhaust away from the vehicle at 30 or 45 degrees. I'm thinking an ablative-coated panel set just below the engines (which I think will be mounded on the sidewalls), that pops out into the exhaust stream.
Would deployable thrust deflectors make sense? Rather than gimballing the engines, just redirect the exhaust away from the vehicle at 30 or 45 degrees. I'm thinking an ablative-coated panel set just below the engines (which I think will be mounded on the sidewalls), that pops out into the exhaust stream.
If you mean the deflectors would be attached to the vehicle itself, then at 45°, the vertical thrust would push down of the panels (and hence the ship) exactly as much as it pushes up; resulting in zero vertical net force. The sideways thrust would presumably be balanced in each horizontal direction by ensuring the panels point in different direction, so you wouldn't get any horizontal movement either. Essentially is would be a very elaborate scheme to stop the rockets from doing anything at all. It would just sit there, burning away, until something gave out.
If you mean to mount the deflectors on the surface, you are essentially talking about a flame-trench. Which means a large concrete construction, along with some way of suspending the vehicle over it, so a launch platform/gantry of some kind. Which wouldn't be a good place to land (too much stuff in the way), so you'd need a vehicle capable of hoisting and transporting the BFR.... basically you've added an awful lot of hardware.
Ok lets make that comparison, touch down engines would be as simple as a dozen super draco engines around the nose of the vehicle. [...]
None of which helps for the re-launch.
I assume you mean weight, the mass would be at least 250mT, and much higher if the ship was to carry enough fuel to return directly to LMO or Earth. MCT is a fuel rich architecture, allowing for many mission profiles.
The landing velocity would be close to zero in either scenario.
Who said anything about gimbaling more than a few degrees? The idea is to allow the canting of the entire engine mount in a single degree of freedom. Canting is required anyway by SRP, so there is no extra engineering required for landing.
Making an engine gimbal a huge amount will basically require a completely second thrust structure to keep the inward thrust vectors from crushing the vehicle, thrust structures are usually more massive then the actual engines they hold. Propellant lines need to use bellows to allow them to flex but they are like springs, the bend radius is large so to get a greater deflection the engine needs to be at the end of a longer and wider thrust structure, the mass of all this would quickly become prohibitive.
To launch you can't be canting, you need vertical thrust or your engine efficiency is ruinous.
First the vehicle is not going to be full of propellant at landing, that is nonsense.
Why would a lander want to land with enough propellant to relaunch? Landing prop for earth return is prohibitive, so where would it go with landed prop, and what would it do when it got there?
Why would a lander want to land with enough propellant to relaunch? Landing prop for earth return is prohibitive, so where would it go with landed prop, and what would it do when it got there?
To reduce the time taken to generate enough for relaunch? Use as a raw material?
To launch you can't be canting, you need vertical thrust or your engine efficiency is ruinous.
You can't be canting for the entire LMO and TEI burns, but what if it was just for the first few seconds, enough to avoid the risk of ejecta impacting the ship?
The heaviest estimate I've seen for BFS dry mass + payload from Mars to Earth is about 160mT. Assuming 7.5 km/s of ΔV is required, that is a takeoff mass of about 1280mT. If there are 5 Raptors, each having around 270mT of thrust, that is 1360mT of thrust at takeoff. For Mars gravity (0.38g), that is a thrust to weight ratio of 1360 / 1280 * 0.38 = 2.8, far more than necessary to achieve liftoff. Allowing a worst case T/W of 1.15, cos^-1(1.15/2.8) = 65.7°, probably more than enough to mitigate the ejecta risk. After a few seconds of flight, the cant could be reduced to zero, and the flight continue with no further cosine losses to TEI.First the vehicle is not going to be full of propellant at landing, that is nonsense.
As shown above, the ship wouldn't need to be completely full of propellant to relaunch, and it certainly wouldn't be necessary once ISRU is established, but why is it a nonsense? Do you have any supporting calculations?
I think ideally a 1/2 to 1/3 scale Raptor would be ideal for the BFS, but I think SpaceX has indicated they are going to only develop one size of this guy.
I think ideally a 1/2 to 1/3 scale Raptor would be ideal for the BFS, but I think SpaceX has indicated they are going to only develop one size of this guy.
Interesting that you should say that. In Noël Bakhtian's thesis 'Drag Augmentation via Supersonic Retropropulsion for Atmospheric Deceleration' she finds that the optimum number of engines for Mars SRP is actually 15 engines in a peripheral array, canted outboard. 15 1/3 scale Raptors would give the same thrust as 5 full scale.
Impaler, I'm drawn to your suggestion re smaller, more responsive engines for Mars landing. But I think Super Dracos aren't the answer for a couple of reasons.
First, you still need the Raptors for Mars ascent (unless you are arguing that the Dracos will be able to lift a fully fueled BFS from Mars surface). And the Raptors will have to be canted, so as to avoid backsplash during ascent. This is exactly the same problem as on descent. So if you have to solve it for ascent, why include a separate method for descent?
Second, the hypergolic fuel used on the Super Dracos is not easily made via ISRU. That means that the fuel used for Earth landing must be carried from Earth to Mars, then back to Earth again. This would really cut into the available payload landed on Mars (given the dictatorship of the rocket equation). Right?
I think ideally a 1/2 to 1/3 scale Raptor would be ideal for the BFS, but I think SpaceX has indicated they are going to only develop one size of this guy.
Isn't the simplest solution to foreign object damage an engine that is robust enough to survive the odd pebble? And maybe some redundancy in case one engine breaks?
I have a really hard time to imagine how something could enter an active combustion chamber against the flow of the combusted fuel.
I think the answer to this issue is landing on bedrock or somewhere else deemed reasonably safe.
I think the answer to this issue is landing on bedrock or somewhere else deemed reasonably safe.
Is "landing on bedrock" compatible with ISRU, e.g. would you have access to water?
Isn't the simplest solution to foreign object damage an engine that is robust enough to survive the odd pebble? And maybe some redundancy in case one engine breaks?
They did this above concrete, which spalls. Short turnaround time.Isn't the simplest solution to foreign object damage an engine that is robust enough to survive the odd pebble? And maybe some redundancy in case one engine breaks?
The odd pebble?....
So picture something with 2/3rd more power than what is depicted here, sitting on landing legs, a metre or two above hard flat surface. Even bedrock would spall. ...
But you can't test your engines before a launch into a direct Earth return trajectory, because there's no infrastructure that can withstand the test (or hold down the vehicle.)
Static fires are fine (if maybe expensive), provided you don't put your payload on top.But you can't test your engines before a launch into a direct Earth return trajectory, because there's no infrastructure that can withstand the test (or hold down the vehicle.)
Actually you could test the engines if they were sufficiently canted, and the ship was fully fuelled. For a T/W of 1 at launch, cos^-1(1/2.8) = 69.1°, so set the cant at 70° or more, and you could perform a short full thrust static fire.
Admittedly, I have some trepidation regarding static fires at the moment.
Making the vehicle strong enough for a static fire in two different directions is a somewhat questionable design choice, even if we were to assume that the gimbals for canting posed no weight issues or technical unknowns.I wasn't suggesting that, by the way. I don't yet believe in the idea of extreme canting for MCT.
Making the vehicle strong enough for a static fire in two different directions is a somewhat questionable design choice, even if we were to assume that the gimbals for canting posed no weight issues or technical unknowns.
I've tried to fit landing some methalox in, but even for a small portion of the vehicle (an 'escape pod'), the math doesn't work for return to orbit (much less return to Earth), it swamps other payload. If you could wring 4.5km/s out of an escape pod, you could return to MLO and a waiting lifeboat there for return to Earth. But that's too much. Landing mostly dry just works a lot better. Every kilogram of propellant landed reduces the ISRU gear landed by 1kg, which reduces the propellant generated by many kilograms per synod.
Now... that's not to say landing *completely* dry is the preferred option. Liquid propellants are useful for a number of things other than full-on Mars ascent. Solar storm shielding, EDL mass redistribution maneuvers, and high-thrust RCS landing maneuvers are all made considerably easier if you pack a sizable quantity of, say, monomethyl hydrazine on board.
Who here still believes that some new pad at Cape Canaveral or some future pad at Boca Chica is still viable for BFR launches? Given the recent pad anomaly, I believe that more stringent safety zones will force the BFR to be launched off shore, assuming any US based launch site, which I do because of ITAR.
There are shallow seas offshore from both SX launch sites. At first a platform could be anchored to the sea floor and serviced by barges and hydrofoils while launch rates remained low. A very expensive causeway could later be constructed to support higher flight rates.
I do not feel that on shore launch remains an option assuming the BFR is really ~2x Saturn V size or more.
The noise dB problem remains as RTLS sonic booms propagate tens of miles with little attenuation.
Who here still believes that some new pad at Cape Canaveral or some future pad at Boca Chica is still viable for BFR launches? Given the recent pad anomaly, I believe that more stringent safety zones will force the BFR to be launched off shore, assuming any US based launch site, which I do because of ITAR....This has no bearing on BFR.
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I do not feel that on shore launch remains an option assuming the BFR is really ~2x Saturn V size or more.
...
I agree an offshore platform is the best option for the BFR we've been discussing. But with the schedule Elon mentioned there really isn't enough time to build a pad or platform before initial BFR testing. Maybe the BFR isn't going to be as big as we think. Perhaps the first generation BFS will be 50 tonnes cargo to Mars instead of 100 tonnes, requiring a BFR about the size of a Saturn V. That's still impressive. Pad 39A can handle that. We'll find out in a few weeks.
My predictions are still going against the grain, I think we will see mostly BFR details with only modest information on the actual mars transit and landing and what ever details are revealed of that phase of the mission will be subject to change as SpaceX actually gets landing experience on Mars, it's too soon to completely commit to the mode or design of the mars landing vehicles at this time.
BFR will be narrower then the 15m diameter many have speculated, I think closer to 12m diameter as the thrust density of the Raptor engine will be high and a narrower vehicle makes for a lighter more efficient thrust structure to transmit the thrust of many engines to the tank side walls and makes for lower air resistance on assent as well. The first stage will be recovered right from the first orbital flight and will be well tested on it's own in sub-orbital flights with a mass-simulating 2nd stage used for testing.
A recoverable second stage with vacuum raptor engines, probably 6 so it has engine out capability, will be used along with a conventional payload fairing to launch large payloads of high volume and make the BFR a direct competitor to SLS. This 2nd stage will also act as the LEO propellant tanker with just the attachment of a minimal nosecone. Recovery of the 2nd stage still be slightly experimental at the time the first orbital launches and the fist few launches will be refine the recovery process for the 2nd stage and several may be lost in the process. The 2nd stage will orbit the Earth one or more times before recovery and might not land at the launch site.
