I think we will see a FH debut only after SpaceX has landed a F9 booster at the Cape (RTLS). They will attempt to recover all three boosters of the FH Demo flight. There is simply no good reason to speed up the FH Demo flight (except us waiting for it... :) ) ...That's not true. There's a big portion of the launch market, the part with the most revenue in it (or at least comparable to F9's market), that can't be addressed until Falcon Heavy has flown.
I think we will see a FH debut only after SpaceX has landed a F9 booster at the Cape (RTLS). They will attempt to recover all three boosters of the FH Demo flight. There is simply no good reason to speed up the FH Demo flight (except us waiting for it... :) ) ...That's not true. There's a big portion of the launch market, the part with the most revenue in it (or at least comparable to F9's market), that can't be addressed until Falcon Heavy has flown.
The sooner they fly Falcon Heavies, the sooner they can compete for all the defense payloads. F9 can only serve about half of those, and the revenue on that lighter half is significantly less than the heavier half.
There are also a lot of commercial payloads which they can't compete for without Falcon Heavy, the highest revenue commercial payloads.
...lots of very good reasons to speed up the Falcon Heavy Demo (although even more crucial will be ensuring it launches without major failure). They may still do the FH flight after the first RTLS F9s, but that's a different issue. (I personally think FH probably won't launch until 2016.)
I think we will see a FH debut only after SpaceX has landed a F9 booster at the Cape (RTLS). They will attempt to recover all three boosters of the FH Demo flight. There is simply no good reason to speed up the FH Demo flight (except us waiting for it... :) ) ...That's not true. There's a big portion of the launch market, the part with the most revenue in it (or at least comparable to F9's market), that can't be addressed until Falcon Heavy has flown.
The sooner they fly Falcon Heavies, the sooner they can compete for all the defense payloads. F9 can only serve about half of those, and the revenue on that lighter half is significantly less than the heavier half.
There are also a lot of commercial payloads which they can't compete for without Falcon Heavy, the highest revenue commercial payloads.
...lots of very good reasons to speed up the Falcon Heavy Demo (although even more crucial will be ensuring it launches without major failure). They may still do the FH flight after the first RTLS F9s, but that's a different issue. (I personally think FH probably won't launch until 2016.)
...link (http://forum.nasaspaceflight.com/index.php?topic=27871.msg1332228#msg1332228)
Falcon Heavy is not using cross-feed. They're not developing it. I was at SpaceX several months ago and asked about cross-feed and was told by one of the people working on the rocket that they are not developing it.
...
@Blackstar posted some interesting tidbits about the Falcon Heavy in the "Proposed Europa Missions" thread.Quote...link (http://forum.nasaspaceflight.com/index.php?topic=27871.msg1332228#msg1332228)
Falcon Heavy is not using cross-feed. They're not developing it. I was at SpaceX several months ago and asked about cross-feed and was told by one of the people working on the rocket that they are not developing it.
...
Back in October I had lunch sitting next to a guy from Aerojet who was working on an upper stage for (I think) Falcon Heavy to enable SpaceX to compete for the Solar Probe Plus mission. Dunno if that's gone public anywhere, but it may be mentioned elsewhere on this site. Anyway, they're locked out of a number of missions unless they upgrade their hardware.
One other thing: Falcon Heavy is not using cross-feed. They're not developing it. I was at SpaceX several months ago and asked about cross-feed and was told by one of the people working on the rocket that they are not developing it. It's a potential upgrade if somebody pays for it, but they're not doing the development. So you shouldn't use it in your calculations.
Hopefully this is the right thread for this question.
Has SpaceX said anything for sure about if they plan to upgrade LC-40 to handle FH?
I'd assume they would, perhaps after 39A is up and fully functional becuase an upgrade to LC-40 like that will mean it'll be out of commission for awhile, and their manifest is packed right now.
But I don't recall anything for sure about it, other than an old comment by Elon saying they might build a FH HIB at a 90 degree angle to the F9 HIB. I think that was before 39A was in the mix, so they could keep flying F9 while upgrading to FH. With 39A operational, I think they can just fly from there and tear down the F9 HIB and build the FH HIB there
Or will they just leave LC-40 launching F9 only?
Hopefully this is the right thread for this question.
Has SpaceX said anything for sure about if they plan to upgrade LC-40 to handle FH?
I'd assume they would, perhaps after 39A is up and fully functional becuase an upgrade to LC-40 like that will mean it'll be out of commission for awhile, and their manifest is packed right now.
But I don't recall anything for sure about it, other than an old comment by Elon saying they might build a FH HIB at a 90 degree angle to the F9 HIB. I think that was before 39A was in the mix, so they could keep flying F9 while upgrading to FH. With 39A operational, I think they can just fly from there and tear down the F9 HIB and build the FH HIB there
Or will they just leave LC-40 launching F9 only?
You know, people keep saying that, neglecting the fact that both Atlas V and Delta IV have flown MANY variants:I think we will see a FH debut only after SpaceX has landed a F9 booster at the Cape (RTLS). They will attempt to recover all three boosters of the FH Demo flight. There is simply no good reason to speed up the FH Demo flight (except us waiting for it... :) ) ...That's not true. There's a big portion of the launch market, the part with the most revenue in it (or at least comparable to F9's market), that can't be addressed until Falcon Heavy has flown.
The sooner they fly Falcon Heavies, the sooner they can compete for all the defense payloads. F9 can only serve about half of those, and the revenue on that lighter half is significantly less than the heavier half.
There are also a lot of commercial payloads which they can't compete for without Falcon Heavy, the highest revenue commercial payloads.
...lots of very good reasons to speed up the Falcon Heavy Demo (although even more crucial will be ensuring it launches without major failure). They may still do the FH flight after the first RTLS F9s, but that's a different issue. (I personally think FH probably won't launch until 2016.)
Yes, these are good points. But then, competing for defense payloads means going for certification, and this means you need to have a near-final version of the vehicle you want to certify ready, otherwise the clock is reset every time you re-introduce a new feature (e.g., cross-feed, fuel densification, uprated Merlins, RTLS etc.). From this point of view, it might be better to have all that ready for FH Demo, which will take time. ...
If that guy's project is official it sounds like Aerojet Rocketdyne is trying to build a future for itself that doesn't rely on ULA. I was going to write that this was surprising back-stabbing of its close business partner ULA but then I remembered that ULA has already cheated on that marriage with their funding of XCOR's RL-10 competitor.
From the Europa science thread Blackstar posted some juicy Falcon Heavy info:A RL10 upper stage would make all the difference on BLEO missions. It is not just RL10 that makes Centuar expensive but also the design. SpaceX should be able to make a considerable cheaper equivalent and maybe IVF in the process. At of the technology would be directly applicable to a methane stage.Back in October I had lunch sitting next to a guy from Aerojet who was working on an upper stage for (I think) Falcon Heavy to enable SpaceX to compete for the Solar Probe Plus mission. Dunno if that's gone public anywhere, but it may be mentioned elsewhere on this site. Anyway, they're locked out of a number of missions unless they upgrade their hardware.
I wonder what sort of fuel that upper stage would use. Aerojet Rocketdyne has a suitable hydrogen engine (RL-10), various hypergolic engines (e.g. Shuttle OMS) and solids experience (e.g. Orion FTS jettison motor) so there are a lot of plausible options.
If that guy's project is official it sounds like Aerojet Rocketdyne is trying to build a future for itself that doesn't rely on ULA. I was going to write that this was surprising back-stabbing of its close business partner ULA but then I remembered that ULA has already cheated on that marriage with their funding of XCOR's RL-10 competitor.QuoteOne other thing: Falcon Heavy is not using cross-feed. They're not developing it. I was at SpaceX several months ago and asked about cross-feed and was told by one of the people working on the rocket that they are not developing it. It's a potential upgrade if somebody pays for it, but they're not doing the development. So you shouldn't use it in your calculations.
Without cross feed Falcon Heavy can send 45 tonnes to LEO according to http://www.spacex.com/falcon-heavy. I suppose that's the best estimate of Falcon Heavy performance to LEO we have currently?
SpaceX's plans for the heaviest payloads involve the BFR so it makes sense that they don't see a need to optimize FH's capacity.
I wonder if the Aerojet guy that Blackstar ate lunch next to was working on the NASA project to build a common upper stage discussed in this thread: http://forum.nasaspaceflight.com/index.php?topic=35144 . Or maybe this Falcon Heavy upper stage project evolved out of that NASA-funded project?
The engine would be DOD certified if not the stage, one plus going for.
There is a lot I do not understand on the Falcon Heavy SpaceX page. It says the core engines throttle down shortly after liftoff (presumably to limit max Q), and throttle up after booster separation. This implies they stay throttled down between those two events. It seems to me that once past max Q, I'd want those engines right back up to 100% again to reduce gravity losses. Maybe they just don't mention those two more throttle moves.
The core is throttled down until separation to leave as much propellant as possible in the core at separation. This improves performance.
The core is throttled down until separation to leave as much propellant as possible in the core at separation. This improves performance.
Right, that's the part I don't understand.
I don't have access to a simulator anymore, but when I used to run simulations, it was always better to burn as much propellant as soon as possible, subject to the limitations of the airframe, and the staging effect. MaxQ is the first limitation, most rockets have to throttle down a bit to avoid getting too fast in the lower atmosphere. I know that the Space Shuttle, at least, throttled all the way back up after maxQ. Typically aero loads have an early maximum just after mach 1, less than 1/3 of the way through the first stage burn, and then trend down from there, as air density drops off faster than velocity^2 increases. The second limitation is maximum acceleration. I know the Saturn V and the Falcon 9 both have to throttle down their first stages shortly before engine cutoff to avoid too much acceleration. But that limitation comes in fairly late in the first stage burn, in the last 20-30 seconds. Finally, staging helps by eliminating the weight of empty tankage and engines that have had to be turned off.
The longer you wait around to burn off your propellant, the more gravity loss you suffer, especially early in the flight. This is a strong effect. There must be some really strong reason to throttle down the core stage on an F9 or D4H launch, to overcome the gravity loss effect. What is it?
I thought maybe the stress loads on the vehicle would be smaller with less thrust, but upon further thought I think stress loads on the interconnect between the boosters and core are minimized by minimizing the difference between core thrust and booster thrust.
Schillings presumably uses cross-feed;
Hope this is the right thread to post in.
I think I have a scheme which gets the Falcon Heavy most of the benefit of crossfeed with only a small portion of the risk.
At liftoff, five of the core engines slurp the core tanks, and two of the core engines run off each of the booster tanks. The five core engines throttle down as necessary to limit max Q and max acceleration. So far, standard crossfeed.
If the center five engines can get down to 50% throttle, the boosters run out of propellant (certainly in the reuse case) before we hit maximum acceleration. So, the booster engines run at 100% for their whole trip. When the booster engine cutoff is signaled around +162 seconds, all 22 engines fed by booster tanks cut off.
The four dead engines on the core stage are never restarted. There is no crossfeed valve. This is the extent of the idea. The five remaining engines ramp back up to 100%, and stay at 100% until core engine cutoff at around 310 seconds into flight.
The new hardware needed are new propellant plenums for the core and booster stages, and unions between the adjacent fuel and oxidizer manifolds that can be isolated, drained, and disconnected in flight.
This scheme is not as good as full-blown crossfeed. It dumps the empty booster weight almost as early as possible. Perfect crossfeed (100% full core at booster engine cutoff) would dump them 15 seconds earlier, gaining perhaps 60-70 m/s more delta-V. Perfect crossfeed would also have all nine core engines running after booster separation, which would make the core engine cutoff about 50 seconds earlier, which would reduce gravity losses. I'm not sure how much delta-V that is worth, perhaps 100 m/s, perhaps more.
This scheme is better than no crossfeed. Without crossfeed, the booster engine cutoff happens around 200 seconds (assuming reuse). The extra momentum carried away by the separated boosters costs 160-180 m/s delta-V compared to my scheme. Also, either the boosters will have to be throttled, or core engines will have to be shut down to avoid overacceleration before booster separation. The first gets the boosters going even faster when they separate, and the second requires either a midflight restart (with the payload still attached) or delaying core engine cutoff, either of which has penalties.
There is a lot I do not understand on the Falcon Heavy SpaceX page. It says the core engines throttle down shortly after liftoff (presumably to limit max Q), and throttle up after booster separation. This implies they stay throttled down between those two events. It seems to me that once past max Q, I'd want those engines right back up to 100% again to reduce gravity losses. Maybe they just don't mention those two more throttle moves.
The diagram also shows the boosters about 15% larger than the core stage, the upper stage identical in propellant mass to the F9, and the total liftoff mass 54,702 kg shy of 3x the Falcon 9. First, why would the FH need a faster liftoff acceleration than the F9? By loading more propellant, you could get more payload. Second, I'd expect the booster propellant load to be approximately the same as the F9's first and second stage propellant loads combined, but the increase is only 2/3 of that. There is plenty of room to stretch those tanks further.
From the Europa science thread Blackstar posted some juicy Falcon Heavy info:
I wonder what sort of fuel that upper stage would use. Aerojet Rocketdyne has a suitable hydrogen engine (RL-10), various hypergolic engines (e.g. Shuttle OMS) and solids experience (e.g. Orion FTS jettison motor) so there are a lot of plausible options.
If that guy's project is official it sounds like Aerojet Rocketdyne is trying to build a future for itself that doesn't rely on ULA. I was going to write that this was surprising back-stabbing of its close business partner ULA but then I remembered that ULA has already cheated on that marriage with their funding of XCOR's RL-10 competitor.
about cross-feed and was told by one of the people working on the rocket that they are not developing it. It's a potential upgrade if somebody pays for it, but they're not doing the development. So you shouldn't use it in your calculations.
If hypergolic's are suitable would/could SpaceX use a kick stage with super dracos?
Hope this is the right thread to post in.
I think I have a scheme which gets the Falcon Heavy most of the benefit of crossfeed with only a small portion of the risk.
At liftoff, five of the core engines slurp the core tanks, and two of the core engines run off each of the booster tanks. The five core engines throttle down as necessary to limit max Q and max acceleration. So far, standard crossfeed.
If the center five engines can get down to 50% throttle, the boosters run out of propellant (certainly in the reuse case) before we hit maximum acceleration. So, the booster engines run at 100% for their whole trip. When the booster engine cutoff is signaled around +162 seconds, all 22 engines fed by booster tanks cut off.
The four dead engines on the core stage are never restarted. There is no crossfeed valve. This is the extent of the idea. The five remaining engines ramp back up to 100%, and stay at 100% until core engine cutoff at around 310 seconds into flight.
The new hardware needed are new propellant plenums for the core and booster stages, and unions between the adjacent fuel and oxidizer manifolds that can be isolated, drained, and disconnected in flight.
It looks like you can bite the bullet and just mount those two engines on the boosters instead of the core. Then at booster separation you do not need any disconnects and you are not carrying dead weight on the core.
The Soyuz rocket uses cross feed and has since the 1950's. They just send a signal and they disconnect.
I read somewhere that the 4 strap ons burn with the core and transfer fuel to the core at the same time. Then they drop off and the core continues. Maybe they don't since they are smaller. However I read that years ago.
Purely FWIW, I think that RL-10C's thrust is too low for use on Falcon Heavy, even if you use cross-feed on the core and boosters. Staging is just too low and slow (because of RTLS performance limitations). Given that limitation and given that I think that Messrs Musk and Bezos would prefer to gouge out their own eyes than work together, the only hydrolox engine likely to be used on Falcon Heavy, IMHO at least (for a certain percentage of 'likely' anyway) is MB-60.
Hopefully this is the right thread for this question.
Has SpaceX said anything for sure about if they plan to upgrade LC-40 to handle FH?