Overall vehicle height is be over 200 ft at the top of the 2nd stage, with payload fairing around 300 ft. Any mars bound spacecraft will be introduced several years after BFR itself is flying and will not fly the super-direct flight plan of departing from LEO directly to mars surface and then from Mars surface directly back to Earth surface, this requires too much delta-V and propellant production on Mars, rather their will be a refueling step somewhere between likely in a high Earth orbit and/or high Mars orbit.
Agree completely. MCT/BFS will be more conceptual and under development than big 1st stage.My totally uninformed and unsubstantiated pet hope is a sudden plan for a falcon heavy replacement:
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Anyone else have predictions/updates?
How short could you plausibly make the full-diameter BFR? Im just thinking a lot shorter and symmetrically remove some engines.If it is only needs 12 meter diameter instead of 15 that is great. However dumb my idea was at 15m, it is probably a bit less dumb at 12m :)
I'm not sure how valid this is, but the tanks have to be some shape. And that shape will probably have hemispherical endcaps for mass efficiency. You can't make a tank with hemispherical endcaps smaller than a sphere, and it is far easier to build a cylinder than other shapes. If this is a crippling constraint, it seems like you probably have to use multiple smaller subscale tanks.Agree completely. MCT/BFS will be more conceptual and under development than big 1st stage.My totally uninformed and unsubstantiated pet hope is a sudden plan for a falcon heavy replacement:
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Anyone else have predictions/updates?How short could you plausibly make the full-diameter BFR? Im just thinking a lot shorter and symmetrically remove some engines.If it is only needs 12 meter diameter instead of 15 that is great. However dumb my idea was at 15m, it is probably a bit less dumb at 12m :)
Thx 4 reply philw1776 I have a few counter and clarifying points.
Lastly I believe it may be possible to send tankers from Earth to a high mars orbit, have them offload propellants and still return to Earth orbit either via chemical propulsion or SEP so all vehicles are recovered. This would allow the Mars landing and relaunch vehicles to be much lighter at liftoff from Mars, using something like 1/3rd as much propellant, then refill in orbit and make a fast return to Earth.
The exact details of how and where refueling occurs aren't clear yet but what I'm sure of is that the direct return approach is not viable, the fixation many have with it is analogous to the fixation on direct earth return in the early days of Apollo, it's attractive for it's simplicity but it a bridge too far in a lot of engineering areas.
Thx 4 reply philw1776 I have a few counter and clarifying points.
Planning for future upgrades to BFR is interesting, but I'm doubtful that a vehicles diameter is made larger then necessary to allow for the insertion of extra engines at a later date. First the extra engines and thrust would require a complete redesign of the plumbing structures and thrust structures which is a significant part of the first stages total design.
As the Big Reveal is nearly upon us (hopefully!), I think that someone (not me) should generate an official Mars MCT Architecture Prediction contest. This would be composed of a list of questions with multiple choice answers, with a fixed value for each question. Prizes TBD.SL thrust of each Raptor on BFR.
One problem is that L2 members may have 'insider information' on this topic. So I guess they will have to be just guided by their own conscience as to their level of participation.
Anyway, I'll kick it off with a few suggested questions (but will be happy for someone else to officially take this over and come up with the canonical list, as well as being judge, jury, etc.)
1) Overall Launch Architecture
a) MCT is composed simply of a BFR 1st stage and BFS 2nd stage/spacecraft
b) Boost phase consists of 2 stages, which put the BFS into orbit
c) Other: 3rd stage, 'half' stages, drop tanks, etc.
2) Number of Raptor Engines on BFR (1st stage)
3) Diameter of BFR (1st stage)
4) Total Raptor 1st stage thrust (sl)
5) LAS Architecture
a) No LAS - BFS is the escape mechanism
b) Traditional LAS - above BFS and is nominally jettisoned during launch phase
c) BFS contains smaller 'ejection pod' where humans reside during launch
d) Other, non-traditional LAS design
6) Shape and Landing Mode of BFS
a) Capsule (perhaps elongated), w/ TPS on base
b) Cylindrical or biconic - horizontal landing
c) Cylindrical or biconic - vertical landing
d) Other
7) Mars and Earth return
a) BFS does direct entry into Mars and Earth atmosphere
b) BFS does orbital capture before performing entry burn and landing
c) Same as b, but upon Earth return, stays in orbit for next synod
8) Use of non-chemical thrust
a) Not part of the plan
b) Will use SEP for some/all of the big transits
c) All chemical for now, but plans to incorporate SEP down the road
Can anyone think of more/better questions?
Thx 4 reply philw1776 I have a few counter and clarifying points.
Lastly I believe it may be possible to send tankers from Earth to a high mars orbit, have them offload propellants and still return to Earth orbit either via chemical propulsion or SEP so all vehicles are recovered. This would allow the Mars landing and relaunch vehicles to be much lighter at liftoff from Mars, using something like 1/3rd as much propellant, then refill in orbit and make a fast return to Earth.
The exact details of how and where refueling occurs aren't clear yet but what I'm sure of is that the direct return approach is not viable, the fixation many have with it is analogous to the fixation on direct earth return in the early days of Apollo, it's attractive for it's simplicity but it a bridge too far in a lot of engineering areas.
Gwynne has said, "We're looking at SEP" so this is a possibility.
On the downside if it's SEP, it is yet another vehicle to be developed and tested.
I haven't looked deeply at delta Vs, masses and run the rocket equation to comment cogently on the feasibility of SEP, etc. tankers to HMO & back. I'm certain that SX has done so in their architectural bakeoffs.
As the Big Reveal is nearly upon us (hopefully!), I think that someone (not me) should generate an official Mars MCT Architecture Prediction contest. This would be composed of a list of questions with multiple choice answers, with a fixed value for each question. Prizes TBD.SL thrust of each Raptor on BFR.
One problem is that L2 members may have 'insider information' on this topic. So I guess they will have to be just guided by their own conscience as to their level of participation.
Anyway, I'll kick it off with a few suggested questions (but will be happy for someone else to officially take this over and come up with the canonical list, as well as being judge, jury, etc.)
1) Overall Launch Architecture
a) MCT is composed simply of a BFR 1st stage and BFS 2nd stage/spacecraft
b) Boost phase consists of 2 stages, which put the BFS into orbit
c) Other: 3rd stage, 'half' stages, drop tanks, etc.
2) Number of Raptor Engines on BFR (1st stage)
3) Diameter of BFR (1st stage)
4) Total Raptor 1st stage thrust (sl)
5) LAS Architecture
a) No LAS - BFS is the escape mechanism
b) Traditional LAS - above BFS and is nominally jettisoned during launch phase
c) BFS contains smaller 'ejection pod' where humans reside during launch
d) Other, non-traditional LAS design
6) Shape and Landing Mode of BFS
a) Capsule (perhaps elongated), w/ TPS on base
b) Cylindrical or biconic - horizontal landing
c) Cylindrical or biconic - vertical landing
d) Other
7) Mars and Earth return
a) BFS does direct entry into Mars and Earth atmosphere
b) BFS does orbital capture before performing entry burn and landing
c) Same as b, but upon Earth return, stays in orbit for next synod
8) Use of non-chemical thrust
a) Not part of the plan
b) Will use SEP for some/all of the big transits
c) All chemical for now, but plans to incorporate SEP down the road
Can anyone think of more/better questions?
First I find focusing on vehicle count as a measure of difficulty is a poor metric which I think has led many people astray, when the mission requires a high performance in a number of conflicting areas trying to achieve all these performance values out of one vehicle is usually far more difficult then breaking up the problem using specialized vehicle.I thought this too, and you still have a point about conflicting performance requirements, but the reason why Apollo was broken up into separate spacecraft instead of a direct ascent architecture was because expensive expendable rockets forced constraints on mass, so there were compromises.
That has been the lesson of all space travel, it's why rockets have stages, it's why the LEM rather then the command module landed on the moon and it's why I consistently foresee more complex mission modalities then the majority.
IMLEO is no longer the primary constraint
Gwynne has said, "We're looking at SEP" so this is a possibility.Random thought: What about a vehicle that can do missions with small crews without SEP, but you have a plan to add SEP later (as a stage, cycler or L2 propellant tug, whatever) when you have multiple flights to spread the costs over and want closer to that 500k/passenger goal, and also to achieve that fast return to earth for immediate reuse?
On the downside if it's SEP, it is yet another vehicle to be developed and tested.
Sure, just change how many tankers are used to refuel (& where) the BFS in orbit. SEP can be to push the fuel to a higher energy orbit before refueling.Gwynne has said, "We're looking at SEP" so this is a possibility.Random thought: What about a vehicle that can do missions with small crews without SEP, but you have a plan to add SEP later (as a stage, cycler or L2 propellant tug, whatever) when you have multiple flights to spread the costs over and want closer to that 500k/passenger goal, and also to achieve that fast return to earth for immediate reuse?
On the downside if it's SEP, it is yet another vehicle to be developed and tested.
Sure, just change how many tankers are used to refuel (& where) the BFS in orbit. SEP can be to push the fuel to a higher energy orbit before refueling.Gwynne has said, "We're looking at SEP" so this is a possibility.Random thought: What about a vehicle that can do missions with small crews without SEP, but you have a plan to add SEP later (as a stage, cycler or L2 propellant tug, whatever) when you have multiple flights to spread the costs over and want closer to that 500k/passenger goal, and also to achieve that fast return to earth for immediate reuse?
On the downside if it's SEP, it is yet another vehicle to be developed and tested.
Sure, just change how many tankers are used to refuel (& where) the BFS in orbit. SEP can be to push the fuel to a higher energy orbit before refueling.Gwynne has said, "We're looking at SEP" so this is a possibility.Random thought: What about a vehicle that can do missions with small crews without SEP, but you have a plan to add SEP later (as a stage, cycler or L2 propellant tug, whatever) when you have multiple flights to spread the costs over and want closer to that 500k/passenger goal, and also to achieve that fast return to earth for immediate reuse?
On the downside if it's SEP, it is yet another vehicle to be developed and tested.
Staging fuel at EML1/2 or similar HEO with SEP would cut refueling launches from Earth by about 50% for crewed flights.
SEP tugs could also put cargo landers directly into TMI from LEO, and then quickly return to LEO. That would cut refueling Earth launches for cargo flights by 65% to 80%. And there will be a lot more cargo flights than crew flights.
Elon Musk ✔ @elonmusk
Turns out MCT can go well beyond Mars, so will need a new name…
8:22 PM - 16 Sep 2016
Hi Mr. Musk,
how about "Interplanetary Colonial Transporter" ?Quote from: Elon MuskTurns out MCT can go well beyond Mars, so will need a new name…
@kzmakino sounds about right
Is he talking about the asteroid belt or something, I can't see the landing vehicle with it's normal atmospheric EDL profile being appropriate on any other planetary body in the solar system other then Venus.
>
Is he talking about the asteroid belt or something, I can't see the landing vehicle with it's normal atmospheric EDL profile being appropriate on any other planetary body in the solar system other then Venus.Maybe for Asteroid missions long duration SEP burns as orbital matching delta-V.
While it could retro-propulsive land on an airless body it would be complete overkill to do that on an asteroid with a vehicle with so much thrust and when something like RCS would be able to land and take off. If he thinks the system will be this flexible it strongly suggests an in space transit vehicle is being used.
From what I can find the Delta V to get to main belt asteroids is around 4.5 km/s burn at LEO and another 4.5 km/s to direct land because their is no significant gravity well or anything to brake against.
Kazunori Makino 12 minutes ago
https://twitter.com/kzmakino/status/776954933820002305Quote from: @kzmakinoHi Mr. Musk,
how about "Interplanetary Colonial Transporter" ?Quote from: Elon MuskTurns out MCT can go well beyond Mars, so will need a new name…
https://twitter.com/elonmusk/status/776956202936782850Quote from: Elon Musk
@kzmakino sounds about right
Takes some cleverness. Maybe use tankers or SEP to get further. MCT should be able to do Ceres just fine. Maybe Callisto, too. Europa is doable but requires more radiation hardening.Kazunori Makino 12 minutes ago
https://twitter.com/kzmakino/status/776954933820002305Quote from: @kzmakinoHi Mr. Musk,
how about "Interplanetary Colonial Transporter" ?Quote from: Elon MuskTurns out MCT can go well beyond Mars, so will need a new name…
https://twitter.com/elonmusk/status/776956202936782850Quote from: Elon Musk
@kzmakino sounds about right
ICT - Columbia Would be very nice.
Back on topic, going doesn't mean it'll be able to land. They are going to be bound to the capacity of manufacturing methane for fueling their ship. I'm a bit skeptical that MCT is going to go anywhere is not Mars or even the Moon. But hey, amazing we are even having this conversation.
Enterprise, Defiant, Voyager, Discovery (new CBS show)....or how about Serenity?
Is he talking about the asteroid belt or something, I can't see the landing vehicle with it's normal atmospheric EDL profile being appropriate on any other planetary body in the solar system other then Venus.
While it could retro-propulsive land on an airless body it would be complete overkill to do that on an asteroid with a vehicle with so much thrust and when something like RCS would be able to land and take off. If he thinks the system will be this flexible it strongly suggests an in space transit vehicle is being used.
From what I can find the Delta V to get to main belt asteroids is around 4.5 km/s burn at LEO and another 4.5 km/s to direct land because their is no significant gravity well or anything to brake against.
Is he talking about the asteroid belt or something, I can't see the landing vehicle with it's normal atmospheric EDL profile being appropriate on any other planetary body in the solar system other then Venus.Maybe for Asteroid missions long duration SEP burns as orbital matching delta-V.
While it could retro-propulsive land on an airless body it would be complete overkill to do that on an asteroid with a vehicle with so much thrust and when something like RCS would be able to land and take off. If he thinks the system will be this flexible it strongly suggests an in space transit vehicle is being used.
From what I can find the Delta V to get to main belt asteroids is around 4.5 km/s burn at LEO and another 4.5 km/s to direct land because their is no significant gravity well or anything to brake against.
Ohh, don't forget the Moon. SpaceX might not have any cis-Lunar plans. Doesn't mean they will not provide transportation services to that region if someone is willing to pay.
I think your delta V budget is way too high.
Just read a paper on NTP for a Jupiter flyby that used a Zubrin # of 4.19 Km/sec (edited) to get to Jupiter in 4.1 years. My MCT model has ~6.5 Km/sec delta V capability fully fueled so I think most asteroids close to the ecliptic, and Saturn's moon Titan with aerobraking are in the cards. Drop the 100 tonnes cargo to 50 tonnes or whatever and pick up almost another Km/sec. Then again add in an onboard nuclear electric power source that weighs, what 25 tonnes net after dropping the solar panel array. Whatever, back of the spreadsheet calcs say that a vehicle with MCT capability to Mars surface can get most anywhere, especially when you drop the cargo # a bit.
EDIT: I agree with your implication in bold, just not strongly, perhaps
"Interplanetary Transport System" is kind of bland, but descriptive. "Mars Colonial Transport" had a definite Bladerunner vibe to me, while this seems more like a municipal rapid transit system.If "Interplanetary Transport System" is kind of bland, how about Intrastellar Colonial Transport since it would stay within the our stellar system? Plus it wouldn't sound limited to going to planets.
It worsens it, LEO is around 4 km/s faster the LMO and your deeper in the suns gravity well too, all of that makes for a higher Oberth effect and higher velocity at infinity.
Fewer pieces, fewer steps.
It worsens it, LEO is around 4 km/s faster the LMO and your deeper in the suns gravity well too, all of that makes for a higher Oberth effect and higher velocity at infinity.
However, refuelling in Mars orbit changes these assumptions. Launch into LEO (or HEO), refuel in LEO (or HEO), launch to Mars, aerobrake into Mars orbit (not surface), refuel from Mars ISRU prop launched into LMO (or HMO) by tugs...
Creating fuelling stations at each stepping stone breaks the rocket equation and changes our normal assumptions of spaceflight "efficiency".
That sounds very inefficient, your scrubbing speed at mars that you then have to regain via propellant picked up at mars. Having any kind of braking necessary to reach your next propellant fill up, even if it is frictional braking is going to be really bad. But lets stop spit balling and crunch some numbers, lets say the destination is the Jovian system. Based on this table http://www.projectrho.com/public_html/rocket/appmissiontable.php I can deduce the mars to Jupiter deltaV.
Direct from LEO we need 6.3 km/s
If we go to mars first we can depart with just 3.6 km/s to mars and if we capture fictionally all the way to LMO, then from LMO a burn of 4.3 km/s is needed to send you on to Jupiter. High mars orbit is 1.44 km/s above LMO so the Jupiter burn their would be 2.9 km/s from there.
That's still a total of 7.9 km/s but it is admittedly broken up into two legs which are considerably less then the single burn from LEO. To convert that into propellant fraction at 380 ISP, at Earth you need 4.4:1 propellant to dry ratio to go to Jupiter, but to go to mars you need 1.6:1 and then at mars you need 2.1:1 to complete the journey.
So total propellant is very similar with the direct from Earth method need 19 percent more total. The question is really one of the trades between availability of propellant at Earth and Mars, as I think propellant in mars orbit is going to be significantly more expensive then propellant in earth orbit so I think the direct approach wins.
If you think 6.3 km/s is too much for one vehicle to handle then simply depart from a high Earth orbit which will split the deltaV very nicely into 3.2 and 3.1 which gets you virtually the same departure burn that you would have needed from high mars orbit, which proves my point their is no advantage to falling into the mars gravity well if your destination is an outer planet.
But the whole idea is really moot anyway cause it would take lots more delta v then is viable upon arrival to just land on a moon like Callisto, Titan would be do able with aero-braking and direct decent but we are not going to send people out that far in any kind of conceivable time frame.
Is he talking about the asteroid belt or something, I can't see the landing vehicle with it's normal atmospheric EDL profile being appropriate on any other planetary body in the solar system other then Venus.
Land on Ceres or the Moon. Musk had earlier made reference to landing MCT on the Moon. Ceres would be similar.
The landing thrust is not necessarily a big problem. You can do a burn above the surface and cut off thrust at just the right moment and fall the rest of the way, perhaps using RCS thrusters to finetune the landing.
The spacecraft formerly known as MCT could be purchased by some future billionaire or government space agency to land a large cargo payload on Titan's surface or just aerobrake at Titan and land on say, Enceladus.
Land on Ceres or the Moon. Musk had earlier made reference to landing MCT on the Moon. Ceres would be similar.
The landing thrust is not necessarily a big problem. You can do a burn above the surface and cut off thrust at just the right moment and fall the rest of the way, perhaps using RCS thrusters to finetune the landing.
Ceres? With chemical propulsion?
The moon is certainly possible, though lunar ISRU (water ice) would clearly favor hydrolox (from what I remember).The spacecraft formerly known as MCT could be purchased by some future billionaire or government space agency to land a large cargo payload on Titan's surface or just aerobrake at Titan and land on say, Enceladus.
Sending cargo one-way doesn't require a vehicle like MCT. There's no need to land the entire Earth departure stage and the heat shield if they're not needed to get back. That said, an aeroshell of similar size is still required. For Venus colonization for example, I would expect payloads to go one way with expendable aeroshells.
The spacecraft formerly known as MCT could be purchased by some future billionaire or government space agency to land a large cargo payload on Titan's surface or just aerobrake at Titan and land on say, Enceladus. Various combinations thereof, possibly dropping off small landing vehicle payloads of several tons. Expensive but clearly matching the name of ITS. Would need a nuclear power source that far out but ITS should have the cargo hold space to fit one. Cut the nominal 100T cargo payload to some lower # to pick up delta V.
Is he talking about the asteroid belt or something, I can't see the landing vehicle with it's normal atmospheric EDL profile being appropriate on any other planetary body in the solar system other then Venus.
Venus it is then. It's the easiest place to colonize anyway.
The problem with Ceres is the extra delta V for out of plane xfer. However, slowing down @Ceres does not need delta V again for plane change. I have not run the #s myself to see what a cargo tonnes reduction offers in increased delta V in a Ceres scenario.
I am one that believes is unfairly overlooked to do to 'surfacism' and I find concepts like Landis2land and HAVOC plausible, deploying large balloons to float in the Venusian atmosphere at an altitude with a hospitable temperature and sunlight.
Wouldn't it make more sense to refuel in lunar orbit
How about this?
Fully fuel MCT and a tanker in LEO. Launch them in tandem to a highly elliptic orbit, chosing the orbit so that the tanker has enough fuel left to fully fuel the MCT again. The tanker returns to earth on that orbit. The MCT does its earth departure burn at perigee with max. efficiency. How much delta-v would that gain over starting in LEO? Close to 3km/s?
It is operationally less complex than a depot. A depot may become more efficient when many flights go beyond Mars.
Though, the question is that could a single tanker refuel the MCT/ITS after it has already spent 3km/s impulse for acccelerating itself(and the fuel), or would multiple tankers be needed/would the tankers need to be much bigger than the MCT/ITS?
When the tanker is returning to LEO it does not need much fuel because it can do aerobraking. Or it could return directly to earth instead of returning to LEO.