I'd assume they would, perhaps after 39A is up and fully functional becuase an upgrade to LC-40 like that will mean it'll be out of commission for awhile, and their manifest is packed right now.
But I don't recall anything for sure about it, other than an old comment by Elon saying they might build a FH HIB at a 90 degree angle to the F9 HIB. I think that was before 39A was in the mix, so they could keep flying F9 while upgrading to FH. With 39A operational, I think they can just fly from there and tear down the F9 HIB and build the FH HIB there
Or will they just leave LC-40 launching F9 only?
They have mentioned doing it, but plans change, and most of those comments were pre-39A acquisition. Now that they have 39A I don't see them being in a hurry do it. Perhaps never. But it also depends on two additional factors:
- What will be the ratio of F9 to FH launches going forward?
- How quickly will the Texas launch pad come online (presumably built to support FH from the beginning)
There is a simple question and simple answer in regards to Falcon Heavy cross-feed development.
Q: Why would SpaceX develop cross-feed for Falcon Heavy?
A: Because it is necessary to meet sufficient demand for customer payload requirements.
It isn't going to happen just because. It would happen because it was necessary for enough paying customers that would justify the development cost, or a customer needed it badly enough to pay for it. And it isn't really clear that that is something that is going to happen in the next 15 years. SpaceX themselves won't need it: their long-term Mars plans revolve around BFR. Any reusable SpaceX BFR in 10-15 years would make a cross-feed FH obsolete.
So who is the customer? And don't say Bigelow. At this point, it seems a rather long shot that he ever pulls together several hundred million dollars to develop, build and fly even one of the smaller modules, much less a larger one. And, again, if you're the *only* customer for a cross-feed Falcon Heavy, guess who gets to foot the tab for development?
Purely FWIW, I think that RL-10C's thrust is too low for use on Falcon Heavy, even if you use cross-feed on the core and boosters. Staging is just too low and slow (because of RTLS performance limitations). Given that limitation and given that I think that Messrs Musk and Bezos would prefer to gouge out their own eyes than work together, the only hydrolox engine likely to be used on Falcon Heavy, IMHO at least (for a certain percentage of 'likely' anyway) is MB-60.
1 - Could the M1D thrust upgrade and prop densification fully compensate no cross feed ?
2 - Just because SpaceX mentioned a few times flying people to Mars on Falcon Heavy doesn't mean it's still in the cards. SpaceX has shown multiple times the ability to evolve its plans. Don't get hung up on words from many years ago if they don't get reinforced (with more words or actions).
http://www.spacex.com/falcon-heavy: Falcon Heavy was designed from the outset to carry humans into space and restores the possibility of flying missions with crew to the Moon or Mars.
There is a simple question and simple answer in regards to Falcon Heavy cross-feed development.When will people stop thinking/stating with certainty that anything is CLEAR in the future, especially 10-15 years hence? In the year 2000, would you or anyone have been certain that we'd be where we are in space launch business in 2015?
Q: Why would SpaceX develop cross-feed for Falcon Heavy?
A: Because it is necessary to meet sufficient demand for customer payload requirements.
It isn't going to happen just because. It would happen because it was necessary for enough paying customers that would justify the development cost, or a customer needed it badly enough to pay for it. And it isn't really clear that that is something that is going to happen in the next 15 years. SpaceX themselves won't need it: their long-term Mars plans revolve around BFR. Any reusable SpaceX BFR in 10-15 years would make a cross-feed FH obsolete.
So who is the customer? And don't say Bigelow. At this point, it seems a rather long shot that he ever pulls together several hundred million dollars to develop, build and fly even one of the smaller modules, much less a larger one. And, again, if you're the *only* customer for a cross-feed Falcon Heavy, guess who gets to foot the tab for development?
The multi-core BFR posts are out of date. Musk has made fairly clear recently (in the MIT talk, I believe?) that the BFR will be single-core (though they were looking at multi-core before), I think due to operational reasons.
The multi-core BFR posts are out of date. Musk has made fairly clear recently (in the MIT talk, I believe?) that the BFR will be single-core (though they were looking at multi-core before), I think due to operational reasons.
[–]FoxhoundBat 12 points 20 minutes ago
In order to use the full MCT design (100 passengers), will BFR be one core or 3 cores?
[–]ElonMuskOfficial 6 points a minute ago
At first, I was thinking we would just scale up Falcon Heavy, but it looks like it probably makes more sense just to have a single monster boost stage.
Well, it's a 3-sigma statement...The multi-core BFR posts are out of date. Musk has made fairly clear recently (in the MIT talk, I believe?) that the BFR will be single-core (though they were looking at multi-core before), I think due to operational reasons.
It was at the reddit AMA :Quote from: https://www.reddit.com/r/IAmA/comments/2rgsan/i_am_elon_musk_ceocto_of_a_rocket_company_ama/cnfpn7r[–]FoxhoundBat 12 points 20 minutes ago
In order to use the full MCT design (100 passengers), will BFR be one core or 3 cores?
[–]ElonMuskOfficial 6 points a minute ago
At first, I was thinking we would just scale up Falcon Heavy, but it looks like it probably makes more sense just to have a single monster boost stage.
" but it looks like it probably makes more sense "
Hardly conclusive. I wouldnt dissmiss the tri core concept altogether based on a triple disclaimed sentence.
When will people stop thinking/stating with certainty that anything is CLEAR in the future, especially 10-15 years hence? In the year 2000, would you or anyone have been certain that we'd be where we are in space launch business in 2015?
Stop, please. (You don't have a clue. None of us has a clue.)
Has anybody seen any recent information on whether or not SpaceX will develop a larger fairing for FH? This to me seems like another item that would only be paid for if a customer wants it, but it also seems like a much cheaper upgrade than cross-feeding.I requested just that information from SpaceX a while back... for a project I'm working.
If you go full throttle on the core and the two boosters you are hauling the mass of all three until you run out of propellant. If you throttle back on the core it still has propellant after you dump the mass of the boosters.
I thought of perhaps another way to do full cross-feed without affecting the bottom of the rocket at all. The boosters are taller than the core stage. You could arrange for each booster to have two LOX tanks and two kerosene tanks, one set half the size of the other. The smaller kerosene tank sits above the main kerosene tank, and significantly, the bottom of the smaller booster kerosene tank sits above the top of the core kerosene tank. It drains into the core kerosene tank by gravity feed. Similarly, the bottom of the smaller booster LOX tank sits above the top of the core LOX tank, and drains by gravity feed as well. Note that gravity feed is fairly powerful, as the accelerations before booster burnout are over 3G.
This scheme is perfect cross-feed, but with the added weight of two more tank bulkheads and the cross piping and unions. No special lines are needed to blow the cross piping clear of propellant before closing the valves and disconnecting. The pressurization system of the booster propellant tank can positively drain all the fluid into the core tanks before separation.
Given that at separation you'd have nearly full core tanks, the propellant burn before separation is always the same and any throttling down just postpones the booster separation and increases gravity losses. Once past max Q you'd ramp all the engines to 100% and leave them there through booster ECO at around 155 seconds.
The longer you wait around to burn off your propellant, the more gravity loss you suffer, especially early in the flight. This is a strong effect. There must be some really strong reason to throttle down the core stage on an F9 or D4H launch, to overcome the gravity loss effect. What is it?
I thought maybe the stress loads on the vehicle would be smaller with less thrust, but upon further thought I think stress loads on the interconnect between the boosters and core are minimized by minimizing the difference between core thrust and booster thrust.
Purely FWIW, I think that RL-10C's thrust is too low for use on Falcon Heavy, even if you use cross-feed on the core and boosters. Staging is just too low and slow (because of RTLS performance limitations).
I thought of perhaps another way to do full cross-feed without affecting the bottom of the rocket at all. The boosters are taller than the core stage. You could arrange for each booster to have two LOX tanks and two kerosene tanks, one set half the size of the other. The smaller kerosene tank sits above the main kerosene tank, and significantly, the bottom of the smaller booster kerosene tank sits above the top of the core kerosene tank. It drains into the core kerosene tank by gravity feed. Similarly, the bottom of the smaller booster LOX tank sits above the top of the core LOX tank, and drains by gravity feed as well. Note that gravity feed is fairly powerful, as the accelerations before booster burnout are over 3G.
This scheme is perfect cross-feed, but with the added weight of two more tank bulkheads and the cross piping and unions. No special lines are needed to blow the cross piping clear of propellant before closing the valves and disconnecting. The pressurization system of the booster propellant tank can positively drain all the fluid into the core tanks before separation.
Given that at separation you'd have nearly full core tanks, the propellant burn before separation is always the same and any throttling down just postpones the booster separation and increases gravity losses. Once past max Q you'd ramp all the engines to 100% and leave them there through booster ECO at around 155 seconds.
that doesn't make any sense. The few feet of head pressure is not going to matter and it is a waste to have separate tanks
also, there is no need to blow the lines clear before closing valves.
There is a simple question and simple answer in regards to Falcon Heavy cross-feed development.When will people stop thinking/stating with certainty that anything is CLEAR in the future, especially 10-15 years hence? In the year 2000, would you or anyone have been certain that we'd be where we are in space launch business in 2015?
Q: Why would SpaceX develop cross-feed for Falcon Heavy?
A: Because it is necessary to meet sufficient demand for customer payload requirements.
It isn't going to happen just because. It would happen because it was necessary for enough paying customers that would justify the development cost, or a customer needed it badly enough to pay for it. And it isn't really clear that that is something that is going to happen in the next 15 years. SpaceX themselves won't need it: their long-term Mars plans revolve around BFR. Any reusable SpaceX BFR in 10-15 years would make a cross-feed FH obsolete.
So who is the customer? And don't say Bigelow. At this point, it seems a rather long shot that he ever pulls together several hundred million dollars to develop, build and fly even one of the smaller modules, much less a larger one. And, again, if you're the *only* customer for a cross-feed Falcon Heavy, guess who gets to foot the tab for development?
Stop, please. (You don't have a clue. None of us has a clue.)
Purely FWIW, I think that RL-10C's thrust is too low for use on Falcon Heavy, even if you use cross-feed on the core and boosters. Staging is just too low and slow (because of RTLS performance limitations).
If someone wanted to replace Falcon Heavy's upper stage with a hydrogen stage it would be reasonable to use a cluster of several RL-10s to get enough thrust (similar to DIRECT and ACES plans) so you can't rule out RL-10 for thrust reasons. Also IMHO a more likely (but still unlikely) use of hydrogen with Falcon Heavy would be use of an existing ULA RL-10-based stage (i.e. DCSS or Centaur) as an in-space kick stage on top of the existing Falcon upper stage (not a replacement), a task for which a single RL-10 is the right size.
Purely FWIW, I think that RL-10C's thrust is too low for use on Falcon Heavy, even if you use cross-feed on the core and boosters. Staging is just too low and slow (because of RTLS performance limitations). Given that limitation and given that I think that Messrs Musk and Bezos would prefer to gouge out their own eyes than work together, the only hydrolox engine likely to be used on Falcon Heavy, IMHO at least (for a certain percentage of 'likely' anyway) is MB-60.
Are there any known plans to install (or use existing) LH2 infrastructure at SLC-4E or 39A? I just don't see SpaceX dealing with LH2 at all. LH2 is extremely high cost, high maintenance technology that limits materials, design and processes not just on the rocket but on the pad.
Purely FWIW, I think that RL-10C's thrust is too low for use on Falcon Heavy, even if you use cross-feed on the core and boosters. Staging is just too low and slow (because of RTLS performance limitations). Given that limitation and given that I think that Messrs Musk and Bezos would prefer to gouge out their own eyes than work together, the only hydrolox engine likely to be used on Falcon Heavy, IMHO at least (for a certain percentage of 'likely' anyway) is MB-60.
Are there any known plans to install (or use existing) LH2 infrastructure at SLC-4E or 39A? I just don't see SpaceX dealing with LH2 at all. LH2 is extremely high cost, high maintenance technology that limits materials, design and processes not just on the rocket but on the pad.
Just to clarify, I don't see SpaceX working towards a high-energy chemical upper stage any time soon. Methalox is fine for their next-gen propulsion development, which will be okay up to cis-lunar. Further than that, I suspect that MCT will be solar-electric or something else even more exotic thatDr EvilElon is hiding up his sleeve. All I was saying is that, with BE-3 ruled out for personality clash reasons, MB-60 is the only hydrolox engine suitably sized for an alternate high-energy upper stage for Falcon Heavy, IMHO at least. An RL-10C cluster would probably be slightly outside what could be fit into the interstage
Regarding the possibility of an RL-10 based upper stage..... I think an RL10 engine is in the price range or above the price of an entire F9 upper stage. Delta IV economics won't fly on a SpaceX rocket.
I wrote this in the other thread, but an RL-10 doesn't make sense. In all reality, they probably were just talking about a solid kick stage.++
Only New Horizons I think has higher delta-v.I believe this mission also used a solid motor kick stage.
Doesn't need to be a particularly high Isp kick stage if the probe is small. Just a typical small solid rocket motor with a vacuum-optimized nozzle.Regarding the possibility of an RL-10 based upper stage..... I think an RL10 engine is in the price range or above the price of an entire F9 upper stage. Delta IV economics won't fly on a SpaceX rocket.
A low thrust Rl-10 or otherwise high ISP kick stage should fit in F9 fairing with a BEO payload. This could be a good combination for europa mission with better performance than DIVH still at far lower price.
As others have already pointed out, the info Blackstar related is most likely referring to a solid kick stage. Just like the Star 48 used with the New Horizons mission. Solar Probe Plus certainly doesn't have the budget to design, certify, and build a new high energy liquid kick stage.
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......, Aerojet and Orbital were asked to design a reference mission for NASA that shows how quickly the companies’ chosen upper stages — whatever those may be — could propel a payload beyond Earth assuming launch by a rocket approximately as powerful as Space Exploration Technologies Corp.’s (SpaceX) Falcon 9 v1.1.
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Punder the main issue with that is not the docking but expecting the Lox not to evaporate too quickly in whatever upper stage(s) it is in. This means both US need at least very efficient passive cooling including attitude control to keep the craft oriented so it radiates heat away from any radiative heat source and has radiative shielding that stays in between it and the major radiative heat sources (Earth, Sun, Moon). This may be a little complex considering they will be docking/birthing nose to nose.
A much smaller side issue is that you will need to dock the two craft top to top which introduces complexity to the final payload needing a frame around it with a docking port, the eventual ejection of that frame and being able to sustain negative g's in the case of the dual KeroLox US propulsion.
Given that we are using the existing US, we have to also be concerned of the stresses on the docking connection. For the dual KeroLox version I suggest the least troublesome arrangement would be a bolting connection like the birthing adaptor AND that release from the payload and it's original US by the 2nd upper stage take place from the frame that distributes the pressure from the 2nd upper stages thrusting so that you are left after the 2nd upper stage separates with just the original upper stage, the final craft and whatever adapting connector you have between them.
Look, I'm not a mechanical engineer, and I perfectly realize there are many factors here of which I'm ignorant. There will have to be precise RCS, detection circuitry, orbital analysis, boiloff reserves, and on and on. I personally just can't think of any showstopper, or really any vastly difficult/expensive problem, that blows this idea away. In fact virtually all of it has been done before, other than the engine bell probe-and-drogue. And on that subject... Isn't the engine bell of a rocket the PERFECT place to direct some force?
Look, I'm not a mechanical engineer, and I perfectly realize there are many factors here of which I'm ignorant. There will have to be precise RCS, detection circuitry, orbital analysis, boiloff reserves, and on and on. I personally just can't think of any showstopper, or really any vastly difficult/expensive problem, that blows this idea away. In fact virtually all of it has been done before, other than the engine bell probe-and-drogue. And on that subject... Isn't the engine bell of a rocket the PERFECT place to direct some force?
NO. :) The nozzle edges are usually very thin and fragile. At least for upper stage engines with nozzle extensions.
Purely FWIW, I think that RL-10C's thrust is too low for use on Falcon Heavy, even if you use cross-feed on the core and boosters. Staging is just too low and slow (because of RTLS performance limitations). Given that limitation and given that I think that Messrs Musk and Bezos would prefer to gouge out their own eyes than work together, the only hydrolox engine likely to be used on Falcon Heavy, IMHO at least (for a certain percentage of 'likely' anyway) is MB-60.
Are there any known plans to install (or use existing) LH2 infrastructure at SLC-4E or 39A? I just don't see SpaceX dealing with LH2 at all. LH2 is extremely high cost, high maintenance technology that limits materials, design and processes not just on the rocket but on the pad.
Just to clarify, I don't see SpaceX working towards a high-energy chemical upper stage any time soon. Methalox is fine for their next-gen propulsion development, which will be okay up to cis-lunar. Further than that, I suspect that MCT will be solar-electric or something else even more exotic thatDr EvilElon is hiding up his sleeve. All I was saying is that, with BE-3 ruled out for personality clash reasons, MB-60 is the only hydrolox engine suitably sized for an alternate high-energy upper stage for Falcon Heavy, IMHO at least. An RL-10C cluster would probably be slightly outside what could be fit into the interstage
Okay, so forget that part. (And I do appreciate the smiley.) Use the ring trusses as the sole thrust structure and align/dock the stages some other way. Anything else in the "big concept" that's utterly nonsensical?
Okay, so forget that part. (And I do appreciate the smiley.) Use the ring trusses as the sole thrust structure and align/dock the stages some other way. Anything else in the "big concept" that's utterly nonsensical?
Back to the boil off issue, if you are only looking for a single impulse from the two stages then, as long as you dock them and get that impulse in the first orbit you should not have lost too much. However, if one of the reasons you wanted all this extra capacity for a probe included a high energy orbit change later on then you still need to solve the boil off (and operation power issues since the main solar panels won't be deployed until after the last high energy burn).
I am not seeing how this gets you a solution that is cheaper than 3'rd stage.
Maybe we should stop trying to pack all the energy for any given mission into a single launch. ...For anything like this, multiple launches don't make any sense.
Okay, so forget that part. (And I do appreciate the smiley.) Use the ring trusses as the sole thrust structure and align/dock the stages some other way. Anything else in the "big concept" that's utterly nonsensical?
Back to the boil off issue, if you are only looking for a single impulse from the two stages then, as long as you dock them and get that impulse in the first orbit you should not have lost too much. However, if one of the reasons you wanted all this extra capacity for a probe included a high energy orbit change later on then you still need to solve the boil off (and operation power issues since the main solar panels won't be deployed until after the last high energy burn).
I am not seeing how this gets you a solution that is cheaper than 3'rd stage.
There may not be a solution. This is what you guys call "hand-waving" after all! But:
1. The 2nd flight US has a bunch of propellant. It had no payload, after all, other than the weight of the docking mechanism and any rendezvous sensors/computers. Maybe it only had a small expendable nosecone rather than a big heavy payload shroud. Seems like that would make for much larger payload than a 3rd stage added to a single launch.
2. Common stages, common engines, common propellants, common software, common power, common procedures, common personnel. All from within SpaceX. No additional companies, personnel, procedures, safety concerns, or pad mods for a non-common stage. Each interaction between dissimilar items eats money and poops paper + lost time.
3. The state of the art wrt cryo boiloff may improve. People have been working on depot tech for awhile now.
Maybe we should stop trying to pack all the energy for any given mission into a single launch. ...For anything like this, multiple launches don't make any sense.
I'm not sure of the exact figures, but I'm pretty sure a small solid rocket kick motor is in the (perhaps low) single-digits millions of dollars range. Another launch (even fully reusable!) would be more than that.
EDIT: A Star-48 is around $6 million, roughly: http://forum.nasaspaceflight.com/index.php?topic=33084.msg1109467#msg1109467
The Falcon 9 - Falcon Heavy upper stage would need a major upgrade to enable it for refuelling. A very much longer loiter time would be one of those. I also would worry more about RP-1 freezing than about some LOX boiloff.
The Falcon 9 - Falcon Heavy upper stage would need a major upgrade to enable it for refuelling. A very much longer loiter time would be one of those. I also would worry more about RP-1 freezing than about some LOX boiloff.
Why major, why couldn't the the fuel depot attach to all the existing umbilical points? giving it the ability to circulate propellants (thereby controlling temps) and to provide power to the US and payload to keep it fresh?
The Falcon 9 - Falcon Heavy upper stage would need a major upgrade to enable it for refuelling. A very much longer loiter time would be one of those. I also would worry more about RP-1 freezing than about some LOX boiloff.
Why major, why couldn't the the fuel depot attach to all the existing umbilical points? giving it the ability to circulate propellants (thereby controlling temps) and to provide power to the US and payload to keep it fresh?
At the very least the stage must remain alive and active until docking is achieved.
Also connections to the stage made in the HIF are not the same as needed for automated docking to a fuel depot. It needs a redesign.
I have the very strong opinion it won't be done for a Falcon upper stage. It will need to be done for the BFR upper stage. But that would be designed from the beginning with this in mind.
F1US and it's Kestrel engine flew several times. It could provide the basis for a kick stage, though I believe it's only ~4t. Extra mass would be no problem as gravity losses are not an issue for a kick stage.
Maybe we should stop trying to pack all the energy for any given mission into a single launch. ...For anything like this, multiple launches don't make any sense.
I'm not sure of the exact figures, but I'm pretty sure a small solid rocket kick motor is in the (perhaps low) single-digits millions of dollars range. Another launch (even fully reusable!) would be more than that.
EDIT: A Star-48 is around $6 million, roughly: http://forum.nasaspaceflight.com/index.php?topic=33084.msg1109467#msg1109467
But a Star-48 wouldn't accomplish the overriding goal of two launches and LEO rendezvous and ...
Wait, that's not the goal?
That's my main concern. Concepts long discarded may become feasible with that kind of cost savings. Maybe not this particular concept from some nobody on the Internet, but you see my point.
This!!!
Bring out those retired visionaries from both sides of the Atlantic and dust off those old bindered studies from when cost was no object (until someone actually tried to spend the money).
We've spend a lot of breath on the economic of reusability, and a good deal of it was conceptualizing missions to take advantage of it....
This is like talking about moving the house across the street to this side of the street so you don't have to cross the street to get to it.
A Star-48 is proven, relatively cheap, with only minor well-understood integration considerations. It is a no-brainer way for SpaceX to be able to provide high-energy escape trajectories for the small number of payloads that require it. Successful booster recovery and reuse doesn't throw any "wrench" into anything. It makes any mission cheaper. It doesn't mean it makes it sensible to throw multiple launches and all of the complexity that would entail at something that a $6 million kick stage solves.
There are other concepts that could greatly benefit from something like the proposed architecture, but this isn't one of them.
If WE are thinking about FH upper stage, the next probabal step is a single raptor.
If WE are thinking about FH upper stage, the next probabal step is a single raptor.
I've long wondered if any performance advantage can be gained from a methane-fuelled Merlin upper stage.
My first stab at a real Falcon Heavy page, after holding out for a while.A proposed typology in a few letters:
www.spacelaunchreport.com/falconH.html (http://www.spacelaunchreport.com/falconH.html)
Comments and criticisms welcomed, because I expect to edit this page substantially this year. I expect to be surprised when the thing rolls out and we actually see it for the first time. I also expect to be surprised by how SpaceX actually uses the machine.
- Ed Kyle
Crossfeed has been put on the back burner and I wouldn't cover barge for boosters. I don't think that will be needed, the difference between pad and barge would be in the overall noise.While crossfeed may not appear for awhile, if ever, I'll have to include it as a future possibility as long as SpaceX continues to talk about 53 tonnes to LEO and 21.2 tonnes to GTO, since those are crossfeed (fully expendable) numbers.
Unless some phase of flight is extremely ill-posed from a TWR perspective, going from WPB to WPE should *not* double payload capacity, it should be a much lower figure. It's only with WPP that you get really severe changes, due to the extreme boostback requirements. The upper stage is heavy enough that the residual fuel needed to land the center core on a barge is not very extreme.Crossfeed has been put on the back burner and I wouldn't cover barge for boosters. I don't think that will be needed, the difference between pad and barge would be in the overall noise.While crossfeed may not appear for awhile, if ever, I'll have to include it as a future possibility as long as SpaceX continues to talk about 53 tonnes to LEO and 21.2 tonnes to GTO, since those are crossfeed (fully expendable) numbers.
My guess is that the booster flyback recovery will usually be attempted, but that core recovery attempts (even downrange) may be less frequent because making the core expendable seems to double beyond LEO payload capacity, but I'm just guessing.
- Ed Kyle
How would prop densification and thrust increase improve FH GEO payload with RTLS ?
Falcon Heavy was announced in 2011-04, with Elon quoting 53 tons to LEO. Are you telling us this was in the plans since almost 4 years ago ?How would prop densification and thrust increase improve FH GEO payload with RTLS ?
I strongly suspect that the FH performance figures already take this thrust upgrade into account.
Ed's article implies an (uprated?) Merlin 1D in the offing, yet I thought the 1D has been flying for some time now in its "final" form. Is there in fact an uprated version of the 1D yet to fly?
I've been looking at Falcon Heavy possibilities a bit. The main puzzle is, and has long been, the SpaceX statement that the Heavy GLOW is 1,463 tonnes, while Falcon 9 v1.1 GLOW is only 506 tonnes. If you put Falcon 9 v1.1 pieces together as a first guess to build a Heavy, you only get a GLOW of 1358 tonnes or so.
But even that rocket can get 45 tonnes to LEO and more than 15 tonnes to GTO in expendable mode (no crossfeed is assumed for any of my figuring right now). If the boosters and cores are recovered, and if 35 tonnes of recovery propellant is assumed for each, the numbers drop to 26 tonnes LEO and a bit more than 7 tonnes GTO, the latter of which lines up well with the 6.4 tonnes advertised capability. Expending the core but recovering the boosters gives 34 tonnes LEO and 11 tonnes GTO.
That missing 100 tonnes GLOW could be packed into the rocket in several ways. One guess was that the boosters would be stretched a bit to carry more propellant than the core stage, which works out to improve the beyond-LEO expendable numbers a bit (GTO goes up to 17 tonnes full expendable, but is still only 8 tonnes or so for full recovery). Propellant densification could increase the mass of all three cores with similar results.
So there is my thinking at present, all in flux.
Now, about prices and costs, note that the current going rate to GTO is about $20 million per tonne. Thus, SpaceX would seem to be giving up $160 million of potential income when it recovers all three cores rather than expending them. Do three core stages cost more than $160 million?
- Ed Kyle
Out of curiosity, what have you included as variables in your simulations?Constraints are: GLOW < 1463 tonnes, Boosters = Stage 1 (PMF = 0.95), Stage 2 propellant mass = Stage 1 propellant mass/4.45 (PMF = 0.94), T/W Stage 2 > 0.55 or so, Core propellant at booster burnout = recovery propellant if any + 20% of liftoff load, 0.5% residuals, 9,200 m/s for LEO, 11,700 m/s for GTO. My variables are the stage gross masses.
I've assumed current thrust levels, but it doesn't matter much because there is plenty of thrust even for the heavier SpaceX GLOW. Stage 2 propellant loading could explain some of the 100 tonnes, but not, I think, all of it. The stage "wants" to gross about 100 tonnes or less to work well beyond-LEO, and it can offload propellant from its GTO type loadings for LEO missions.
Ed, is that's with Merlin-1D at 85% or 100%?
There has also talk that S2 has been flying with propellant offload on F9 1.1.. Could that be part of the mass delta?
That missing 100 tonnes GLOW could be packed into the rocket in several ways. One guess was that the boosters would be stretched a bit to carry more propellant than the core stage, which works out to improve the beyond-LEO expendable numbers a bit (GTO goes up to 17 tonnes full expendable, but is still only 8 tonnes or so for full recovery). Propellant densification could increase the mass of all three cores with similar results.
Seems like no one thing accounts for the full 100 tons, but if you combine all three (propellant densification, booster stretch, increased S2 loading) you might get pretty close?
Seems like no one thing accounts for the full 100 tons, but if you combine all three (propellant densification, booster stretch, increased S2 loading) you might get pretty close?
The S2 not launching full on F9 is just forum speculation, and non-sensical forum speculation at that. Don't believe it.
One guess was that the boosters would be stretched a bit to carry more propellant than the core stage
One guess was that the boosters would be stretched a bit to carry more propellant than the core stage
If that's without cross-feed and no reusability or disposable boosters and reusable core then you're really going to have to have the boosters running at a considerably higher throttle than the core. Otherwise the core runs out of prop before the boosters do. That's pretty problematic! :o
One guess was that the boosters would be stretched a bit to carry more propellant than the core stage
If that's without cross-feed and no reusability or disposable boosters and reusable core then you're really going to have to have the boosters running at a considerably higher throttle than the core. Otherwise the core runs out of prop before the boosters do. That's pretty problematic! :o
Why is that problematic? It is the intent to run the core at reduced thrust until booster staging. (it throttles down just after liftoff) Delta IV Heavy operates the same way. M1D has enough of a throttle range to still make that possible with slightly stretched boosters.
One guess was that the boosters would be stretched a bit to carry more propellant than the core stage
If that's without cross-feed and no reusability or disposable boosters and reusable core then you're really going to have to have the boosters running at a considerably higher throttle than the core. Otherwise the core runs out of prop before the boosters do. That's pretty problematic! :o
Why is that problematic? It is the intent to run the core at reduced thrust until booster staging. (it throttles down just after liftoff) Delta IV Heavy operates the same way. M1D has enough of a throttle range to still make that possible with slightly stretched boosters.
One guess was that the boosters would be stretched a bit to carry more propellant than the core stage
If that's without cross-feed and no reusability or disposable boosters and reusable core then you're really going to have to have the boosters running at a considerably higher throttle than the core. Otherwise the core runs out of prop before the boosters do. That's pretty problematic! :o
Why is that problematic? It is the intent to run the core at reduced thrust until booster staging. (it throttles down just after liftoff) Delta IV Heavy operates the same way. M1D has enough of a throttle range to still make that possible with slightly stretched boosters.
One guess was that the boosters would be stretched a bit to carry more propellant than the core stage
If that's without cross-feed and no reusability or disposable boosters and reusable core then you're really going to have to have the boosters running at a considerably higher throttle than the core. Otherwise the core runs out of prop before the boosters do. That's pretty problematic! :o
Why is that problematic? It is the intent to run the core at reduced thrust until booster staging. (it throttles down just after liftoff) Delta IV Heavy operates the same way. M1D has enough of a throttle range to still make that possible with slightly stretched boosters.
You surely have not read carefully. Go back and re-read please!!! (sigh)
Why is that problematic? It is the intent to run the core at reduced thrust until booster staging. (it throttles down just after liftoff) Delta IV Heavy operates the same way. M1D has enough of a throttle range to still make that possible with slightly stretched boosters.
Regardless, the boosters on the Falcon Heavy are stretched, end of story. But in your defense, I guess the Falcon Heavy boosters (I believe?) did have a slight run-in with an overpass or something on their way to Texas. :DWhy is that problematic? It is the intent to run the core at reduced thrust until booster staging. (it throttles down just after liftoff) Delta IV Heavy operates the same way. M1D has enough of a throttle range to still make that possible with slightly stretched boosters.
From what I've heard, the boosters have already been stretched to the maximum extent possible. Any longer and they would not be road transportable.