Without calculation it seems to me that one tanker could be enough or nearly enough to give the max advantage.
A simulation and analysis of this concept has been in L2 for some time.
It's OK to post an L2 link here as non-members simply can't access it
Would a BFR first stage be able to single-stage-to-orbit without the BFS?Without any payload? Quite possibly.
I'm wondering how hard it would be to get a BFR into orbit, fuel it, send it to Titan, propulsively land, and use ISRU to use it repeatedly as a Titan launch vehical.
That should be thousands of tons into Titan orbit, not just hundreds. Could also land thousands of tons.Would a BFR first stage be able to single-stage-to-orbit without the BFS?Without any payload? Quite possibly.
I'm wondering how hard it would be to get a BFR into orbit, fuel it, send it to Titan, propulsively land, and use ISRU to use it repeatedly as a Titan launch vehical.
Your idea is impractical (BFR is not designed itself for interplanetary travel) but still very interesting. Should be capable of putting hundreds of tons into Titan orbit IF (big if!!!) fueled up on the surface.
BFR as a system comprises the booster, the pad and the "refurb street". Having only the booster landed on Titan gives you nothing, usability wise.Well, sure. Kind of is assumed you're landing a bunch of ISRU equipment first.
Heck, you could make a space elevator really easily on Ceres...
Reaching Titan is actually easier then that, it would be 7.3 km/s burn at LEO but you can use friction to do all the deceleration at Titan. Taking off again on Titan would be a nightmare because of the atmospheric thickness, per unit of surface area a column of Titans atmosphere is 7.3 times more massive then Earths and has huge scale height so the launch vehicle must spend a lot of time at low speed plowing through this atmosphere which adds to gravity losses too, finally the density of the atmosphere means rocket engines would yield less then sea-level ISP due to under-expansion in nozzles.That's actually why I proposed the reusable Titan BFR-varient idea- because even if it requires some performace-penalizing modifications to make the trip, the core problems of lift through a thick atmosphere are solved for Earth, and many of the same solutions would be equally viable for Titan.
So don't expect to get a vehicle back from Titan, the transit times alone means their is little point anyway as the vehicle can't get amortized over enough missions to make reuse attractive.
Point of note, SpaceX has repeatedly stated that what they have in mind for the MCT was a Mars Colonial Transport SYSTEM not simply a stage and lander.
Heck, you could make a space elevator really easily on Ceres...I make the Clarke orbit for Ceres to be ~1,200 km from the centre, or about 740 km above the surface. So, a lot easier to build a space elevator for Ceres than Earth!
Taking off again on Titan would be a nightmare because of the atmospheric thickness
However, Musk has explicitly described the system as BFR + BFS. With BFR being a giant booster, and BFS being the upper-stage cum Mars transport cum Mars lander cum Earth return vehicle.
He may have changed his mind, and we'll hopefully find out soon, but there's nothing announced yet by SpaceX that implies he has.
He has said "land the whole thing", which is not just "interpretation."
However, Musk has explicitly described the system as BFR + BFS. With BFR being a giant booster, and BFS being the upper-stage cum Mars transport cum Mars lander cum Earth return vehicle.
He may have changed his mind, and we'll hopefully find out soon, but there's nothing announced yet by SpaceX that implies he has.
Musk has never said BFS is an upper stage or that it is a lander or that it is even a monolithic vehicle rather then a vehicle stack. That's what many people have chosen to interpret it as.
...A figure shown by a user on twitter showing essentially the BFR-boosted BFS flying to Mars, refueling there and flying back to Earth entry was commented by him to be the best representation of the MCT to date...
...A figure shown by a user on twitter showing essentially the BFR-boosted BFS flying to Mars, refueling there and flying back to Earth entry was commented by him to be the best representation of the MCT to date...
Anyone have that image or a link to it? That's something I would like to see!
Hi Everyone, this is a concept of a monolithic BFS: http://imgur.com/gallery/fGzkH
And a previous modular concept: http://imgur.com/a/15fO2
...A figure shown by a user on twitter showing essentially the BFR-boosted BFS flying to Mars, refueling there and flying back to Earth entry was commented by him to be the best representation of the MCT to date...
Anyone have that image or a link to it? That's something I would like to see!
Was this the video?
https://m.youtube.com/playlist?list=PLBU9UJfqaRooKnHY8QtQ399qqRwBqU6W3
...A figure shown by a user on twitter showing essentially the BFR-boosted BFS flying to Mars, refueling there and flying back to Earth entry was commented by him to be the best representation of the MCT to date...
Anyone have that image or a link to it? That's something I would like to see!
Hi Everyone, this is a concept of a monolithic BFS: http://imgur.com/gallery/fGzkH
And a previous modular concept: http://imgur.com/a/15fO2
What are the legs made of?
Not specified, they look thin but ithis concept is more about the layout of the components and the general idea.Landing legs for such a rig would be a major component and the general idea (of anything spaceflight, let alone an ITS) would be to comply with laws of physics and waltz your design around the constraints they impose.
Not specified, they look thin but ithis concept is more about the layout of the components and the general idea.Landing legs for such a rig would be a major component and the general idea (of anything spaceflight, let alone an ITS) would be to comply with laws of physics and waltz your design around the constraints they impose.
Do you know why people find rockets beautiful?
Uh... it's better, but only marginally... I suggest you try to thoughtfully brake some things in your spare time. All kinds of things, of different materials, sizes and shapes. Also, try to be creative with how you apply the force needed to brake a thing - fast, slow, point, area, etc. While you are at it, remember that force has a vector and try to visualize it.Not specified, they look thin but ithis concept is more about the layout of the components and the general idea.Landing legs for such a rig would be a major component and the general idea (of anything spaceflight, let alone an ITS) would be to comply with laws of physics and waltz your design around the constraints they impose.
Do you know why people find rockets beautiful?
I've done this prevously, I thought a little about the system, but I simply didn't mind to put on the final renderings.
...A figure shown by a user on twitter showing essentially the BFR-boosted BFS flying to Mars, refueling there and flying back to Earth entry was commented by him to be the best representation of the MCT to date...
Anyone have that image or a link to it? That's something I would like to see!
I remember seeing this (https://twitter.com/John_Gardi/status/356823217299664896), was this it? It looks like Elon's comment about it being the closest guess he's seen so far is above the linked tweet and is actually about a diagram of a Hyperloop track. I don't see a reply from Elon about the MCT diagram.
Hi Everyone, this is a concept of a monolithic BFS: http://imgur.com/gallery/fGzkH
Uh... it's better, but only marginally... I suggest you try to thoughtfully brake some things in your spare time. All kinds of things, of different materials, sizes and shapes. Also, try to be creative with how you apply the force needed to brake a thing - fast, slow, point, area, etc. While you are at it, remember that force has a vector and try to visualize it.I've done this prevously, I thought a little about the system, but I simply didn't mind to put on the final renderings.Not specified, they look thin but ithis concept is more about the layout of the components and the general idea.Landing legs for such a rig would be a major component and the general idea (of anything spaceflight, let alone an ITS) would be to comply with laws of physics and waltz your design around the constraints they impose.
Do you know why people find rockets beautiful?
While a horizontal lander would be better for cargo access, the MCT/ITS would need to launch from the surface of Mars, so the main thrust would need to be along the longitudinal axis for better aerodynamics during ascent, and there would be extra weight and complexity to support the loads of the vehicle landing on its side in addition to the loads along the vehicle's longitudinal axis.Hi Everyone, this is a concept of a monolithic BFS: http://imgur.com/gallery/fGzkH
Nice one, wouldn't be surprised if the final thing looked a lot like this, but landing this thing vertically doesn't look practical.
It's just a hunch, but I think it's gonna land horizontally, using a separate set of smaller engines. I don't think the MCT/ITS/whatever will have more than 2-3 Raptors and they will be used for orbital stuff, not landing. Oh, and deployable skids on the sides of the thermal shielding would be more robust than current F9 legs for the same mass.
Well, only 4 days left until, hopefully, we know more.
Hi Everyone, this is a concept of a monolithic BFS: http://imgur.com/gallery/fGzkH
Nice one, wouldn't be surprised if the final thing looked a lot like this, but landing this thing vertically doesn't look practical.
It's just a hunch, but I think it's gonna land horizontally, using a separate set of smaller engines. I don't think the MCT/ITS/whatever will have more than 2-3 Raptors and they will be used for orbital stuff, not landing. Oh, and deployable skids on the sides of the thermal shielding would be more robust than current F9 legs for the same mass.
Well, only 4 days left until, hopefully, we know more.
While a horizontal lander would be better for cargo access, the MCT/ITS would need to launch from the surface of Mars, and there would be extra weight and complexity to support the loads of the vehicle landing on its side in addition to the loads generated by the main engine's thrust along the vehicle's longitudinal axis.
Its really puzzling. If you ask me I would think about a 20 years test programme with SEP, expandable disposable habitats, etc. But the rumour mill quasi-consensus point other way. On the Mars landing field, I agree, but other engines would add mass just for landing, and there is the coming back problem, the need to reorient the ship for take-off.
Its really puzzling. If you ask me I would think about a 20 years test programme with SEP, expandable disposable habitats, etc. But the rumour mill quasi-consensus point other way. On the Mars landing field, I agree, but other engines would add mass just for landing, and there is the coming back problem, the need to reorient the ship for take-off.
Think of smallish engines in nacelles placed high on both sides of the ship. You could use the side engines for take off and fire the big ones (Raptors) high in the atmosphere. The mass argument is reasonable, but this alternative provides for a more flexible design. Also, smaller engines far from the ground would prevent damage from dust or rocks blasted off. Several small motors provide redundancy, so a single damaged engine would not risk the whole mission. Add a detachable nose and a ramp like some cargo planes or ferries and you get an easy to operate roll-on roll-of ship.
Just a thought anyway.
...A figure shown by a user on twitter showing essentially the BFR-boosted BFS flying to Mars, refueling there and flying back to Earth entry was commented by him to be the best representation of the MCT to date...
Anyone have that image or a link to it? That's something I would like to see!
I remember seeing this (https://twitter.com/John_Gardi/status/356823217299664896), was this it? It looks like Elon's comment about it being the closest guess he's seen so far is above the linked tweet and is actually about a diagram of a Hyperloop track. I don't see a reply from Elon about the MCT diagram.
That is definetly the image I was thinking of. but I can't find Elon's tweet/comment on it. Perhaps it was deleted. The best I can come up with is a blog entry repeating what the tweet said (including the picture):
https://rocketry.wordpress.com/2013/07/15/spacex-f9-next-generation-booster-gets-full-duration-burn/
"Pretty close to what I have in mind". So its not the Hyperloop tweet, although the content is similar.