From what I've heard, the boosters have already been stretched to the maximum extent possible. Any longer and they would not be road transportable.The *Falcon 9* has been stretched as far as possible (as far as we know). It's not road-transport which is the issue (that's the constraint on the *diameter* of the stage), it's the bending modes. A wobbly thin rocket is tricky to push from the bottom. (See Fineness ratio (http://en.m.wikipedia.org/wiki/Fineness_ratio) on the wiki.). But when you take the second stage and payload off the boosters, you apparently gain the ability to do some extra stretch on them. Which makes sense, right? Naively you'd think you could extend the boosters to at least the size of S1+S2 on the core, and the amount of stretch depicted is much more modest (roughly the size of the interstage).
Regardless, the boosters on the Falcon Heavy are stretched, end of story. But in your defense, I guess the Falcon Heavy boosters (I believe?) did have a slight run-in with an overpass or something on their way to Texas. :DUnfortunately, it isn't the end of the story. The big diagrams of Falcon Heavy on the SpaceX web site show what might be stretched boosters, but the recent video release showed boosters that looked the same length as the core with the extra space above the tanks used to support grid fins, etc. Also, I seem to remember SpaceX officials stating that the core and boosters would be identical, or nearly so.
I look at it this way.
The interstage is basically part of the first stage. Since there is no need for the interstage on the side boosters they might as well extend the tanks that far.
Or....do they need something akin to the interstage for the side booster grid fins and avionics?
Ship it in two sections, one for the RP and one for the LOX tank. Weight wise it costs you one additional bulkhead since you wouldn't have a common bulkhead anymore but you can join the two sections similar to how you would join stage one and stage two. Probably a few more bolts so everything stays together for landing.
I would give you 20:1 odds and still take the bet. The boosters are stretched, we have it from multiple sources.Who?
SpaceX multiple times and Jim.I would give you 20:1 odds and still take the bet. The boosters are stretched, we have it from multiple sources.Who?
- Ed Kyle
Maybe all three are stretched from v1.1. That would make sense. Boosters carrying more propellant than core makes no sense unless crossfeed is used, which is apparently not going to be the case. There is no performance advantage. If anything, there is a performance penalty.SpaceX multiple times and Jim.I would give you 20:1 odds and still take the bet. The boosters are stretched, we have it from multiple sources.Who?
- Ed Kyle
If you were sufficiently brave (or perhaps foolhardy) you could use active control of bending modes. The Falcon should be in a particularly good position to do this, if it's willing to use the grid fins on the way up.From what I've heard, the boosters have already been stretched to the maximum extent possible. Any longer and they would not be road transportable.The *Falcon 9* has been stretched as far as possible (as far as we know). It's not road-transport which is the issue (that's the constraint on the *diameter* of the stage), it's the bending modes. A wobbly thin rocket is tricky to push from the bottom.
Boosters carrying more propellant than core makes no sense unless crossfeed is used, which is apparently not going to be the case. There is no performance advantage. If anything, there is a performance penalty.
Boosters carrying more propellant than core makes no sense unless crossfeed is used, which is apparently not going to be the case. There is no performance advantage. If anything, there is a performance penalty.
I am not an expert, so could you elaborate more on why this would not provide any performance advantage? Center core will be throttled down = it feels like the more propelant you have in boosters, the higher delta V you can give to the center core.
The comparison I made is for a given GLOW, assumes no crossfeed, and assumes full recovery of boosters and core. The choice is between all three cores being identical and having the boosters carry, say, 10% more propellant than the core. In my modeling, I find that the identical cores example provides slightly more LEO payload than the bigger boosters example, partly because there is less propellant left in the core at staging for the latter. The goal isn't just to boost the core to some velocity, it is also to stage it with as much propellant still on board as possible. The longer-burning boosters would cause the core to burn more propellant before staging.Boosters carrying more propellant than core makes no sense unless crossfeed is used, which is apparently not going to be the case. There is no performance advantage. If anything, there is a performance penalty.
I am not an expert, so could you elaborate more on why this would not provide any performance advantage? Center core will be throttled down = it feels like the more propelant you have in boosters, the higher delta V you can give to the center core.
The longer-burning boosters would cause the core to burn more propellant before staging.
The interstage, man.Maybe all three are stretched from v1.1. That would make sense. Boosters carrying more propellant than core makes no sense unless crossfeed is used, which is apparently not going to be the case. There is no performance advantage. If anything, there is a performance penalty.SpaceX multiple times and Jim.I would give you 20:1 odds and still take the bet. The boosters are stretched, we have it from multiple sources.Who?
- Ed Kyle
Meanwhile, we have these images to ponder.
- Ed Kyle
The longer-burning boosters would cause the core to burn more propellant before staging.Can't the core be throttled even more? Or some core engines may even be shut down to save fuel? But I think I see your point - adding more fuel to boosters makes them carry their own weight for a longer period and the goal is to get rid of the boosters as soon as possible.
The core can be throttled to the minimum for both cases, but will still have less propellant left for the longer-burning booster example. The only advantage of a longer booster would be if it transferred propellant to the core while still attached, but crossfeed seems unlikely to appear for awhile if ever.The longer-burning boosters would cause the core to burn more propellant before staging.
Can't the core be throttled even more? Or some core engines may even be shut down to save fuel? But I think I see your point - adding more fuel to boosters makes them carry their own weight for a longer period and the goal is to get rid of the boosters as soon as possible.
I would give you 20:1 odds and still take the bet. The boosters are stretched, we have it from multiple sources.
The first mission is really a demonstration flight. It's there to prove that Falcon Heavy will work. That it will deliver the payload that we say it can, and we don't have a primary customer for it, but we are likely to have several smaller secondary satellites on-board that will do a variety of things, and if we get lucky, maybe there will be a big satellite at the last minute that wants to buy the flight at a reduced price.
The snag is USAF/NASA certification, unless they could upgrade with a single launch to confirm everything is ok, SpaceX has a big incentive for doing all much as possible on V1.0.I would give you 20:1 odds and still take the bet. The boosters are stretched, we have it from multiple sources.
I'll take the bet. 20:1 odds are pretty good!
I think unstretched boosters makes sense given SpaceX's incremental development philosophy, plus the fact that FH is already rather late.
FHv1: unstretched boosters, no crossfeed, identical tanks, interstage & grid fins for all cores, "just" stage interconnects, core throttling, and everything else which is "essential" for FH. Get 'er flying ASAP!
FHv1.1: stretched boosters, new interstage for boosters, grid fin relocation, etc. Probably all in development, just not on the critical path for first flight anymore. Low-hanging fruit.
FHv1.2: crossfeed. Developed only if/when a customer has need, may be bypassed by current events.
...
FHv2: S2 reusability? (A man can dream!)
That seems like a nice incremental development path, assuming that all three of FHv1, FHv1.1, FHv1.2 offer monotonic performance improvements over F9R.
Question: how much will the increased thrust of the Merlin 1D+ affect the benefits of crossfeed? If the side boosters have more thrust, you can throttle down the center core more without being eaten up alive by gravity drag, and the benefits of crossfeed should be smaller.Precisely.
Ok, I'm aware this sound insane from a risk standpoint. But what if you not only throttled down the core engines but actually shut down the center engine as well? They've had a lot of experience now re-lighting it, just not while the other ones are running. So you throttle the core mains to say 80% and shut down the center engine while the boosters keep burning at 115%. Now you'd have almost half the fuel left in the core when you stage. That's significant. If the stage can survive this thrust differential at Max Q, and the additional gravity losses...hmm, needs more work...If engines aren't needed, why have them at all? Thrust costs money.
Huh? Where'd you get that from? You'd still want full thrust right at lift-off but throttle down the core (perhaps shut down an engine or two) soon after then throttle the core to full right after the boosters stage off.Ok, I'm aware this sound insane from a risk standpoint. But what if you not only throttled down the core engines but actually shut down the center engine as well? They've had a lot of experience now re-lighting it, just not while the other ones are running. So you throttle the core mains to say 80% and shut down the center engine while the boosters keep burning at 115%. Now you'd have almost half the fuel left in the core when you stage. That's significant. If the stage can survive this thrust differential at Max Q, and the additional gravity losses...hmm, needs more work...If engines aren't needed, why have them at all? Thrust costs money.
It seems more likely that the rocket is designed to use, and need, the thrust.
- Ed Kyle
Ok, I'm aware this sound insane from a risk standpoint. But what if you not only throttled down the core engines but actually shut down the center engine as well? They've had a lot of experience now re-lighting it, just not while the other ones are running. So you throttle the core mains to say 80% and shut down the center engine while the boosters keep burning at 115%. Now you'd have almost half the fuel left in the core when you stage. That's significant. If the stage can survive this thrust differential at Max Q, and the additional gravity losses...hmm, needs more work...If engines aren't needed, why have them at all? Thrust costs money.
It seems more likely that the rocket is designed to use, and need, the thrust.
- Ed Kyle
I'm curious - does Falcon Heavy automatically fall under the Reusable Rockets category?Yeah, anything which needs its full performance. None really exist, yet, except maybe some high energy missions like New Horizons.
Is there any mission type where none of it gets retrieved or re-used?
Yeah, anything which needs its full performance. None really exist, yet, except maybe some high energy missions like New Horizons.
Not if you really, really need the performance (although for high energy, a small kick stage may be cheaper than throwing away 2 boosters).Yeah, anything which needs its full performance. None really exist, yet, except maybe some high energy missions like New Horizons.
But not even the side-boosters get recovered in that case?
I'm curious - does Falcon Heavy automatically fall under the Reusable Rockets category?Suggest you look at "reusable" as a mode or part of a evolving CONOPs at a longer scale.
Yeah, anything which needs its full performance. None really exist, yet, except maybe some high energy missions like New Horizons.Ok, but there are no unawarded new horizons type missions, right ?
I don't think SpaceX will end up doing any fully expendable FH missions.
At a minimum recovery of the side boosters.
With recovery of the side boosters, F9R + FHR can do what, 95% of all missions up to date ?
Yeah, anything which needs its full performance. None really exist, yet, except maybe some high energy missions like New Horizons.Ok, but there are no unawarded new horizons type missions, right ?
Doesn't it make more sense to leave those type missions for Raptor rockets instead ?
Pick the battles you can win, and those you might not win, but some other spoils can come from.
SpaceX has its plate full.
I don't think SpaceX will end up doing any fully expendable FH missions.
At a minimum recovery of the side boosters.
With recovery of the side boosters, F9R + FHR can do what, 95% of all missions up to date ?
Anyhow, what payload class is new horizons ? Is it the same type as JWST ?
Huh? Where'd you get that from? You'd still want full thrust right at lift-off but throttle down the core (perhaps shut down an engine or two) soon after then throttle the core to full right after the boosters stage off.Ok, I'm aware this sound insane from a risk standpoint. But what if you not only throttled down the core engines but actually shut down the center engine as well? They've had a lot of experience now re-lighting it, just not while the other ones are running. So you throttle the core mains to say 80% and shut down the center engine while the boosters keep burning at 115%. Now you'd have almost half the fuel left in the core when you stage. That's significant. If the stage can survive this thrust differential at Max Q, and the additional gravity losses...hmm, needs more work...If engines aren't needed, why have them at all? Thrust costs money.
It seems more likely that the rocket is designed to use, and need, the thrust.
- Ed Kyle
If you shut down all the engines on the second stage (or didn't use them at all for launch) is that the same a making it a three stage rocket. After all the second stage isn't lit until later in flight. Or does the maths simply not work out?
If you shut down all the engines on the second stage (or didn't use them at all for launch) is that the same a making it a three stage rocket. After all the second stage isn't lit until later in flight. Or does the maths simply not work out?
If you do not light center core engines until booster cores' separation, that is the canonical definition of a three stage rocket. If you do light the center core engines before separation it's a 'hybrid' 2.5 stage rocket, for lack of a better term.
EDIT: Really the idea of a 'stage' is a conceptual construct we use to describe a rocket that is much more nuanced and flexible in reality.
Regardless, the boosters on the Falcon Heavy are stretched, end of story. But in your defense, I guess the Falcon Heavy boosters (I believe?) did have a slight run-in with an overpass or something on their way to Texas. :DWhy is that problematic? It is the intent to run the core at reduced thrust until booster staging. (it throttles down just after liftoff) Delta IV Heavy operates the same way. M1D has enough of a throttle range to still make that possible with slightly stretched boosters.
From what I've heard, the boosters have already been stretched to the maximum extent possible. Any longer and they would not be road transportable.
If you shut down all the engines on the second stage (or didn't use them at all for launch) is that the same a making it a three stage rocket. After all the second stage isn't lit until later in flight. Or does the maths simply not work out?
SpaceX may have an increased confidence in restarting engines now... May be shutting down engines on the center core and then starting them again does not sound that dangerous any more (after so many successfull restarts for reentry and landing).
If you shut down all the engines on the second stage (or didn't use them at all for launch) is that the same a making it a three stage rocket. After all the second stage isn't lit until later in flight. Or does the maths simply not work out?
I think you'd need the central core burning to get FH off the pad. But they probably could be shut down shortly into ascent. Since the Merlin can be relit after shutdown, it could stay off until booster sep, and then relight. Gravity losses may be a little more in this ascent profile, but the benefit would be the stack is going pretty low and slow at booster sep so they should be pretty easy to recover at LC-13 without the central core helping very much prior to that point.
Sigh. That's not gonna happen. M1D can throttle to 50% - perhaps even lower - and this is more than enough to leave propellant in the core.
Sigh. That's not gonna happen. M1D can throttle to 50% - perhaps even lower - and this is more than enough to leave propellant in the core.
Sigh. That's not gonna happen. M1D can throttle to 50% - perhaps even lower - and this is more than enough to leave propellant in the core.
I'm curious, do you have a source for this 50% throttle capability?
If you shut down all the engines on the second stage (or didn't use them at all for launch) is that the same a making it a three stage rocket.
I was really thinking of a stack rather than side by side
Maybe all three are stretched from v1.1.
If you shut down all the engines on the second stage (or didn't use them at all for launch) is that the same a making it a three stage rocket.
I thought second stage meant core until I read this:I was really thinking of a stack rather than side by side
Second stage engines cannot be used at launch. They cannot be used until the first stage is jettisoned, therefore you would not shut them down, because you cannot light them while the first stage is running. Think about it, if you light the second stage while the first stage is still firing, the T/W on stage 1 is higher than stage 2; stage 2 cannot pull away from stage 1; the flame from stage 2 would likely ignite any prop still in stage 1. (I know, there are rare exceptions.) Stage 1 has MECO (main engine cut off). Stage 1 jettisons. Interstage adaptor jettisons (on most rockets, but not this one). Stage 2 ignites. This is very basic..
It's not road-transport which is the issue (that's the constraint on the *diameter* of the stage), it's the bending modes.Road transport constraints involve both core diameter and core length.
The boosters are stretched, we have it from multiple sources.I agree. But again, if you include the interstage as an integral part the center core, the length of all 3 cores is pretty close. In fact, if you remove the domes, the booster cores would be shorter.
IIRC, the cross-feed scheme would have four core engines fuelled from the portside booster and four others from the starboard booster with one core engine fuelled from the core tanks. So, a larger core tank would seem to make sense.The one diagram I recall showed cross feed between distribution nodes that feed all engines on a particular core. So the center core has a node that is fed by three tanks and distributing to all nine engines. That diagram was very cartoonish but it also seems more intuitive. Either way I don't see the relationship to tank size. Same size core and boosters or not ... cross feed or not ... lots of choices to spend money (or not) for performance options.
I agree. But again, if you include the interstage as an integral part the center core, the length of all 3 cores is pretty close. In fact, if you remove the domes, the booster cores would be shorter.The interstage protects the upper-stage engine. The core tank ends below it.
I agree. But again, if you include the interstage as an integral part the center core, the length of all 3 cores is pretty close. In fact, if you remove the domes, the booster cores would be shorter.The interstage protects the upper-stage engine. The core tank ends below it.
I think it's likely that the domes of the side boosters are part of the tank structure and filled with propellant. For example the space shuttle external tank's nose was filled with LOX.
IIRC, the cross-feed scheme would have four core engines fuelled from the portside booster and four others from the starboard booster with one core engine fuelled from the core tanks.I thought it was 3 for each booster and 3 for the center core, but I'm not sure. Anyone have a reference?
So, a larger core tank would seem to make sense.if they could get a larger core from Hawthorne to the launch site.
So, a larger core tank would seem to make sense.if they could get a larger core from Hawthorne to the launch site.
IIRC, the cross-feed scheme would have four core engines fuelled from the portside booster and four others from the starboard booster with one core engine fuelled from the core tanks.I thought it was 3 for each booster and 3 for the center core, but I'm not sure. Anyone have a reference?So, a larger core tank would seem to make sense.if they could get a larger core from Hawthorne to the launch site.
Maybe all three are stretched from v1.1.
From the SpaceX website:
http://www.spacex.com/falcon9
Height 68.4 m 224.4 ft
http://www.spacex.com/falcon-heavy
Height 68.4 m 224.4 ft
If they can throttle the new upgraded engines, cross feed might not be necessary. Throttle down the core and throttle up the boosters. The second stage engine has/is being upgraded also, and it seems to be a much more powerful upgrade.
Nobody said that 120' was an absolute limit for road transport. Just that the number of possible routes drops off a cliff once you pass 120'. It doesn't matter whether you can transport 220' towers *somewhere*. The issues is the maximum load length possible on coast-to-coast routes between Hawthorne, McGregor, and the Cape.IIRC, the cross-feed scheme would have four core engines fuelled from the portside booster and four others from the starboard booster with one core engine fuelled from the core tanks.I thought it was 3 for each booster and 3 for the center core, but I'm not sure. Anyone have a reference?So, a larger core tank would seem to make sense.if they could get a larger core from Hawthorne to the launch site.
I've seen wind generator towers transported on two-lane roads as one unit. They are typically 212 feet in length.
https://www.wind-watch.org/faq-size.php (https://www.wind-watch.org/faq-size.php)
The extended side booster of the Falcon Heavy doesn't look like it's much longer than 150 feet, so the length concerns posted here are simply not realistic.
Wait, is it quite true, though, that 120' is such a hard (okay, "firm") limit? I got the idea that length has a much less firm limit than height, which obviously over any distance greater than a few miles has serious issues due to bridges. I mean, once you're on the Interstate, it's not like there are that many sharp turns.Turns are probably less of a big deal than height problems. Slap a 300' cargo a lot of different places in the interstate system (we call them 'hills') and your wheels will be off the ground while the center of the cargo see-saws. Raise this cargo high up enough that it can survive a gentle hill, and you start to scrape overhead bridges. Put in dynamic shocks so you can raise or lower it, and even then you can't deal with a deliberate depression under a bridge.
Wait, is it quite true, though, that 120' is such a hard (okay, "firm") limit? I got the idea that length has a much less firm limit than height, which obviously over any distance greater than a few miles has serious issues due to bridges. I mean, once you're on the Interstate, it's not like there are that many sharp turns.
How often would that be necessary?Wait, is it quite true, though, that 120' is such a hard (okay, "firm") limit? I got the idea that length has a much less firm limit than height, which obviously over any distance greater than a few miles has serious issues due to bridges. I mean, once you're on the Interstate, it's not like there are that many sharp turns.
On-ramps and off-ramps, however, present all manner of pain-in-the-ass problems for transporting long objects.
I'm not buying it. Look up Tonopah, NV,and U.S. 95. Small town, mountainous terrain. That's where I've seen those 212' long towers being pulled. Unless there are major obstacles right outside the front gates of either Hawthorne, McGregor (which is a pretty wide open space, so not likely there) or Cape Canaveral, I doubt there will be any on the interstate route in between.
Something to remember is that there are different levels of oversized. There is oversized with a permit that can travel night and day and does not require escort vehicles. Then there is the kind that DOES require escort and can only travel certain hours. I don't know the details, but there must be a huge difference in cost and time involved.
If you add the u/s stretch, presumably both of those numbers become obsolete. The FH one possibly before the first time it flies.
It doesn't matter whether you can transport 220' towers *somewhere*. The issues is the maximum load length possible on coast-to-coast routes between Hawthorne, McGregor, and the Cape.and to the new Boca Chica launch site.
I'm not buying it. Look up Tonopah, NV,and U.S. 95. Small town, mountainous terrain. That's where I've seen those 212' long towers being pulled...
How are the wind turbine components transported?
Transport of such large items and the cranes needed to assemble them often presents problems in the remote areas where they are typically built. Roads must be widened, curves straightened, and in wild areas new roads built altogether.
I'm not buying it. Look up Tonopah, NV,and U.S. 95. Small town, mountainous terrain. That's where I've seen those 212' long towers being pulled. Unless there are major obstacles right outside the front gates of either Hawthorne, McGregor (which is a pretty wide open space, so not likely there) or Cape Canaveral, I doubt there will be any on the interstate route in between.
How often would that be necessary?Wait, is it quite true, though, that 120' is such a hard (okay, "firm") limit? I got the idea that length has a much less firm limit than height, which obviously over any distance greater than a few miles has serious issues due to bridges. I mean, once you're on the Interstate, it's not like there are that many sharp turns.
On-ramps and off-ramps, however, present all manner of pain-in-the-ass problems for transporting long objects.
Anyway, this is all a good argument for Return to Launch Site.
All I see in that post is a clear indication of a stage becoming longer with time.If you add the u/s stretch, presumably both of those numbers become obsolete. The FH one possibly before the first time it flies.
Is a second stage stretch confirmed?
Some have speculated that the added 10% volume can be done without a stretch.
For example:
http://forum.nasaspaceflight.com/index.php?topic=36815.msg1340204#msg1340204
All I see in that post is a clear indication of a stage becoming longer with time.If you add the u/s stretch, presumably both of those numbers become obsolete. The FH one possibly before the first time it flies.
Is a second stage stretch confirmed?
Some have speculated that the added 10% volume can be done without a stretch.
For example:
http://forum.nasaspaceflight.com/index.php?topic=36815.msg1340204#msg1340204
+10% vol means stretched, period.
Actually, upper stage tank volume +10% does not necessarily mean the outer structure needs to be stretched. A good example is the good old Ariane 4 - its LH2 third stage was modified twice in history (first flight 1988, first modification in 1992 and the second in 1994) to allow for slight increases in the fuel capacity with just internal tank stretches of inches. Each stretch brought about 100-150 kg increase in GTO payload capacity.
Yes, but the graphic shows that the first two stretches did, in fact, grow the overall length of the stage.
Only the final stretch maintained the same stage length, by tweaking the tanks and skirt lengths by a few centimeters.
The scale of the tank growth anticipated on F9 is more consistent with the first two stretches shown in the diagram, both of which increased stage length.
No confirmation that I've seen.
No confirmation that I've seen.
You mean confirmation besides Elon Musk saying it in a Twitter?
I'm not buying it. Look up Tonopah, NV,and U.S. 95. Small town, mountainous terrain. That's where I've seen those 212' long towers being pulled...
From the link you provided earlier:
https://www.wind-watch.org/faq-size.php (https://www.wind-watch.org/faq-size.php)QuoteHow are the wind turbine components transported?
Transport of such large items and the cranes needed to assemble them often presents problems in the remote areas where they are typically built. Roads must be widened, curves straightened, and in wild areas new roads built altogether.
So it appears that, while there's no fixed limit, longer stages would tend to present more issues with road transport.
In any case, an informative discussion.
Sigh. Yes, we know Elon has tweeted that tank volume will increase by 10%. That's why we're having this discussion.
The question is whether that will result in a length increase in the overall stage. I believe it will. Others believe it will not. But no one from SpaceX has confirmed what the actual effect on overall stage length, if any, will be.
How long are the second stage tanks, without engine, just the tanks? That gives how much in cm a 10% stretch is.About 9 meters, consistently with 90 t propellant load.
Everything can be imagined or speculated: bulges, external auxiliary tanks, modified pressurization with removal of He COPV tanks etc...A different meaning that could, hypothetically, be possible:
But a one meter stretch of the stage is so simple that "+10% vol" can't mean anything different.
How long are the second stage tanks, without engine, just the tanks? That gives how much in cm a 10% stretch is.About 9 meters, consistently with 90 t propellant load.
A different meaning that could, hypothetically, be possible:
Chilled LOX gets more "densification gains" than chilled RP-1 (~9% vs. ~5%, respectively), thereby altering the balance of prop:LOX to excess LOX when tanks are full. Additionally, increased thrust M1d may also have optimum performance with an altered fuel mixture (slightly less LOX). -----> Between higher density and lower fuel mix, less LOX volume is needed. Current tank sizes are no longer optimized and moving the bulkhead to increase RP-1 tank volume (at the expense of LOX tank volume) re-establishes optimum. +10% vol (in RP-1 tank) without stage stretch.
Sigh. Yes, we know Elon has tweeted that tank volume will increase by 10%. That's why we're having this discussion.It has to increase the length. The stage uses a common bulkhead and is cylindrical. I doubt there is any room in the interstage for a 10% volume increase. More thrust needs more propellant to achieve best results, so this stretch seems in line with the up-thrusted engines and with the needs of Falcon Heavy.
The question is whether that will result in a length increase in the overall stage. I believe it will. Others believe it might not. But no one from SpaceX has confirmed what the actual effect on overall stage length, if any, will be.
v1.0 was constrained by width, not length. Or at least there was no indication length was a concern, only width.Something to remember is that there are different levels of oversized. There is oversized with a permit that can travel night and day and does not require escort vehicles. Then there is the kind that DOES require escort and can only travel certain hours. I don't know the details, but there must be a huge difference in cost and time involved.
F9 v1.0 was sized at the maximum to be transported with minimum fuss. By implication, the constraints on v1.1 transport must be greater.
Cheers, Martin
v1.0 was constrained by width, not length. Or at least there was no indication length was a concern, only width.Something to remember is that there are different levels of oversized. There is oversized with a permit that can travel night and day and does not require escort vehicles. Then there is the kind that DOES require escort and can only travel certain hours. I don't know the details, but there must be a huge difference in cost and time involved.
F9 v1.0 was sized at the maximum to be transported with minimum fuss. By implication, the constraints on v1.1 transport must be greater.
Cheers, Martin
Sigh. Yes, we know Elon has tweeted that tank volume will increase by 10%. That's why we're having this discussion.It has to increase the length. The stage uses a common bulkhead and is cylindrical. I doubt there is any room in the interstage for a 10% volume increase. More thrust needs more propellant to achieve best results, so this stretch seems in line with the up-thrusted engines and with the needs of Falcon Heavy.
The question is whether that will result in a length increase in the overall stage. I believe it will. Others believe it might not. But no one from SpaceX has confirmed what the actual effect on overall stage length, if any, will be.
- Ed Kyle
More thrust needs more propellant to achieve best results...
hammerhead? the second stage is much smaller and easier to transport, so a small width increase might not cause too many problems.
A hammerhead pressure vessel, while not impossible, is relatively novel and incompatible with present production methods - it represents a greater technical risk and greater amount of development than SpaceX has taken with nearly any aspect of their business. A silly double-F9 upper, with two welded cylinders each holding independent pressure vessels side by side, is more likely than a hammerhead tank, but without an extra orbital stage doubling the dry mass removes most of the benefit.Sigh. Yes, we know Elon has tweeted that tank volume will increase by 10%. That's why we're having this discussion.It has to increase the length. The stage uses a common bulkhead and is cylindrical. I doubt there is any room in the interstage for a 10% volume increase. More thrust needs more propellant to achieve best results, so this stretch seems in line with the up-thrusted engines and with the needs of Falcon Heavy.
The question is whether that will result in a length increase in the overall stage. I believe it will. Others believe it might not. But no one from SpaceX has confirmed what the actual effect on overall stage length, if any, will be.
- Ed Kyle
hammerhead? the second stage is much smaller and easier to transport, so a small width increase might not cause too many problems.
SpaceX has a very big Friction Stir Welding machine optimized for a given diameter.
This FSW machine can likely assemble cylindrical sections of any lenght up to the maximum sheet lenght available.
Seems to me that KISS goes for stretching.
Note that the first and second stages use a common architecture such as the same 3.7 meter (12 foot) diameter aluminum-lithium barrels and domes, and we manufacture them utilizing the same systems and tooling. This approach greatly reduces overhead, inventory and production costs, and simultaneously contributes to increased reliability. These are essential aspects of how SpaceX improves reliability and lowers the cost of access to space.
A hammerhead seems a lot less likely than a simple stretch of the barrel. Musk mentioned this machine and its capabilities as a competitive advantage of SpaceX... why make a hammerhead if it makes things more complex when a simple stretch of maybe 90cm will do?
A hammerhead seems a lot less likely than a simple stretch of the barrel. Musk mentioned this machine and its capabilities as a competitive advantage of SpaceX... why make a hammerhead if it makes things more complex when a simple stretch of maybe 90cm will do?
Yep...
I must nevertheless point out that pictures you posted are of their old circumferential FSW.
Now SpaceX has a longitudinal FSW that can be seen in the circled area.
Joining of barrels is likely done by electric arc welding in the so called "Paint Booth".
SpaceX used a Garmat USA spray booth that was designed, manufactured, and installed through RelyOn Technologies. This spray booth system was design to finish the outer aesthetics and protective layers of their Falcon 9 v1.1 Rocket, which is also known as the Grasshopper. After receiving an initial concept of what the rocket scientists at SpaceX needed, RelyOn was able design a paint booth that was specific to their needs. These types of rockets require external coatings that are equipped for harsh atmospheres.
From http://shitelonsays.com/transcript/spacex-dragon-2-unveil-qa-2014-05-29, talking about F9 v1.1:v1.0 was constrained by width, not length. Or at least there was no indication length was a concern, only width.And nowhere other than speculation here have I heard that 1.1 or the FH boosters are any different.
It's a nice shot, it's a tall rocket. It's as skinny as I thought we could possibly make it. We stretched to it as long as .. [is that for road transport?] Yeah. It's twelve feet in diameter and when you add little bits and pieces, it gets out to almost 14 feet, and then we have to tuck the little bits and pieces in the corners because the key thing is that the total height above the road has got to be less than 14.5 feet. [Question about bigger rockets.] Need a boat. Or they've got to fly themselves. Not going to fit on the roads, that's for sure.
From http://shitelonsays.com/transcript/spacex-dragon-2-unveil-qa-2014-05-29 (http://shitelonsays.com/transcript/spacex-dragon-2-unveil-qa-2014-05-29), talking about F9 v1.1:v1.0 was constrained by width, not length. Or at least there was no indication length was a concern, only width.And nowhere other than speculation here have I heard that 1.1 or the FH boosters are any different.QuoteIt's a nice shot, it's a tall rocket. It's as skinny as I thought we could possibly make it. We stretched to it as long as .. [is that for road transport?] Yeah. It's twelve feet in diameter and when you add little bits and pieces, it gets out to almost 14 feet, and then we have to tuck the little bits and pieces in the corners because the key thing is that the total height above the road has got to be less than 14.5 feet. [Question about bigger rockets.] Need a boat. Or they've got to fly themselves. Not going to fit on the roads, that's for sure.
But Falcon 9 can't burn its propellant at very high G. It is already G-limited on purpose to give the payload a soft ride, requiring those fancy Merlins to throttle down after the early phases of flight. Thrust costs money, and usually adds weight, so it is wasted on a rocket like this unless it allows more propellant to be carried, which allows more payload (or frees up more propellant for recovery burns).More thrust needs more propellant to achieve best results...