The overall idea behind that makes sense if the goal is simple reusability operations. However, "no refueling" in Earth orbit is not an option unless you want the second stage to require SSTO-like delta-v performance or have extremely low mass margins. There will probably be MCTs/ITSs acting as tankers in the plan....A figure shown by a user on twitter showing essentially the BFR-boosted BFS flying to Mars, refueling there and flying back to Earth entry was commented by him to be the best representation of the MCT to date...
Anyone have that image or a link to it? That's something I would like to see!
I remember seeing this (https://twitter.com/John_Gardi/status/356823217299664896), was this it? It looks like Elon's comment about it being the closest guess he's seen so far is above the linked tweet and is actually about a diagram of a Hyperloop track. I don't see a reply from Elon about the MCT diagram.
That is definetly the image I was thinking of. but I can't find Elon's tweet/comment on it. Perhaps it was deleted. The best I can come up with is a blog entry repeating what the tweet said (including the picture):
https://rocketry.wordpress.com/2013/07/15/spacex-f9-next-generation-booster-gets-full-duration-burn/
"Pretty close to what I have in mind". So its not the Hyperloop tweet, although the content is similar.
Maybe it actually was deleted, I just found this reddit comment (https://www.reddit.com/r/spacex/comments/2mi22w/im_writing_a_website_about_spacex_and_i_want/cm6jytv) that has a screenshot (https://pbs.twimg.com/media/BahB8qnCAAAaer5.jpg:large) of John Gardi's MCT diagram and Elon's reply to it saying it's "pretty close to what I have in mind."
Also, this Google+ post (https://plus.google.com/117039636053462680924/posts/asyj9Udoeyz) has a direct link (https://twitter.com/elonmusk/status/356826725805588482) to Elon's tweet, but the tweet doesn't exist anymore.
The overall idea behind that makes sense if the goal is simple reusability operations. However, "no refueling" in Earth orbit is not an option unless you want the second stage to require SSTO-like delta-v performance or have extremely low mass margins. There will probably be MCTs/ITSs acting as tankers in the plan....A figure shown by a user on twitter showing essentially the BFR-boosted BFS flying to Mars, refueling there and flying back to Earth entry was commented by him to be the best representation of the MCT to date...
Anyone have that image or a link to it? That's something I would like to see!
I remember seeing this (https://twitter.com/John_Gardi/status/356823217299664896), was this it? It looks like Elon's comment about it being the closest guess he's seen so far is above the linked tweet and is actually about a diagram of a Hyperloop track. I don't see a reply from Elon about the MCT diagram.
That is definetly the image I was thinking of. but I can't find Elon's tweet/comment on it. Perhaps it was deleted. The best I can come up with is a blog entry repeating what the tweet said (including the picture):
https://rocketry.wordpress.com/2013/07/15/spacex-f9-next-generation-booster-gets-full-duration-burn/
"Pretty close to what I have in mind". So its not the Hyperloop tweet, although the content is similar.
Maybe it actually was deleted, I just found this reddit comment (https://www.reddit.com/r/spacex/comments/2mi22w/im_writing_a_website_about_spacex_and_i_want/cm6jytv) that has a screenshot (https://pbs.twimg.com/media/BahB8qnCAAAaer5.jpg:large) of John Gardi's MCT diagram and Elon's reply to it saying it's "pretty close to what I have in mind."
Also, this Google+ post (https://plus.google.com/117039636053462680924/posts/asyj9Udoeyz) has a direct link (https://twitter.com/elonmusk/status/356826725805588482) to Elon's tweet, but the tweet doesn't exist anymore.
Well, the "no refueling" part is actually the only point that is contradicted by something Elon said himself (that there will be re-fueling before departure). So its a safe bet that this point is (among others, perhaps) why he said "pretty close" and not "spot on".Oh yeah, he did say that. (http://shitelonsays.com/transcript/elon-musk-at-mits-aeroastro-centennial-part-2-of-6-2014-10-24) (9:40 in video)
I mean, if you do a densified liquid methalox rocket with on-orbit refueling, so like you load the spacecraft into orbit and then you send a whole bunch of refueling missions to fill up the tanks and you have the Mars colonial fleet - essentially - that gets built up during the time between Earth-Mars synchronizations, which occur every 26 months, then the fleet all departs at the optimal transfer point.
Of course is science fiction. I took more time on these first concepts, before the "put the entire thing on the surface of mars" rumour, I guess.
1) a
2) 9-12 engines
3) 8.5-10 m diameter
4)
5)c) BFS contains smaller 'ejection pod' where humans reside during launch
6)
7) a)direct entry
8.) c) All chemical for now, but plans to incorporate SEP down the road
I hope it will be called Bender 8)
Well, this didn't get the traction that I'd hoped, but I'm still wondering if anyone would like to 'put their money where their mouth is' and declare their positions on the (hopefully) upcoming reveal details. Only a few days left! I'll start with my guesses:
1-a, 2-31, 3-15m, 4-17.5Mlbs, 5a, 6a, 7a, 8cQuotea) MCT is composed simply of a BFR 1st stage and BFS 2nd stage/spacecraft (only)
1) Overall Launch Architecture
a) MCT is composed simply of a BFR 1st stage and BFS 2nd stage/spacecraft (only)
b) Boost phase consists of 2 stages, which put the BFS into orbit
c) Other: 3rd stage, 'half' stages, drop tanks, etc.Quote2) Number of Raptor Engines on BFR (1st stage)
27 engines.Quote3) Diameter of BFR (1st stage)
15-20 metersQuote4) Total Raptor 1st stage thrust (sl)Don't know.Quote5) LAS Architectured) Other, non-traditional LAS design
a) No LAS - BFS is the escape mechanism
b) Traditional LAS - above BFS and is nominally jettisoned during launch phase
c) BFS contains smaller 'ejection pod' where humans reside during launch
d) Other, non-traditional LAS designQuote6)
Shape and Landing Mode of BFS
a) Capsule (perhaps elongated), w/ TPS on base
b) Cylindrical or biconic - horizontal landing
c) Cylindrical or biconic - vertical landing
d) Other
Capsule, will probably be a human rated Red Dragon.Quote7) Mars and Earth return
a) BFS does direct entry into Mars and Earth atmosphere
b) BFS does orbital capture before performing entry burn and landing
c) Same as b, but upon Earth return, stays in orbit for next synod
c) Same as b, but upon Earth return, stays in orbit for next synodQuote8) Use of non-chemical thrust
a) Not part of the plan
b) Will use SEP for some/all of the big transits
c) All chemical for now, but plans to incorporate SEP down the road
C, all chemical for now, but will probably look into electrical engines once the technology has developed.
Lot of talk here seem to be about a single vehicle that will launch from earth, land on mars, and then return to earth.
Musk isn't one to waste things, surely he will be using Red Dragon vehicle to handle the Mars orbit to ground bit and the IST will just handle the Earth to Mars bit?
It seems a bit risky to me to design a vehicle capable of being launch from earth with dozens of passengers and all their supplies sent to land to Mars and expect the same vehicle to do the reverse again.
Dragons might make excellent Star Wars style escape pods for the MCT, lifeboats able to save the human cargo...
An affordable, fully reusable HLV changes the rules of the game.To think that several years ago, this would have been considered an oxymoron. That reusability would have to be designed for existing markets, that they would have small payload capacities because they would need very high flight rates to be economically viable.
If there was a demand of 100-200mT/yr propellant in LEO, then a very small RLV could be profitable. No sufficiently large markets seem likely for Falcon 9 sized payloads in the near future.
SpaceX may decide to bite the bullet and dev. an F-1 class version of Raptor for BFR to avoid an N-1 type design and to carry over the octaweb design heritage of F9. Raptor is said to be highly scalable so there is no reason whatsoever that several sizes of it for different tasks can be dev. maybe except for dev. cost.Not to disagree with your point, but to nitpick on the phrasing:
EM did say the plan do use only one size of Raptor throughout the BFR/MCT architecture may change during his AMA in January 2015.
I am sure EM knows what happening to the N-1 and I can't imagine him being that stupid.
We will find out on Tuesday.
SpaceX may decide to bite the bullet and dev. an F-1 class version of Raptor for BFR to avoid an N-1 type design and to carry over the octaweb design heritage of F9. Raptor is said to be highly scalable so there is no reason whatsoever that several sizes of it for different tasks can be dev. maybe except for dev. cost.
EM did say the plan do use only one size of Raptor throughout the BFR/MCT architecture may change during his AMA in January 2015.
I am sure EM knows what happening to the N-1 and I can't imagine him being that stupid.
We will find out on Tuesday.
SpaceX may decide to bite the bullet and dev. an F-1 class version of Raptor for BFR to avoid an N-1 type design and to carry over the octaweb design heritage of F9. Raptor is said to be highly scalable so there is no reason whatsoever that several sizes of it for different tasks can be dev. maybe except for dev. cost.
EM did say the plan do use only one size of Raptor throughout the BFR/MCT architecture may change during his AMA in January 2015.
I am sure EM knows what happening to the N-1 and I can't imagine him being that stupid.
We will find out on Tuesday.
http://m.imgur.com/a/87OOT
- Was this proposal, by 'Coborop' on Reddit, ever discussed in this thread?
It only takes one engine to explode out of 30 to bring down an LV. I believe this happened to at least one of the N-1's. Anyone including SpaceX even attempting something like the N-1 beggars belief. The more engines you have the greater the risk of one of them exploding causing catastrophic LV failure. 5-9 engines is the optimum no. on a 1st stage to strike a balance between the risk of engine explosion and good engine out capability for benign engine failures. Lets hope that SpaceX have learned the N-1 lesson and the BFR design we should see on Tuesday will have much less than 30 engines on it.SpaceX may decide to bite the bullet and dev. an F-1 class version of Raptor for BFR to avoid an N-1 type design and to carry over the octaweb design heritage of F9. Raptor is said to be highly scalable so there is no reason whatsoever that several sizes of it for different tasks can be dev. maybe except for dev. cost.
EM did say the plan do use only one size of Raptor throughout the BFR/MCT architecture may change during his AMA in January 2015.
I am sure EM knows what happening to the N-1 and I can't imagine him being that stupid.
We will find out on Tuesday.
N-1 failed due to fuel system problems that would not exist under present engineering conventions.
Number of engines does not increase risk. Indeed, there's an argument that a high number of engines reduces risk.
It's usually safe to assume that SpaceX isn't "stupid", they're better informed and acting under better data.
1) Overall Launch Architecture
a) MCT is composed simply of a BFR 1st stage and BFS 2nd stage/spacecraft
b) Boost phase consists of 2 stages, which put the BFS into orbit
c) Other: 3rd stage, 'half' stages, drop tanks, etc.