No it doesn't. You can burn the same volume of prop. at a faster rate and thus have a lesser metric of time x gravity losses. You may have some issues @ MaxQ and MaxDrag, but at other parts of the profile you endure higher G forces in order to induce lower gravity losses. Heck, the MX Peacekeeper missile burns for what, between 1 and 2 minutes before burning out @ almost orbital velocity. Burning the prop in such a short time at very high G means that far more impulse goes into true delta V and much less gravity losses.
Looks to me like he's talking about width as the hard and fast limit.
Anything above about 13.5' greatly reduces the routes you can take. We tried for weeks to plot a route for a 14' 3" trailer from Edison, New Jersey to Portland, Maine and never did find a way. Interstate bridges don't always meet interstate standards.Looks to me like he's talking about width as the hard and fast limit.
No. Width can be much more. Up to 33 feet with special permits.
Height is the hard limit. 14.5 feet is the max to get under bridges, power lines, etc.
Length is also limiting, but that's not a hard limit.
Some agreement could be therapeutic!
Looks to me like he's talking about width as the hard and fast limit.
No. Width can be much more. Up to 33 feet with special permits.
Height is the hard limit. 14.5 feet is the max to get under bridges, power lines, etc.
Length is also limiting, but that's not a hard limit.
Elliptical cross section?.....maybe just for trucking.
Looks to me like he's talking about width as the hard and fast limit.
No. Width can be much more. Up to 33 feet with special permits.
Height is the hard limit. 14.5 feet is the max to get under bridges, power lines, etc.
Length is also limiting, but that's not a hard limit.
Everything can be imagined or speculated: bulges, external auxiliary tanks, modified pressurization with removal of He COPV tanks etc...A different meaning that could, hypothetically, be possible:
But a one meter stretch of the stage is so simple that "+10% vol" can't mean anything different.
Chilled LOX gets more "densification gains" than chilled RP-1 (~9% vs. ~5%, respectively), thereby altering the balance of prop:LOX to excess LOX when tanks are full. Additionally, increased thrust M1d may also have optimum performance with an altered fuel mixture (slightly less LOX). -----> Between higher density and lower fuel mix, less LOX volume is needed. Current tank sizes are no longer optimized and moving the bulkhead to increase RP-1 tank volume (at the expense of LOX tank volume) re-establishes optimum. +10% vol (in RP-1 tank) without stage stretch.
The problem is that Elon announced this all via twitter which really doesn't allow much flexibility for complex/nuanced explanations. Personally, I think that there is a small stretch, but the above hypothesis could also fit with the currently available public information.
FH Trajectory | M1D and current US: Max Payload | M1D+ and 10%> US: Max Payload |
LEO | 53 | 58.4 |
TMI | 13.2 | 14.62 |
For FH the increased thrust M1D+ and 10% tank stretch for US implies the following FH payload capabilities increase is possible based on maintaining a constant PMF (propellant mass fraction) for the US by increasing the payload weight and increase the US dry weight for the tank stretch:This assumes that the long-stated 53 tonne goal was achievable using the current length stage. That is something I've always doubted. It could be that the stretch is needed to achieve, or approach, that goal.
FH Trajectory M1D and current US: Max Payload M1D+ and 10%> US: Max Payload LEO 53 58.4 TMI 13.2 14.62
The stretch is in order to come near the performance goal, without the need to develop cross-feed. Cross-feed adds complexity and actually is considered by SpaceX engineers to be an obstacle to rapid first stage reusability.For FH the increased thrust M1D+ and 10% tank stretch for US implies the following FH payload capabilities increase is possible based on maintaining a constant PMF (propellant mass fraction) for the US by increasing the payload weight and increase the US dry weight for the tank stretch:This assumes that the long-stated 53 tonne goal was achievable using the current length stage. That is something I've always doubted. It could be that the stretch is needed to achieve, or approach, that goal.
FH Trajectory M1D and current US: Max Payload M1D+ and 10%> US: Max Payload LEO 53 58.4 TMI 13.2 14.62
- Ed Kyle
This assumes that the long-stated 53 tonne goal was achievable using the current length stage. That is something I've always doubted. It could be that the stretch is needed to achieve, or approach, that goal.
The stretch is in order to come near the performance goal, without the need to develop cross-feed. Cross-feed adds complexity and actually is considered by SpaceX engineers to be an obstacle to rapid first stage reusability.
Shotwell: Falcon Heavy side booster will be the same as the enhanced F9 v1.1 to be introduced later this year.
https://twitter.com/jeff_foust/status/577878260949864448QuoteShotwell: Falcon Heavy side booster will be the same as the enhanced F9 v1.1 to be introduced later this year.
So Gwynne Shotwell confirmed what was read out of the animation video. The side boosters are not longer than the central core any more.
But then we know they'll have uprated Merlin, densified propellant and 10% bigger upper stage, right?https://twitter.com/jeff_foust/status/577878260949864448QuoteShotwell: Falcon Heavy side booster will be the same as the enhanced F9 v1.1 to be introduced later this year.
So Gwynne Shotwell confirmed what was read out of the animation video. The side boosters are not longer than the central core any more.
I agree with you, but in the interests of suppressing a tedious rehashing of this contentious topic, let's note that 140 characters is not *quite* enough to unambiguously confirm your interpretation. They could be "the same as the enhanced F9"... except for a tank stretch.
Let's all hold off on debating this to death and wait until we see pictures of the hardware.
I agree with you, but in the interests of suppressing a tedious rehashing of this contentious topic, let's note that 140 characters is not *quite* enough to unambiguously confirm your interpretation. They could be "the same as the enhanced F9"... except for a tank stretch.But then we know they'll have uprated Merlin, densified propellant and 10% bigger upper stage, right?
Anything above about 13.5' greatly reduces the routes you can take. We tried for weeks to plot a route for a 14' 3" trailer from Edison, New Jersey to Portland, Maine and never did find a way. Interstate bridges don't always meet interstate standards.Looks to me like he's talking about width as the hard and fast limit.
No. Width can be much more. Up to 33 feet with special permits.
Height is the hard limit. 14.5 feet is the max to get under bridges, power lines, etc.
Length is also limiting, but that's not a hard limit.
And there is precedence for this with the Falcon 9, where SpaceX was selling the V1.1 capability far in advance of fielding the V1.1.
And there is precedence for this with the Falcon 9, where SpaceX was selling the V1.1 capability far in advance of fielding the V1.1.
Yeah, that's called lying. Or false advertising. The problem is they were stating it as Falcon 9 performance, not future-enhancement-version performance (v1.1). It was widely believed that that was the performance of the original Falcon 9.
Either it's lying, or really awful systems analysis. Maybe this has been one of the problems with certification.....
No, those numbers were always for a "Block II" F9, it was common knowledge. The Block II morphed into the v1.1. Future capabilities are frequently advertised, because the lead up from order to launch of most payloads is measured in YEARS.Nice try.
No, those numbers were always for a "Block II" F9, it was common knowledge. The Block II morphed into the v1.1. Future capabilities are frequently advertised, because the lead up from order to launch of most payloads is measured in YEARS.Nice try.
I have the original pamphlet they were handing out at the 1st Space Symposium they attended. It has Falcon 9 performance numbers on it. No mention of any Block II. The numbers are off by 30 to 40%.
So where was the actual Falcon 9 1.0 performance ever quoted???
Wouldn't the original pictures of the Falcon Heavy reflect the longer booster(s)? If they knew that was never going to fly they would have showed the enhanced version with longer cores and upper stage.
No, those numbers were always for a "Block II" F9, it was common knowledge. The Block II morphed into the v1.1. Future capabilities are frequently advertised, because the lead up from order to launch of most payloads is measured in YEARS.Nice try.
I have the original pamphlet they were handing out at the 1st Space Symposium they attended. It has Falcon 9 performance numbers on it. No mention of any Block II. The numbers are off by 30 to 40%.
So where was the actual Falcon 9 1.0 performance ever quoted???
Wouldn't the original pictures of the Falcon Heavy reflect the longer booster(s)? If they knew that was never going to fly they would have showed the enhanced version with longer cores and upper stage.
The original F9 did not end up reaching the performance they targeted.
The original F9 did not end up reaching the performance they targeted.
But that's my "issue". Did they not reach the performance because of poor analysis? Weight came in much heavier?
It clearly shows a short Falcon 9, the 3x3 engine configuration, thrust and ISP for the propulsion (1C).
And then the performance numbers are for another rocket that will be flying 5 years later?? I just don't buy it. But that's my opinion.
Which is worth how much? Are you threatening to sue SpaceX? Are you trying to drum up popular opposition? Are you trying to be a whistleblower?
The business relationships, and whatever agreements are reached, between SpaceX and its clients are exactly between them -- not us. Indignation serves you no purpose.
And there is precedence for this with the Falcon 9, where SpaceX was selling the V1.1 capability far in advance of fielding the V1.1.
Yeah, that's called lying. Or false advertising. The problem is they were stating it as Falcon 9 performance, not future-enhancement-version performance (v1.1). It was widely believed that that was the performance of the original Falcon 9.
Either it's lying, or really awful systems analysis. Maybe this has been one of the problems with certification.....
All payloads have ended up in their prescribed orbits, so I guess performance never fell short of promises.
What's giving you acid is that public information wasn't detailed enough. First show me a launch company that lists performance and pricing in a neat little table on their public web site, then we can talk about certification.
So... SpaceX deliberately misled the public by publicizing a slightly better version of F9 v1.0 but INSTEAD giving us a rocket with nearly twice v1.0's performance. I know, I know, it's HORRIBLE, but true!
All payloads have ended up in their prescribed orbits, so I guess performance never fell short of promises.
What's giving you acid is that public information wasn't detailed enough. First show me a launch company that lists performance and pricing in a neat little table on their public web site, then we can talk about certification.
You mean like the Orbcom-OG2 payload? Why bother upgrade to 1.1 if performance never fell short?
I don't want more detail. Doesn't ULA have a mission planner's guide that lists performance? I also never requested ANYTHING about pricing. All prices are negotiable to some extent. I really don't want to talk to you about certification.
After browsing through the forum topics, I'm starting to realize this is a SpaceX forum. It seems like 80 to 90% of all posts are "SpaceX will/can do everything topics" Even the launch coverage has about a 20 to 1 ratio.
It's cute, because I could see the Orbcom red herring coming even as I was typing my post... But figured I might as well give you the benefit of the doubt.
So for kicks - what does Orbcomm have to do with anything? That was an engine-out event, combined with HSF flight safety rules, and absolutely nothing to do with published or actual performance.
Or is it that they also lied because they didn't write on the website that anomalies happen and the chance for success is <100%
Almost all the published figures for v1.0's performance were for block II's performance. They were to do a few block I flights before transitioning to the better block II.So... SpaceX deliberately misled the public by publicizing a slightly better version of F9 v1.0 but INSTEAD giving us a rocket with nearly twice v1.0's performance. I know, I know, it's HORRIBLE, but true!
If it has twice the performance, then initial estimates were off 100%. So what was v1.0 performance capability?
I never claimed v1.1 didn't live up to it's estimates. Just that v1.0 didn't come close.
Almost all the published figures for v1.0's performance were for block II's performance. They were to do a few block I flights before transitioning to the better block II.
Apparently, they said "hell with it" and just did v1.1, which kept all the promises they made before plus a whole lot more. But just because you weren't careful enough to listen to them telling you that they were (in virtually every case) talking about block II performance (and not block I) doesn't mean that they "lied" about F9's performance.
Yeah, at one point you could see the performance here (just has the newer version now... although actually it might be the upgrade-to-v1.1, but regardless): http://elvperf.ksc.nasa.govAlmost all the published figures for v1.0's performance were for block II's performance. They were to do a few block I flights before transitioning to the better block II.
Apparently, they said "hell with it" and just did v1.1, which kept all the promises they made before plus a whole lot more. But just because you weren't careful enough to listen to them telling you that they were (in virtually every case) talking about block II performance (and not block I) doesn't mean that they "lied" about F9's performance.
Thanks for your answer. So was the capability of v1.0 ever published then? Anywhere or anytime?
v1.1 isn't "block II," actually.Almost all the published figures for v1.0's performance were for block II's performance. They were to do a few block I flights before transitioning to the better block II.
Apparently, they said "hell with it" and just did v1.1, which kept all the promises they made before plus a whole lot more. But just because you weren't careful enough to listen to them telling you that they were (in virtually every case) talking about block II performance (and not block I) doesn't mean that they "lied" about F9's performance.
Thanks for your answer. So was the capability of v1.0 ever published then? Anywhere or anytime?
Falcon 9 first flew in June of 2010. When did the first information about v1.1's (Block II) come into public. Was it not until 2012?
Are you suggesting that SpaceX was "lying" just because they decided to give their customers an even more capable rocket?
If you watched the Air Force side of the House hearing today they were lamenting that SpaceX was not having earlier test launches of the Heavy. ...
Couldn't you lot start a splinter thread if you want to talk about all this non-FH stuff?
If you watched the Air Force side of the House hearing today they were lamenting that SpaceX was not having earlier test launches of the Heavy. ...
I thought the second panel at the hearing was quite complimentary of SpaceX today. They're just trying to be realistic in figuring out a schedule of launch vehicle availability for the next 7-8 years.
The Falcon heavy should be shortly after they finish work a the cape. Once the launch pad and facilities are ready, it should be shortly (1-2 months) probably after that. I think that is a bigger hold up than the rocket itself.
The Falcon heavy should be shortly after they finish work a the cape. Once the launch pad and facilities are ready, it should be shortly (1-2 months) probably after that. I think that is a bigger hold up than the rocket itself.
If the rocket was ready they would just do it from Vandenberg, but it's not. The qualification articles still need to be sent to McGregor. A new vehicle, albeit built on existing cores, is not going to just roll onto the pad for launch.
Not that I see LC-39A being completed this year either.
During the hearing yesterday it was mentioned that in order for the Falcon Heavy to be certified to handle all the payloads the Delta IV Heavy currently handles it would need to direct inject 14,500 kilos of payload into GEO, and then endure a coast phase of 3 hours followed by another burn. Gen. Mitchell stated that the stage to do this doesn't currently exist.
If that is the case, what sort of upgrades would be needed to the second stage to be able to do this or is Gen. Mitchell mistaken? The other question would be how many payloads need this capability? What is the longest coast so far they have done? Would gelling of the RP1 and stage power be the issues?
During the hearing yesterday it was mentioned that in order for the Falcon Heavy to be certified to handle all the payloads the Delta IV Heavy currently handles it would need to direct inject 14,500 kilos of payload into GEO, and then endure a coast phase of 3 hours followed by another burn. Gen. Mitchell stated that the stage to do this doesn't currently exist.
If that is the case, what sort of upgrades would be needed to the second stage to be able to do this or is Gen. Mitchell mistaken? The other question would be how many payloads need this capability? What is the longest coast so far they have done? Would gelling of the RP1 and stage power be the issues?
I heard that as well, but think the General is confused. The D4H cannot even do 14mT direct GEO, but can do that to GTO. Can do half of that to GEO, though. The NRO requirements are direct GEO capability. ULA can do it, but SpaceX has not done it, and never even mentions it. Direct GEO is tough and extremely rare, and FH should be capable of that, but has to prove it. SpaceX U/S is not really set up for direct GEO, or high dV missions. Hence the concern of the AF.
Unless there have been classified payloads, almost all payloads are less than 7000kg and go to GTO. But during wartime, they might need a payload operational ASAP, so the NRO says they need to have that option.
During the hearing yesterday it was mentioned that in order for the Falcon Heavy to be certified to handle all the payloads the Delta IV Heavy currently handles it would need to direct inject 14,500 kilos of payload into GEO, and then endure a coast phase of 3 hours followed by another burn. Gen. Mitchell stated that the stage to do this doesn't currently exist.
If that is the case, what sort of upgrades would be needed to the second stage to be able to do this or is Gen. Mitchell mistaken? The other question would be how many payloads need this capability? What is the longest coast so far they have done? Would gelling of the RP1 and stage power be the issues?
I heard that as well, but think the General is confused. The D4H cannot even do 14mT direct GEO, but can do that to GTO. Can do half of that to GEO, though. The NRO requirements are direct GEO capability. ULA can do it, but SpaceX has not done it, and never even mentions it. Direct GEO is tough and extremely rare, and FH should be capable of that, but has to prove it. SpaceX U/S is not really set up for direct GEO, or high dV missions. Hence the concern of the AF.
Unless there have been classified payloads, almost all payloads are less than 7000kg and go to GTO. But during wartime, they might need a payload operational ASAP, so the NRO says they need to have that option.
I have a feeling more birds like NROL-32 (13mT 328foot antenna ELINT) would be up there if they could get more rockets to push them up there...if the cost is lessened by SpaceX FH....Is that the real mission patch? At first I thought it was a joke about the symbolism of expensive rockets burning money. :o
The "all-seeing eye" is often used for intelligence sats. The real question is how it got onto our money. ;)I have a feeling more birds like NROL-32 (13mT 328foot antenna ELINT) would be up there if they could get more rockets to push them up there...if the cost is lessened by SpaceX FH....Is that the real mission patch? At first I thought it was a joke about the symbolism of expensive rockets burning money. :o
The "all-seeing eye" is often used for intelligence sats. The real question is how it got onto our money. ;)The free masons were mostly religious. The eye at the top of the pyramid represents god. The idea is that god can see all 4 sides of the pyramid at once, while humans, from our vantage point, can only see 1 or 2.
I always thought there was a pyramid stuck in His eye, because He's just so big.The "all-seeing eye" is often used for intelligence sats. The real question is how it got onto our money. ;)The free masons were mostly religious. The eye at the top of the pyramid represents god. The idea is that god can see all 4 sides of the pyramid at once, while humans, from our vantage point, can only see 1 or 2.
Not trying to promote religion or anything, just explain what their thinking was behind the eye on our money.
Just a question: Over on the Falcon 9 v.1.2 thread, it was stated that the projected performance figures for that upgrade were above the published Falcon 9 Block-II figures. If this is the case, will this push Falcon Heavy's performance about 53t IMLEO? Additionally, will the improved Merlin-1d performance in any way affect Merlin-1d-VAC's vacuum ISP?Falcon Heavy's 53t performance assumed Falcon v1.1 and cross feeding. As I understand it, all this improvements are done to allow Falcon Heavy's performance targets without cross feeding. BTW, I believe the target would be 6tonnes to GSO (expendable) and 6.5tonnes to GTO with full boosters and core reusability. But that's my estimation. It was generally accepted that v1.1 FH without cross-feeding would be about 45tonnes to LEO or so. Thus, all this improvements might allow for closer to 50tonnes OR around 35 tonnes to LEO with the three cores RTLS. And that's where it gets really interesting. Think about it for a potential LEO station, like Bigelow, you can launch 35tonnes to LEO for about (may be even less) than a Falcon 9 expendable flight. Unproven, but if achieved, it would allow for really cheap stations and supply runs. Something like an LM's Jupiter/Exoliner but with a mass of 35 tonnes could probably take 15tonnes per trip, which is all that's currently required for the ISS USOS :D.
this was posted to twitter by spacex, then removed within minutes?
associated text was something about "F9H wind tunnel model now decorates the factory".
Shows up in my timeline! Very sexy!!
Tweet and FB post have now been deleted. I can't tell for sure but it looks like a larger fairing and 2nd stage. Also confirms that the boosters and core are the same size after all, too.
Jeezes....allowing for foreshortening due to the angle, it almost looks like a 5.2x20m long fairing.
the other shot
the other shot
And anticipating some people now speculating that FH will have 8 engines per core in 3....2...1... Sigh, it is inevitable.
the other shot
And anticipating some people now speculating that FH will have 8 engines per core in 3....2...1... Sigh, it is inevitable.
the other shot
And anticipating some people now speculating that FH will have 8 engines per core in 3....2...1... Sigh, it is inevitable.
and those center core nozzles... ;)
During the hearing yesterday it was mentioned that in order for the Falcon Heavy to be certified to handle all the payloads the Delta IV Heavy currently handles it would need to direct injectLook here: http://forum.nasaspaceflight.com/index.php?topic=30544.1514,500 kilosof payload into GEO, and then endure a coast phase of 3 hours followed by another burn. Gen. Mitchell stated that the stage to do this doesn't currently exist.
If that is the case, what sort of upgrades would be needed to the second stage to be able to do this or is Gen. Mitchell mistaken? The other question would be how many payloads need this capability? What is the longest coast so far they have done? Would gelling of the RP1 and stage power be the issues?
As pointed out below I mistook pounds for kilos, so it is 14,500 pounds to direct inject. The length of the coast phase is still curious though.
Reposted from the General Falcon discussion thread:
I notice that model shows what appeared to be LOX feedlines running down the outside of the stage. I was wondering why this was the case, as it had been previously established that the LOX feed runs down the center of the kerosene tank, with the tank making a torus shape around it.
But then I realized that by moving the LOX feed to the outside of the rocket, they gain a significant amount of volume on the inside. This could be a way to accomplish their stated goals of increasing tank capacity without requiring a tank stretch and the cascade of changes that would entail.
Oh yeah, gone from my timeline now. Wonder what's going on there? Accidental delete of the tweet? I've done that before! :)
Hmmm, saw that to and thought 3 things in addition to it being a feedline:Reposted from the General Falcon discussion thread:
I notice that model shows what appeared to be LOX feedlines running down the outside of the stage. I was wondering why this was the case, as it had been previously established that the LOX feed runs down the center of the kerosene tank, with the tank making a torus shape around it.
But then I realized that by moving the LOX feed to the outside of the rocket, they gain a significant amount of volume on the inside. This could be a way to accomplish their stated goals of increasing tank capacity without requiring a tank stretch and the cascade of changes that would entail.
Is that a LOX feedline? It looks like it runs the full length of the upper stage. If a LOX feed line wouldn't it only run from the bottom of the LOX tank down the outside?
the other shot
And anticipating some people now speculating that FH will have 8 engines per core in 3....2...1... Sigh, it is inevitable.
the center core's eight engines are sliced in half in an odd way, too.. would have thought those changes would make wind tunnel testing a bit off.
Hmmm, saw that to and thought 3 things in addition to it being a feedline:Reposted from the General Falcon discussion thread:
I notice that model shows what appeared to be LOX feedlines running down the outside of the stage. I was wondering why this was the case, as it had been previously established that the LOX feed runs down the center of the kerosene tank, with the tank making a torus shape around it.
But then I realized that by moving the LOX feed to the outside of the rocket, they gain a significant amount of volume on the inside. This could be a way to accomplish their stated goals of increasing tank capacity without requiring a tank stretch and the cascade of changes that would entail.
Is that a LOX feedline? It looks like it runs the full length of the upper stage. If a LOX feed line wouldn't it only run from the bottom of the LOX tank down the outside?
-Part of a new pressurization system
-Part of a new GSE system
-Needed for new chill system
On the Falcon Heavy rocket’s inaugural flight: “Later this year. The pad will be ready for it by September or October of this year. We’ll get it launched as quickly as we can."
SpaceNews article has more quotes from Gwynne this week: http://spacenews.com/spacex-aims-to-debut-new-version-of-falcon-9-this-summer/ (http://spacenews.com/spacex-aims-to-debut-new-version-of-falcon-9-this-summer/)The article *also* says:QuoteOn the Falcon Heavy rocket’s inaugural flight: “Later this year. The pad will be ready for it by September or October of this year. We’ll get it launched as quickly as we can."
In March 16 and 17 appearances at the Satellite 2015 conference here, Shotwell said the new-version Falcon 9, which has yet to be named, will be about 30 percent more powerful than the rocket’s current version.
Is this the 147,000 lb thrust engine or is it greater than that?
“Falcon Heavy is two different cores — the inner core and the two side sticks,” Shotwell said. “The new Falcon 9 will basically be a Falcon Heavy side booster. So we’re building [only two different] cores to make sure we don’t have a bunch of configurations around the factory so we can streamline operations and hit a launch cadence of one or two a month from every launch site we have.” - See more at: http://spacenews.com/spacex-aims-to-debut-new-version-of-falcon-9-this-summer/#sthash.Kk4wUTpC.dpuf
SpaceNews article has more quotes from Gwynne this week: http://spacenews.com/spacex-aims-to-debut-new-version-of-falcon-9-this-summer/ (http://spacenews.com/spacex-aims-to-debut-new-version-of-falcon-9-this-summer/)The article *also* says:QuoteOn the Falcon Heavy rocket’s inaugural flight: “Later this year. The pad will be ready for it by September or October of this year. We’ll get it launched as quickly as we can."QuoteIn March 16 and 17 appearances at the Satellite 2015 conference here, Shotwell said the new-version Falcon 9, which has yet to be named, will be about 30 percent more powerful than the rocket’s current version.
30%? Where did 30% come from?
We were told that the M1D engines were certified to 85% design spec before. A 17.6% increase in existing thrust would bring it to 100% design spec.
Are they going *further* than that?
And anticipating some people now speculating that FH will have 8 engines per core in 3....2...1... Sigh, it is inevitable.
I notice that model shows what appeared to be LOX feedlines running down the outside of the stage. I was wondering why this was the case, as it had been previously established that the LOX feed runs down the center of the kerosene tank, with the tank making a torus shape around it.I thought it had to do with crossfeed. But probably not. ..?
But then I realized that by moving the LOX feed to the outside of the rocket, they gain a significant amount of volume on the inside. This could be a way to accomplish their stated goals of increasing tank capacity without requiring a tank stretch and the cascade of changes that would entail.
do legs look the same as current ones? thought they would look different ..isn't there a new leg being developed?
jb
Reposted from the General Falcon discussion thread:
I notice that model shows what appeared to be LOX feedlines running down the outside of the stage. I was wondering why this was the case, as it had been previously established that the LOX feed runs down the center of the kerosene tank, with the tank making a torus shape around it.
But then I realized that by moving the LOX feed to the outside of the rocket, they gain a significant amount of volume on the inside. This could be a way to accomplish their stated goals of increasing tank capacity without requiring a tank stretch and the cascade of changes that would entail.
Pre-existing.
Credit: SpaceX.
Found on:
http://space.skyrocket.de/doc_lau_det/falcon-9_v1-1.htm
SpaceNews article has more quotes from Gwynne this week: http://spacenews.com/spacex-aims-to-debut-new-version-of-falcon-9-this-summer/ (http://spacenews.com/spacex-aims-to-debut-new-version-of-falcon-9-this-summer/)The article *also* says:QuoteOn the Falcon Heavy rocket’s inaugural flight: “Later this year. The pad will be ready for it by September or October of this year. We’ll get it launched as quickly as we can."QuoteIn March 16 and 17 appearances at the Satellite 2015 conference here, Shotwell said the new-version Falcon 9, which has yet to be named, will be about 30 percent more powerful than the rocket’s current version.
30%? Where did 30% come from?
We were told that the M1D engines were certified to 85% design spec before. A 17.6% increase in existing thrust would bring it to 100% design spec.
Are they going *further* than that?
SpaceNews article has more quotes from Gwynne this week: http://spacenews.com/spacex-aims-to-debut-new-version-of-falcon-9-this-summer/ (http://spacenews.com/spacex-aims-to-debut-new-version-of-falcon-9-this-summer/)The article *also* says:QuoteOn the Falcon Heavy rocket’s inaugural flight: “Later this year. The pad will be ready for it by September or October of this year. We’ll get it launched as quickly as we can."QuoteIn March 16 and 17 appearances at the Satellite 2015 conference here, Shotwell said the new-version Falcon 9, which has yet to be named, will be about 30 percent more powerful than the rocket’s current version.
30%? Where did 30% come from?
We were told that the M1D engines were certified to 85% design spec before. A 17.6% increase in existing thrust would bring it to 100% design spec.
Are they going *further* than that?
Careful, don't jump to conclusions. The 30% may simply be how much payload performance is improved. (To LEO or GTO) Not M1D Performance by itself.
SpaceNews article has more quotes from Gwynne this week: http://spacenews.com/spacex-aims-to-debut-new-version-of-falcon-9-this-summer/ (http://spacenews.com/spacex-aims-to-debut-new-version-of-falcon-9-this-summer/)The article *also* says:QuoteOn the Falcon Heavy rocket’s inaugural flight: “Later this year. The pad will be ready for it by September or October of this year. We’ll get it launched as quickly as we can."QuoteIn March 16 and 17 appearances at the Satellite 2015 conference here, Shotwell said the new-version Falcon 9, which has yet to be named, will be about 30 percent more powerful than the rocket’s current version.
30%? Where did 30% come from?
We were told that the M1D engines were certified to 85% design spec before. A 17.6% increase in existing thrust would bring it to 100% design spec.
Are they going *further* than that?
Careful, don't jump to conclusions. The 30% may simply be how much payload performance is improved. (To LEO or GTO) Not M1D Performance by itself.
That's exactly what it sounds like to me. 15% thrust upgrade + LOX subcooling + upper stage 10% stretch = 30% payload increase.
SpaceNews article has more quotes from Gwynne this week: http://spacenews.com/spacex-aims-to-debut-new-version-of-falcon-9-this-summer/ (http://spacenews.com/spacex-aims-to-debut-new-version-of-falcon-9-this-summer/)The article *also* says:QuoteOn the Falcon Heavy rocket’s inaugural flight: “Later this year. The pad will be ready for it by September or October of this year. We’ll get it launched as quickly as we can."QuoteIn March 16 and 17 appearances at the Satellite 2015 conference here, Shotwell said the new-version Falcon 9, which has yet to be named, will be about 30 percent more powerful than the rocket’s current version.
30%? Where did 30% come from?
We were told that the M1D engines were certified to 85% design spec before. A 17.6% increase in existing thrust would bring it to 100% design spec.
Are they going *further* than that?
Careful, don't jump to conclusions. The 30% may simply be how much payload performance is improved. (To LEO or GTO) Not M1D Performance by itself.
That's exactly what it sounds like to me. 15% thrust upgrade + LOX subcooling + upper stage 10% stretch = 30% payload increase.
Yup, me three - same thought. I also wondered if there might be a small Isp change, especially in light of the change to mixture ratio.
Don't forget also mild chilling of RP-1:-
AvWeek Paris (AvWeekParis):
Shotwell on Falcon 9 Merlin engine upgrade: We're doing slightly chilled RP, mostly chilled oxidizer. "Hey, if Russia can do it..." #satshow
http://twitter.com/AvWeekParis/status/577534249420468224
Cheers, Martin
That's about 5.88%. The CPI went up about 1.7% last year, but that's an imperfect inflation indicator.CPI is just about as good as anything else (although in the near-term, I'd probably exclude gas and food as they fluctuate a lot... you'll end up just looking at the variation in gas and food).
That's about 5.88%. The CPI went up about 1.7% last year, but that's an imperfect inflation indicator.
I expect it will be like with F9. Biggest jump in price will happen after paper rocket changes into actual rocket that launched successfully.
In the ULA thread reaching GEO was discussed. So Falcon upper stage will need the ability to restart after app. 12h coast. That's quite long. Is it a very large stretch to go from there to 3 days? 3 days would mean it could do LOI.I believe it's more like 8hs. It's a quarter of an orbit, and a GEO that would mean 6hrs, a GTO must be less.