Going with (a)
2) Number of Raptor Engines on BFR (1st stage)
< 30, my best estimate is 25-27 if thrust stays close to 230 tonnes range
3) Diameter of BFR (1st stage)
Range 12.5m-15m, best estimate 15m 1st stage
4) Total Raptor 1st stage thrust (sl)
60 Meganewtons and T/W > 1.3
5) LAS Architecture
a) No LAS - BFS is the escape mechanism
b) Traditional LAS - above BFS and is nominally jettisoned during launch phase
c) BFS contains smaller 'ejection pod' where humans reside during launch
d) Other, non-traditional LAS design
Best guess is (a)
6) Shape and Landing Mode of BFS
a) Capsule (perhaps elongated), w/ TPS on base
b) Cylindrical or biconic - horizontal landing
c) Cylindrical or biconic - vertical landing
d) Other
Going with (c), definitely no horizontal landing
7) Mars and Earth return
a) BFS does direct entry into Mars and Earth atmosphere
b) BFS does orbital capture before performing entry burn and landing
c) Same as b, but upon Earth return, stays in orbit for next synod
(a)
8) Use of non-chemical thrust
a) Not part of the plan
b) Will use SEP for some/all of the big transits
c) All chemical for now, but plans to incorporate SEP down the road
(c) strongly favor
Can anyone think of more/better questions?
And announces a >F-1 thrust version of Raptor to keep no. of engines on BFR to no more than 9.
1) Overall Launch Architecture
a) MCT is composed simply of a BFR 1st stage and BFS 2nd stage/spacecraft
b) Boost phase consists of 2 stages, which put the BFS into orbit
c) Other: 3rd stage, 'half' stages, drop tanks, etc.
Going with (a)
2) Number of Raptor Engines on BFR (1st stage)
< 30, my best estimate is 25-27
3) Diameter of BFR (1st stage)
Range 12.5m-15m, best estimate 15m 1st stage
4) Total Raptor 1st stage thrust (sl)
60 Meganewtons and T/W > 1.3
5) LAS Architecture
a) No LAS - BFS is the escape mechanism
b) Traditional LAS - above BFS and is nominally jettisoned during launch phase
c) BFS contains smaller 'ejection pod' where humans reside during launch
d) Other, non-traditional LAS design
Best guess is (a)
6) Shape and Landing Mode of BFS
a) Capsule (perhaps elongated), w/ TPS on base
b) Cylindrical or biconic - horizontal landing
c) Cylindrical or biconic - vertical landing
d) Other
Going with (c), definitely no horizontal landing
7) Mars and Earth return
a) BFS does direct entry into Mars and Earth atmosphere
b) BFS does orbital capture before performing entry burn and landing
c) Same as b, but upon Earth return, stays in orbit for next synod
(a)
8) Use of non-chemical thrust
a) Not part of the plan
b) Will use SEP for some/all of the big transits
c) All chemical for now, but plans to incorporate SEP down the road
(c) strongly favor
Can anyone think of more/better questions?
Predict Musk will miss 1st crewed landing by >= 3 synods
5-6 Rvacs on BFS stage 2
Raptor sea level will have 10s of tonnes thrust more than the 230 tonnes mentioned by Elon
Entire BFR/BFS GLOW masses under 5.000 tonnes; my estimate ~4,500
Height of BFR/BFS stack under 120m; my estimate <100m
Cargo version, tanker version, crewed version of BFS
1st crewed landing on Mars 8-12 humans planned
Just over 48 hours until Musk makes fools of us
http://m.imgur.com/a/87OOT
- Was this proposal, by 'Coborop' on Reddit, ever discussed in this thread?
It only takes one engine to explode out of 30 to bring down an LV. I believe this happened to at least one of the N-1's. Anyone including SpaceX even attempting something like the N-1 beggars belief. The more engines you have the greater the risk of one of them exploding causing catastrophic LV failure. 5-9 engines is the optimum no. on a 1st stage to strike a balance between the risk of engine explosion and good engine out capability for benign engine failures. Lets hope that SpaceX have learned the N-1 lesson and the BFR design we should see on Tuesday will have much less than 30 engines on it.SpaceX has been there and done that.
Up-feed plumbing of propellant to Roc makes this way too complex & heavy. Cosine losses and reduced engine performance by running stage 2 Roc at sea level for Earth launch a bad tradeoff. Just add engines on 1st stage Sling until it's big enough to launch upper stage.
It only takes one engine to explode out of 30 to bring down an LV. I believe this happened to at least one of the N-1's. Anyone including SpaceX even attempting something like the N-1 beggars belief.
>
And announces a >F-1 thrust version of Raptor to keep no. of engines on BFR to no more than 9.
...
Also for purposes of reliability and repair.That's what I meant, wouldn't it take longer to repair/inspect 30 engines on one stage compared to 9 or 20? Extra person-hours required for turnaround, etc.
Production Raptor goal is specific impulse of 382 seconds and thrust of 3 MN (~310 metric tons) at 300 bar
Chamber pressure is almost 3X Merlin, so engine is about the same size for a given area ratio
Looks like crazy EM is going for an N-1 type design for BFR after all judging from the announced Raptor thrust. Lets hope the 1st 4 BFR's don't explode on or just after launch.
And announces a >F-1 thrust version of Raptor to keep no. of engines on BFR to no more than 9.
Why do you assume early Sixties engine logic still applies? The base principles of rocketry do indeed move that slowly. The engineering, materials science, and practically everything else, do not.
Target BFR thrust stated by EM and Raptor announced thrust seems to indicate 25-27 Raptors which is still too many for my liking.
My best guess
1 - a (BFS 2nd stage is MCT)
2 - 46x raptor on 1st stage, multiple used for landing, for redundancy
3 - 15m
4 - 120 MN SL thrust
5 - Redundant raptors on MCT for LAS / no LAS needed
6 - c
7 - a (or combination with b where there is more than one pass through atmosphere before landing)
8 - a / c
1) Overall Launch Architecture
a) MCT is composed simply of a BFR 1st stage and BFS 2nd stage/spacecraft
b) Boost phase consists of 2 stages, which put the BFS into orbit
c) Other: 3rd stage, 'half' stages, drop tanks, etc.
Going with (a)
2) Number of Raptor Engines on BFR (1st stage)
< 30, my best estimate is 25-27 if thrust stays close to 230 tonnes range
3) Diameter of BFR (1st stage)
Range 12.5m-15m, best estimate 15m 1st stage
4) Total Raptor 1st stage thrust (sl)
60 Meganewtons and T/W > 1.3
5) LAS Architecture
a) No LAS - BFS is the escape mechanism
b) Traditional LAS - above BFS and is nominally jettisoned during launch phase
c) BFS contains smaller 'ejection pod' where humans reside during launch
d) Other, non-traditional LAS design
Best guess is (a)
6) Shape and Landing Mode of BFS
a) Capsule (perhaps elongated), w/ TPS on base
b) Cylindrical or biconic - horizontal landing
c) Cylindrical or biconic - vertical landing
d) Other
Going with (c), definitely no horizontal landing
7) Mars and Earth return
a) BFS does direct entry into Mars and Earth atmosphere
b) BFS does orbital capture before performing entry burn and landing
c) Same as b, but upon Earth return, stays in orbit for next synod
(a)
8) Use of non-chemical thrust
a) Not part of the plan
b) Will use SEP for some/all of the big transits
c) All chemical for now, but plans to incorporate SEP down the road
(c) strongly favor
Can anyone think of more/better questions?
Predict Musk will miss 1st crewed landing by >= 3 synods
5-6 Rvacs on BFS stage 2
Raptor sea level will have 10s of tonnes thrust more than the 230 tonnes mentioned by Elon
Entire BFR/BFS GLOW masses under 5.000 tonnes; my estimate ~4,500
Height of BFR/BFS stack under 120m; my estimate <100m
Cargo version, tanker version, crewed version of BFS
1st crewed landing on Mars 8-12 humans planned
Just over 48 hours until Musk makes fools of us
I'm guessing now that this won't be a very huge problem (or will it?), because when Falcon 9's first stage has 9 engines, people aren't talking about how this is harder to maintain than one big engine + at least 2 small engines on the sides for landing; people instead talk about how it is more expensive to maintain production of multiple engine types instead.Also for purposes of reliability and repair.That's what I meant, wouldn't it take longer to repair/inspect 30 engines on one stage compared to 9 or 20? Extra person-hours required for turnaround, etc.
I'm guessing now that this won't be a very huge problem (or will it?), because when Falcon 9's first stage has 9 engines, people aren't talking about how this is harder to maintain than one big engine + at least 2 small engines on the sides for landing; people instead talk about how it is more expensive to maintain production of multiple engine types instead.Also for purposes of reliability and repair.That's what I meant, wouldn't it take longer to repair/inspect 30 engines on one stage compared to 9 or 20? Extra person-hours required for turnaround, etc.
I agree with you, but N1 isn't something that "speaks in favour" of having many smaller engines instead of few big engines, though.
I agree with you, but N1 isn't something that "speaks in favour" of having many smaller engines instead of few big engines, though.
I think Musk's talk tomorrow will follow more or less the following logic:
...
7) Makes case for public-private partnerships. "Cannot do this without NASA. But NASA might not be able to do this without us." Presents clever funding scheme which will keep them running with a small fraction of what NASA invests in SLS every year. "NASA should diversify", not only SLS but also alternatives in case rocket grounded after mishap. Multiple plans, "also Blue Origin", should keep us on track to Mars.
O.K. I want to change my guess for number of engines from 31 to 27.You can have 27 engines with a single center engine. The configuration for BFR would be a single center engine surrounded by rings of 8 and 18 engines.
And this isn't just because Musk informed us that the thrust is going to be higher than previously hinted at. ;)
It's also because I realized that, while 31 gives a very nice, minimum diameter packing pattern with a single center engine (like the F9), having a single center (landing) engine is probably not appropriate or desired here. It's unlikely that the BFR could land on a single engine - just too heavy. 3 centralized engines would probably be appropriate. And the nice symetrical, minimum pattern for 3 center engines comes out to be 27.
We'll know in just over a day (I hope).
I think Musk's talk tomorrow will follow more or less the following logic:
...
7) Makes case for public-private partnerships. "Cannot do this without NASA. But NASA might not be able to do this without us." Presents clever funding scheme which will keep them running with a small fraction of what NASA invests in SLS every year. "NASA should diversify", not only SLS but also alternatives in case rocket grounded after mishap. Multiple plans, "also Blue Origin", should keep us on track to Mars.
I like your thinking save for #7. Musk will, as usual, praise NASA as an excellent partner. But he will be wise to avoid any mention of SLS or NASA 'needing' SpaceX. I can think of many a Senator and Congressperson who will already feel threatened by the reveal of BFR.
Musk's best bet is to avoid either-or comparisons between The US Govt.'s Rocket and his own. Rather, emphasize that SpaceX's system will be yet another tool available in the public/private toolbox of US Space Exploration over the years to come.
Finally, I would add one to your list:
2.5) Red Dragon in ~2018; emphasize SX-NASA cooperation, demonstration of ISRU on-planet.
I think Musk's talk tomorrow will follow more or less the following logic:
...
7) Makes case for public-private partnerships. "Cannot do this without NASA. But NASA might not be able to do this without us." Presents clever funding scheme which will keep them running with a small fraction of what NASA invests in SLS every year. "NASA should diversify", not only SLS but also alternatives in case rocket grounded after mishap. Multiple plans, "also Blue Origin", should keep us on track to Mars.
I like your thinking save for #7. Musk will, as usual, praise NASA as an excellent partner. But he will be wise to avoid any mention of SLS or NASA 'needing' SpaceX. I can think of many a Senator and Congressperson who will already feel threatened by the reveal of BFR.
Musk's best bet is to avoid either-or comparisons between The US Govt.'s Rocket and his own. Rather, emphasize that SpaceX's system will be yet another tool available in the public/private toolbox of US Space Exploration over the years to come.
Finally, I would add one to your list:
2.5) Red Dragon in ~2018; emphasize SX-NASA cooperation, demonstration of ISRU on-planet.
I'm revising BFR diameter down to a range of 10 - 12 m based on Musk comment on raptor being comparable to Merlin size but carrying 3x thrust.
That completely rules out 15 m diameters, a vehicle that size is approximately 9x the base area of F9, with triple thrust density of Raptor would lead to 27x the thrust of F9 or something close to 45 million pounds of thrust, 3x times what Musk has aimed for.
Fundamentally the 15 m speculation was nothing more then a crude attempt to multiply the thrust of F9 by a factor of 9 while completely ignoring the thrust density improvements that come from the full-flow staged combustion cycle.
...
I'm revising BFR diameter down to a range of 10 - 12 m based on Musk comment on raptor being comparable to Merlin size but carrying 3x thrust.Raptor could be tiny and still obtain 3x as much thrust if you're willing to tolerate low Isp. Nozzles get exponentially larger as you raise Isp closer to its maximum possible figure; That max figure itself scales with chamber pressure (though not only chamber pressure, and not linearly). The base area of F9 is 3.66m, so a 15m BFR would be a factor 16.8x as large.
That completely rules out 15 m diameters, a vehicle that size is approximately 9x the base area of F9, with triple thrust density of Raptor would lead to 27x the thrust of F9 or something close to 45 million pounds of thrust, 3x times what Musk has aimed for.
Fundamentally the 15 m speculation was nothing more then a crude attempt to multiply the thrust of F9 by a factor of 9 while completely ignoring the thrust density improvements that come from the full-flow staged combustion cycle.
I'm revising BFR diameter down to a range of 10 - 12 m based on Musk comment on raptor being comparable to Merlin size but carrying 3x thrust.
That completely rules out 15 m diameters, a vehicle that size is approximately 9x the base area of F9, with triple thrust density of Raptor would lead to 27x the thrust of F9 or something close to 45 million pounds of thrust, 3x times what Musk has aimed for.
Fundamentally the 15 m speculation was nothing more then a crude attempt to multiply the thrust of F9 by a factor of 9 while completely ignoring the thrust density improvements that come from the full-flow staged combustion cycle.
The 15 m diameter speculation had little to do with individual engine thrust.
1. Various attempts have been made to optimise the BFR mass, these give least tank mass for a 13-15 m stage.
2. The BFS has been said by Musk to be very large, a scaled Dragon capsule design might have to be over 23 m, which would fit much more easily on a 15 m stage. Biconic or semi-lifting body BFS designs would seem to need 12-15 m diameter to enclose the inidcated volume.
3. L2 info and various leaks on reddit
but then, if they cannot do it on their own, he has to make the case why it would be a good thing for NASA to invest some money in the SpaceX plan
SpaceX has what NASA needs. SX are being paid 1.6B for it. And then there's CRS-2. That's quite some money invested by NASA into the SpaceX plan
Heck, even if it was just NASA giving SpaceX money, NASA will be paid back in spades just Red Dragon succeeds.SpaceX has what NASA needs. SX are being paid 1.6B for it. And then there's CRS-2. That's quite some money invested by NASA into the SpaceX plan
That's NASA purchasing services at a very good price.
O.K. I want to change my guess for number of engines from 31 to 27.
And this isn't just because Musk informed us that the thrust is going to be higher than previously hinted at. ;)
It's also because I realized that, while 31 gives a very nice, minimum diameter packing pattern with a single center engine (like the F9), having a single center (landing) engine is probably not appropriate or desired here. It's unlikely that the BFR could land on a single engine - just too heavy. 3 centralized engines would probably be appropriate. And the nice symetrical, minimum pattern for 3 center engines comes out to be 27.
O.K. I want to change my guess for number of engines from 31 to 27.
And this isn't just because Musk informed us that the thrust is going to be higher than previously hinted at. ;)
It's also because I realized that, while 31 gives a very nice, minimum diameter packing pattern with a single center engine (like the F9), having a single center (landing) engine is probably not appropriate or desired here. It's unlikely that the BFR could land on a single engine - just too heavy. 3 centralized engines would probably be appropriate. And the nice symetrical, minimum pattern for 3 center engines comes out to be 27.
3 center engines is much worse than 1.
If any of those 3 center engines fail, then it's crash and burn.
But with many more engines, the landing burn can be done with center engine + ANY two engines that are on oppisite sides of the craft.
I'm revising BFR diameter down to a range of 10 - 12 m based on Musk comment on raptor being comparable to Merlin size but carrying 3x thrust.
That completely rules out 15 m diameters, a vehicle that size is approximately 9x the base area of F9, with triple thrust density of Raptor would lead to 27x the thrust of F9 or something close to 45 million pounds of thrust, 3x times what Musk has aimed for.
Fundamentally the 15 m speculation was nothing more then a crude attempt to multiply the thrust of F9 by a factor of 9 while completely ignoring the thrust density improvements that come from the full-flow staged combustion cycle.
Target BFR thrust stated by EM and Raptor announced thrust seems to indicate 25-27 Raptors which is still too many for my liking.
42 engines!!!! good god!
That will be like 25 million pounds of thrust at liftoff!!!! :o 8) ;D
3 center engines is much worse than 1.
If any of those 3 center engines fail, then it's crash and burn.
But with many more engines, the landing burn can be done with center engine + ANY two engines that are on oppisite sides of the craft.
Musk: spaceship can serve as own abort system from booster, but on Mars, either you’re taking off or you’re not. #IAC2016
Asked about abort modes for launcher @elonmusk said "make it very reliable ... you do not have parachutes for commercial airliners" #IAC2016
I guessed 36 engines, did not see anybody guessing more, so if we use The Price Is Right rules....
Matthew
I guessed 36 engines, did not see anybody guessing more, so if we use The Price Is Right rules....
Matthew
My guess was 46
I guessed 36 engines, did not see anybody guessing more, so if we use The Price Is Right rules....
Matthew
My guess was 46
Must have missed that. Are you familiar with The Price Is Right over there in Sweden?
My best guess
4 - 120 MN SL thrust
I'm thrilled by the Musk presentation. However, I'm worried by what I didn't see.
Two MCT concerns in particular I have; heat and power. How, exactly, will they preserve the subcooled LOX and methane needed for landing during interplanetary cruise (80 to 115 days)? And also, how will they handle the heat generated on MCT by the 100 people aboard, plus life support, plus lighting, etc? On ISS, cooling is very complex, and if I'm remembering correctly, the cooling system (ammonia based) outmasses the solar arrays.
Also, regarding solar arrays for power... what we see in the demo looks like an area on par with the ISS arrays, so how does one get sufficient power from that area to handle life support, etc, for 100 people? And that's without considering the diminished sunlight at Mars.
One guess I have regarding mass fractions, etc, is that refueling won't be done in LEO, but in an elliptical orbit - lower delta/v needed for the TMI burn that way, thus reducing the mass fraction issue.
...
One guess I have regarding mass fractions, etc, is that refueling won't be done in LEO, but in an elliptical orbit - lower delta/v needed for the TMI burn that way, thus reducing the mass fraction issue.
The phrase used on the show was "closest without going over" (in other words, the largest guess which is equal to or below the target value).Must have missed that. Are you familiar with The Price Is Right over there in Sweden?I'm sorry, I'm not familiar with that.
Here is my rough MSPaint size comparison between the ICT and some space stations/craft.
Here is my rough MSPaint size comparison between the ICT and some space stations/craft.
Can you do the IST/MCT sitting on the deck of the ASDS? :o
Might be fun for folks to go back and edit their prediction lists to show how they did (me? not that great)
A good look at those enormous rainbirds in that photo. Assuming the yellow railing is 4' high, I get a quick-and dirty estimate of the rainbirds at 57'. :o That seems ridiculously big.
Could they be backed off from the rocket far enough that that actually makes sense? Or am I way off on the height?
Pure speculation - These are oversized for an F9 family vehicle but are right sized for a larger follow on vehicle. 39A is the eventual BFR launch site.
Might be fun for folks to go back and edit their prediction lists to show how they did (me? not that great)
I don't know if I ever made real "prediction checkbox list", but I have made several posts with the following predictions that came true:
- spacecraft as integrated 2nd stage
- biconic/side entry
- powered by multiple raptors instead of a single one (although I was off on the count, I expected 5)
- no launch abort capsule (abort the whole thing)
The booster was roughly what I expected, but bigger and lacking landing gear.
Might be fun for folks to go back and edit their prediction lists to show how they did (me? not that great)
I don't know if I ever made real "prediction checkbox list", but I have made several posts with the following predictions that came true:
- spacecraft as integrated 2nd stage
- biconic/side entry
- powered by multiple raptors instead of a single one (although I was off on the count, I expected 5)
- no launch abort capsule (abort the whole thing)
The booster was roughly what I expected, but bigger and lacking landing gear.
To be precise the BFS we were shown is cylindrical with a rounded nose and thick fins that look to hold the landing gear, not bi-conic which would involve the vehicle widening all the way down to the base, the fins can give a false sense that's whats happening but the cut-away is clear the central body is cylindrical.
That said both cylindrical and bi-conic were combined on the questionnaire and the entry profile is indeed horizontal which was the more important distinction.
Really, you of all people nitpicking me on the 'spaceship' predictions? ;) Cylindrical is indeed what all my previously posted MCT drawings showed... Here are two old sketches of a side re-entering cylindrical MCT, pretty close in the end. (other than size, flipping the propellant and crew/cargo, and engine count)
Really, you of all people nitpicking me on the 'spaceship' predictions? ;) Cylindrical is indeed what all my previously posted MCT drawings showed... Here are two old sketches of a side re-entering cylindrical MCT, pretty close in the end. (other than size, flipping the propellant and crew/cargo, and engine count)
I've seen some other posts claiming the presented vehicle was bi-conic and wanted to clarify, your perdition was indeed for a cylindrical vehicle which is why I though it odd that you wrote bi-conic and that you might have been misinformed by people throwing that word around.
Also I don't see why 'I' of all people should be considered an invalid source of nitpicking even in jest, I did correctly predict nearly everything about the booster in the face of a strong opposing consensus and while I was wrong about the 2nd stage everyone else was basing their prediction on rumors and L2 information that I don't have access too. And I stick by my opinion that the massive BFS isn't workable and think the design as presented will go the way to the reusable F9 2nd stage.
My one BFR speculation was dead on. Helodriver predicts the future (again) ;)A good look at those enormous rainbirds in that photo. Assuming the yellow railing is 4' high, I get a quick-and dirty estimate of the rainbirds at 57'. :o That seems ridiculously big.
Could they be backed off from the rocket far enough that that actually makes sense? Or am I way off on the height?
Pure speculation - These are oversized for an F9 family vehicle but are right sized for a larger follow on vehicle. 39A is the eventual BFR launch site.
Those who insisted (pretty strenuously as I recall) on additional stages, 'pusher stages' for LEO departure, aerocapture (at Mars or Earth), early transition to SEP, and especially Mars orbit refueling were pretty much out to lunch.
From interplanetary space, the ship enters the atmosphere, either capturing into orbit or proceeding directly to landing
Those who insisted (pretty strenuously as I recall) on additional stages, 'pusher stages' for LEO departure, aerocapture (at Mars or Earth), early transition to SEP, and especially Mars orbit refueling were pretty much out to lunch.
The aerocapture is on the table (see slide 38 of the presentation):QuoteFrom interplanetary space, the ship enters the atmosphere, either capturing into orbit or proceeding directly to landing
My one BFR speculation was dead on. Helodriver predicts the future (again) ;)A good look at those enormous rainbirds in that photo. Assuming the yellow railing is 4' high, I get a quick-and dirty estimate of the rainbirds at 57'. :o That seems ridiculously big.
Could they be backed off from the rocket far enough that that actually makes sense? Or am I way off on the height?
Pure speculation - These are oversized for an F9 family vehicle but are right sized for a larger follow on vehicle. 39A is the eventual BFR launch site.
Come on. You couldn't even predict that you yourself will be at the Grand Mars Reveal.
a few more teasers before the september reveal:Bold mine.
https://www.washingtonpost.com/news/the-switch/wp/2016/06/10/elon-musk-provides-new-details-on-his-mind-blowing-mission-to-mars/
1 red dragon in 2018, 'at least 2' in 2020, then first flight of MCT in 2022...
From the article:QuoteThen in 2022, Musk said he hoped to launch what the company now sometimes refers to as the Mars Colonial Transporter, designed to bring a colony to Mars.
I'm sorry, but this isnutssomewhat optimistic. You all realize that 2022 is only six years away, right? Regardless of the fact that Dragon v2 hasn't flown yet
, and regardless of the fact that FH hasn't flown yet; NOTHING concrete about BFR/MCT has even been released, and Musk is talking about launching one in six years. Six. Years.
Six years to get BFR off the ground, literally. To build a factory on the scale of Michoud (only bigger) for fabrication and assembly of BFR and MCT. To build a huge HIF to handle the 12.5m or 15m cores, or heck even to lease one of the VAB high bays and get it fitted out for BFR. To build all of the ground support infrastructure and ground transportation.
To get the entire Raptor engine (not just components) off of the drawing board and into the test stands and validated.
Heck, you guys are still arguing over where the thing will be built and launched from. Do you think that would really be the case if they were going to be rolling off the assembly line in less than six years?
Approaching 30 days from what (we hope) is the big reveal, I thought it a good time to revisit and post revised BFR/MCT speculation before any info leaks out. Trying to stay within the parameters of what Musk has said as I best understand. A TSTO vehicle launched by a re-useable, single core BFR that puts the BFS a.k.a. the MCT into LEO where it is re-fueled, travels to and lands on Mars where it is again refueled for the journey back to Earth carrying a quarter of the outbound “cargo” mass. The outbound cargo masses 100 tonnes which I assume means either cargo or people or a combination thereof. BFS/MCT mass not included in the 100T.
Correct mission profile. Most here with a few notable exceptions agreed with this, so no great insight.
Myriad unknowns led by the dry mass of the BFS. Rocket equation dictates various mass assumptions here can produce wildly different answers.
My predictions, metric unless otherwise stated:
1. Entire launch vehicle BFR+BFS masses under 5,000T. Guestimate ~4,500T.
WAY off!
2. BFS dry mass < 100T, my pick is 85T carbon composites BUT heavier than some predictions because ruggedized to allow for minimal maintenance.
WAY off
3. BFR absolutely > 10m diameter to fit enough engines. Likely between 12.5 and 15m. My guess 15m. Allows addition of more engines in the future.
Another miss as I was confident of 12.5m or greater. Not counting 17m flare outs on ITS craft.
4. My guestimate BFR+BFS stack <100m height. Certainly <125m.
Barely made my "certainly under 125m but missed on 15m. Skinny rocket 12m increased height.
5. Sticking with the “over 230T” Raptor thrust Elon mentioned, I get 25-27 engines. My guestimate is 26 with “over 230T” as 235T in my spreadsheet. Around 13.5 million Lbs force.
Engine # most likely wrong because…
I was right that I was wrong :)
6. Predict that Raptor engine design goal thrust changed to higher than 230T previously stated, but only by several 10s of tonnes, not hundreds.
Hit! I was confident that more detailed design would increase Raptor SL thrust by several 10s of tonnes.
It helped that Bezos BE-4 thrust was higher than Elon's earlier 230T for Raptor. That shall not stand! :)
7. BFS with 5 Rvac engines
Close but no cigar.
8. RTLS minimizes cost, turnaround time, effort. Changed my opinion from max payload ASDS for those reasons. Just make the BFR bigger. Stages low and slow ~2.2 Km/sec. “Easy” recovery & re-flight vs F9 GTO flights.
Hit!
9. Initial BFR test flights likely equipped with less engines and less payload.
Unknown
10. Large crew volume design >2,000m3. Initial flights with less people & people space but more cargo space.
I believe crew volume is ~3,000 m3 so this is a hit. Felt that nuclear sub range 20 something m3 was the goal. Never agreed with those in the 10m3 and under range. You know who you are.
11. Initial crewed Mars mission will carry 6-12 people. 10 is my latest #. Why?
NASA & other nations will buy seats.
http://forum.nasaspaceflight.com/index.php?topic=40683.msg1557261#msg1557261
Unknown
12. SEP still under development awaits later opposition cargo transits
Scoring as a miss in that not mentioned
13. BFS will have “exotic” upper mounted engines for rough terrain Mars landing &takeoff (just echoing others’ analysis here)
Big miss. I am concerned about SX's approach here.
14. BFS will be a lifting body for EDL, but not a scaled up Dragon capsule shape. It will look badass.
HIT!
You know we’re totally screwed trying to predict Musk because he already warned us,
“When it looks more like an alien dreadnought, that’s when you know you’ve won.”
I’ve attached a spreadsheet showing different assumptions, BFS mass, etc.
Anyone else want to update their speculations?
1) Overall Launch Architecture
a) MCT is composed simply of a BFR 1st stage and BFS 2nd stage/spacecraft
b) Boost phase consists of 2 stages, which put the BFS into orbit
c) Other: 3rd stage, 'half' stages, drop tanks, etc.
Going with (a)
Hit
2) Number of Raptor Engines on BFR (1st stage)
< 30, my best estimate is 25-27 if thrust stays close to 230 tonnes range
Miss
3) Diameter of BFR (1st stage)
Range 12.5m-15m, best estimate 15m 1st stage
Miss
4) Total Raptor 1st stage thrust (sl)
60 Meganewtons and T/W > 1.3
Miss
5) LAS Architecture
a) No LAS - BFS is the escape mechanism
b) Traditional LAS - above BFS and is nominally jettisoned during launch phase
c) BFS contains smaller 'ejection pod' where humans reside during launch
d) Other, non-traditional LAS design
Best guess is (a)
HIT!!!
6) Shape and Landing Mode of BFS
a) Capsule (perhaps elongated), w/ TPS on base
b) Cylindrical or biconic - horizontal landing
c) Cylindrical or biconic - vertical landing
d) Other
Going with (c), definitely no horizontal landing
HIT!!!
7) Mars and Earth return
a) BFS does direct entry into Mars and Earth atmosphere
b) BFS does orbital capture before performing entry burn and landing
c) Same as b, but upon Earth return, stays in orbit for next synod
(a)
HIT
8) Use of non-chemical thrust
a) Not part of the plan
b) Will use SEP for some/all of the big transits
c) All chemical for now, but plans to incorporate SEP down the road
(c) strongly favor
a seems more correct
Can anyone think of more/better questions?
Predict Musk will miss 1st crewed landing by >= 3 synods
UNKNOWN but too easy
5-6 Rvacs on BFS stage 2
HIT sort of. Missed on R SL engines. frakked at self for not seeing that.
Raptor sea level will have 10s of tonnes thrust more than the 230 tonnes mentioned by Elon
HIT!!!
Entire BFR/BFS GLOW masses under 5.000 tonnes; my estimate ~4,500
BIG miss!
Height of BFR/BFS stack under 120m; my estimate <100m
MISS
Cargo version, tanker version, crewed version of BFS
HIT
1st crewed landing on Mars 8-12 humans planned
UNKNOWN
Just over 48 hours until Musk makes fools of us
Got that right!
everyone else was basing their prediction on rumors and L2 information that I don't have access too.
Yes, I meant cylindrical. Although the three "fins" do alter the shape slightly, so I wonder if they have seen some aerodynamic benefit of stretching them so far forward.
Although I never came out and specifically said so, I've favoured what I called a "semi-lifting body", that is something that gets some lift from its body shape and maybe has small fins but is not a traditional lifting body shape.
It looks like the Ship qualifies as a semi-lifting body.
I underestimated the Tanker fuel load so expected 5-8 Tanker flights to refuel the Ship.