In the ULA thread reaching GEO was discussed. So Falcon upper stage will need the ability to restart after app. 12h coast. That's quite long. Is it a very large stretch to go from there to 3 days? 3 days would mean it could do LOI.I believe it's more like 8hs. It's a quarter of an orbit, and a GEO that would mean 6hrs, a GTO must be less.
I'm lazy, so sue me :P It does require the knowledge that you increase your apogee at perigee, and then circularize and do the plane change simultaneously at apogee. This is for a stock GTO, if they were doing super-syncronous that would be a different profile and might well take 12hs or even more for bi-elliptic transfer. But that also require an extra restart and more coast time. I don't believe they'll want the added risk if they can spare the performance (which a Falcon Heavy should have plenty).In the ULA thread reaching GEO was discussed. So Falcon upper stage will need the ability to restart after app. 12h coast. That's quite long. Is it a very large stretch to go from there to 3 days? 3 days would mean it could do LOI.I believe it's more like 8hs. It's a quarter of an orbit, and a GEO that would mean 6hrs, a GTO must be less.
I especially like how you used the most advanced orbital mechanics simulation to conclude this. :)
I'm lazy, so sue me :P It does require the knowledge that you increase your apogee at perigee, and then circularize and do the plane change simultaneously at apogee. This is for a stock GTO, if they were doing super-syncronous that would be a different profile and might well take 12hs or even more for bi-elliptic transfer. But that also require an extra restart and more coast time. I don't believe they'll want the added risk if they can spare the performance (which a Falcon Heavy should have plenty).In the ULA thread reaching GEO was discussed. So Falcon upper stage will need the ability to restart after app. 12h coast. That's quite long. Is it a very large stretch to go from there to 3 days? 3 days would mean it could do LOI.I believe it's more like 8hs. It's a quarter of an orbit, and a GEO that would mean 6hrs, a GTO must be less.
I especially like how you used the most advanced orbital mechanics simulation to conclude this. :)
If the new engine pushes the Falcon heavy towards 60 tonnes to LEO, won't this make the $1 billion Per launch 70 tonne SLS look DOA. $150 mill for 60 tonnes.$1000 mill for 70 tonnes ummmSince cross feeding is no longer being used on the Falcon Heavy I suspect that the upgrades will be used to get it to its advertised lift capacity. SLS is not a good comparison to make. SLS will be capable of almost 90 tonnes from the start. The lift capacity is under reported. The program wanted room for mass growth should it happen, but it hasn't.
Does anyone know the figures for falcon heavy with updated engines to LEO
The better metric to use instead of payload to low orbit is payload to L2, TLI, TMI. Those are the likely destinations based on the current mission architecture concepts. The only official figures I could find for the Falcon Heavy were to LEO, GTO, or Mars. So the comparison will have to use Mars instead of the more likely L2 or TLI. The Falcon Heavy is capable of 13.2 tonnes to Mars. SLS will initially be able to do 20.2 to Mars, and 31.7 tonnes with the EUS which could be as soon as the 2nd or 3rd launch. SLS will have almost 2.5 times the lift capacity to BEO destinations. It will be able to send more on a Mars trajectory than any currently flying rocket can to LEO.
No it can't the math doesn't work out. Two Falcon Heavies can't throw as much through TMI. Its not just 13.2 + 13.2. You have to subtract the mass of the extra hardware needed to preform the docking, loiter in LEO, and control the stages. Even if one could just add those two numbers it wouldn't equal what an SLS can do once it gets the EUS which will be very soon in the program.If the new engine pushes the Falcon heavy towards 60 tonnes to LEO, won't this make the $1 billion Per launch 70 tonne SLS look DOA. $150 mill for 60 tonnes.$1000 mill for 70 tonnes ummmSince cross feeding is no longer being used on the Falcon Heavy I suspect that the upgrades will be used to get it to its advertised lift capacity. SLS is not a good comparison to make. SLS will be capable of almost 90 tonnes from the start. The lift capacity is under reported. The program wanted room for mass growth should it happen, but it hasn't.
Does anyone know the figures for falcon heavy with updated engines to LEO
The better metric to use instead of payload to low orbit is payload to L2, TLI, TMI. Those are the likely destinations based on the current mission architecture concepts. The only official figures I could find for the Falcon Heavy were to LEO, GTO, or Mars. So the comparison will have to use Mars instead of the more likely L2 or TLI. The Falcon Heavy is capable of 13.2 tonnes to Mars. SLS will initially be able to do 20.2 to Mars, and 31.7 tonnes with the EUS which could be as soon as the 2nd or 3rd launch. SLS will have almost 2.5 times the lift capacity to BEO destinations. It will be able to send more on a Mars trajectory than any currently flying rocket can to LEO.
Which means that the job of SLS can be done with an Earth Departure Stage lifted by a second Falcon Heavy to rendezvous in orbit with the payload. Two Falcon Heavy's plus an EDS are bound to be a lot cheaper than SLS.
Will it get it sooner than BFR flies?
No it can't the math doesn't work out. Two Falcon Heavies can't throw as much through TMI. Its not just 13.2 + 13.2. You have to subtract the mass of the extra hardware needed to preform the docking, loiter in LEO, and control the stages. Even if one could just add those two numbers it wouldn't equal what an SLS can do once it gets the EUS which will be very soon in the program.
Will it get it sooner than BFR flies?I would wager on it but my attention span doesn't allow me to make bets that far into the future. We know what the Block 1B SLS will look like. We know when they are planning to fly it. The factory to build it exists. The launch pad to launch it exists. We know what the engines to power it will look like. The engines are being qualified right now. The RL-10 is flying now. The RS-25s and SRB casings have already been built for the first EUS flight. SLS want to fly the EUS soon in the program with the second of third launch putting its inaugural flight in the 2021/2022 time frame.
I don't have time to give a detailed response to all these points. Also this is starting to become an SLS vs SpaceX thread so I will limit myself to just the SpaceX aspect of this. Falcon Heavy will not bet getting an LH2 upper stage. That goes against everything Elon has said about that fuel. Besides the gain in payload capacity with an LH2 upper stage would only push the payload LEO capacity into the low 60mt area. The LH2 upper stage would certainly be more expensive and only used by one customer on a very small number of flights. SpaceX's other customers don't have a need for cross-feeding let alone capacity beyond what that would provide. Not worth the extra cost for the improved performance.
No it can't the math doesn't work out. Two Falcon Heavies can't throw as much through TMI. Its not just 13.2 + 13.2. You have to subtract the mass of the extra hardware needed to preform the docking, loiter in LEO, and control the stages. Even if one could just add those two numbers it wouldn't equal what an SLS can do once it gets the EUS which will be very soon in the program.
-Use 3 or more launches.
-Use a hydrogen EDS
-Use a SEP EDS
-Not all missions may stretch top specs or perhaps can be configurable to lesser spec’ed layouts. For instance, if Orion flies with a module, then perhaps that module can be launched separately from the crew in the Falcon Heavy based approach. Not all Orion flights may carry along such modules and stretch the top specs of the system, as an example.
-You can use Falcon Heavy now and commission a SLS exact matching or exceeding LV later.
-You can investigate upgrades to Falcon Heavy like a hydrogen upper stage.
-Accept a lesser matching Falcon Heavy solution for the benefits a Falcon Heavy based approach brings to the table than top spec obsession.
Falcon Heavy has its own positive qualities like low cost mass delivery to orbit and greater budget space allocation to payload/technology/mission development with less ongoing parasitic LV cost and architectures can be tailored to harness those strengths rather than fit a SLS tailored plan to it, or hypothetical ones. A comparative basket that each approach could deliver for us I think would see the Falcon Heavy approach be more compelling even if they were not identical.
Also this is starting to become an SLS vs SpaceX thread so I will limit myself to just the SpaceX aspect of this.
Also this is starting to become an SLS vs SpaceX thread so I will limit myself to just the SpaceX aspect of this.
Even though it's not MCT, the Falcon Heavy will be able to launch bigger payloads to Mars and to elsewhere in the solar system than any rocket to date. I'm hoping there are institutions beyond SpaceX alone that will be able to take advantage of this capability.
If Raptor is 500k lb thrust and Merlin vacuum is 200k lb thrust, that is a lot more thrust. Would the upper be widened to hold more fuel, or stretched? Unless the Raptor vacuum can be throttled down.No doubt it'd be widened and/or stretched, but yeah, it can be throttled down (or at least they could certainly design it that way since they'll need to be able to throttle the first stage Raptors anyway).
Does anyone think they will make a metholox Merlin engine? If so, what kind of capacity would the Falcon heavy have converting all to metholox?
Does anyone think they will make a metholox Merlin engine? If so, what kind of capacity would the Falcon heavy have converting all to metholox?
My bold
What I would like to know is the expected isp of a gas generator methalox engine compared to a kerolox gg engine, and that would be a good beginning to answer your question.
If the difference is not substantial , I would stick to the speculation (adopted from the mini BFR thread) that with\without crossfeed, the next step is a reusable single raptor upper stage for FH, followed by replacing both F9R and FH cores with a 9 raptor reusable booster.
In a practical way you don't obtain substantial improvement with methane over querosene, the slight improvement on ISP (3,8%) is more than lost in the higher volume (22%). Falcon 9 is designed for road transport, if you switch to methane you have to increase the size of the tanks and you lose this capability, so it doesn't make sense.Does anyone think they will make a metholox Merlin engine? If so, what kind of capacity would the Falcon heavy have converting all to metholox?
My bold
What I would like to know is the expected isp of a gas generator methalox engine compared to a kerolox gg engine, and that would be a good beginning to answer your question.
If the difference is not substantial , I would stick to the speculation (adopted from the mini BFR thread) that with\without crossfeed, the next step is a reusable single raptor upper stage for FH, followed by replacing both F9R and FH cores with a 9 raptor reusable booster.
This might not be quite as big of an issue as is normally argued if they can launch from VAFB. It might be possible to transport a wider diameter stage from Hawthorne to VAFB by boat/barge. Though, they might lose the ability to test it in TX first.
In a practical way you don't obtain substantial improvement with methane over querosene, the slight improvement on ISP (3,8%) is more than lost in the higher volume (22%). Falcon 9 is designed for road transport, if you switch to methane you have to increase the size of the tanks and you lose this capability, so it doesn't make sense.
This might not be quite as big of an issue as is normally argued if they can launch from VAFB. It might be possible to transport a wider diameter stage from Hawthorne to VAFB by boat/barge. Though, they might lose the ability to test it in TX first.
What is the advantage? Kerosene is already great, very dense and easy to storage. If you don't need to collect from Mars is better than methane.
Does anyone think they will make a metholox Merlin engine? If so, what kind of capacity would the Falcon heavy have converting all to metholox?
My bold
What I would like to know is the expected isp of a gas generator methalox engine compared to a kerolox gg engine, and that would be a good beginning to answer your question.
If the difference is not substantial , I would stick to the speculation (adopted from the mini BFR thread) that with\without crossfeed, the next step is a reusable single raptor upper stage for FH, followed by replacing both F9R and FH cores with a 9 raptor reusable booster.
In a practical way you don't obtain substantial improvement with methane over querosene, the slight improvement on ISP (3,8%) is more than lost in the higher volume (22%). Falcon 9 is designed for road transport, if you switch to methane you have to increase the size of the tanks and you lose this capability, so it doesn't make sense.
Methane improves the reutilization of the engines through a cleaner combustion with les maintenance. The second good point of methane over querosene is the capacity of ISRU production in Mars.
What it would work would be a querosene closed cycle Merlin Engine to improve the ISP, with methane you would end up with a too long stick.
Methane has never been used for rockets because querosene and hydrogen work better. This might change with reutilization once you don't care about road transportation (BFR).
Well, the question was about a methalox Merlin (gas generator) , not about a kerolox mini Raptor (staged combustion) .
I reckon that was assuming the first option is easier to develop.
Thanks for your answer -
"the slight improvement on ISP (3,8%) is more than lost in the higher volume (22%)."
As I get it, since the rocket equation is exponential for isp, the 3,6% isp increase can't compare with the 22% volume reduction. Some math has to be made in order to know if the payload increases or decreases with these changes and the same total volume of the F9.
Falcon 9 is designed for road transport, if you switch to methane you have to increase the size of the tanks and you lose this capability, so it doesn't make sense.
But in the particular case of Falcon rockets there is no discussion as the tank size is the big constraint that can not grow.
But in the particular case of Falcon rockets there is no discussion as the tank size is the big constraint that can not grow. But even if it could grow I don't see any advantage.
Tank diameter is not an absolute constraint. It is a constraint that SpaceX has placed upon itself to reduce costs, if it ever truly gets in the way of an important goal it will be discarded. A true need to get FH a much higher energy upper stage might be enough reason if that market opportunity ever materializes.
But that would be years down the road, and SpaceX will likely never encounter the business need.
But this is no case for SpaceX. They are going from RP-1 Gas Generator to Methane Full Flow. Big difference. And actual RSC Energyia engineer has calculated that a methane Falcon 9 with the same dimensions as v1.1, but engine performance as Raptor, would get 25tonnes to LEO and 8tonnes to GTO. That covers 100% of current market. And that's without the 10% tank lengething and propellant densification that SpaceX is implementing on the enhanced Falcon 9, nor 2050 aluminum tanks and other "cheap" enhancements. Probably could hit 30/10 with that.Well, the question was about a methalox Merlin (gas generator) , not about a kerolox mini Raptor (staged combustion) .
I reckon that was assuming the first option is easier to develop.
Thanks for your answer -
"the slight improvement on ISP (3,8%) is more than lost in the higher volume (22%)."
As I get it, since the rocket equation is exponential for isp, the 3,6% isp increase can't compare with the 22% volume reduction. Some math has to be made in order to know if the payload increases or decreases with these changes and the same total volume of the F9.
If you keep the same tanks you have 22% less fuel, no way it compensates throught ISP. Otherwise Musk would have gone for methane from the beginning.
In case you make bigger tanks (22%) then you increase the tank weight in small percentage and is here where you loose the small advantage in ISP.
Same engine for Kerosene, Methane, Hydrogen give ISP of: 355,368,456s. For Hydrogen you always need bigger tanks with isolation, so for same weight of fuel you have much more weight, but the ISP (+24%) here is making big big difference. In kerosene VS methane the difference is really small and you probably end up in something symilar. But in the particular case of Falcon rockets there is no discussion as the tank size is the big constraint that can not grow. But even if it could grow I don't see any advantage.
But this is no case for SpaceX. They are going from RP-1 Gas Generator to Methane Full Flow. Big difference. And actual RSC Energyia engineer has calculated that a methane Falcon 9 with the same dimensions as v1.1, but engine performance as Raptor, would get 25tonnes to LEO and 8tonnes to GTO. That covers 100% of current market. And that's without the 10% tank lengething and propellant densification that SpaceX is implementing on the enhanced Falcon 9, nor 2050 aluminum tanks and other "cheap" enhancements. Probably could hit 30/10 with that.Well, the question was about a methalox Merlin (gas generator) , not about a kerolox mini Raptor (staged combustion) .
I reckon that was assuming the first option is easier to develop.
Thanks for your answer -
"the slight improvement on ISP (3,8%) is more than lost in the higher volume (22%)."
As I get it, since the rocket equation is exponential for isp, the 3,6% isp increase can't compare with the 22% volume reduction. Some math has to be made in order to know if the payload increases or decreases with these changes and the same total volume of the F9.
If you keep the same tanks you have 22% less fuel, no way it compensates throught ISP. Otherwise Musk would have gone for methane from the beginning.
In case you make bigger tanks (22%) then you increase the tank weight in small percentage and is here where you loose the small advantage in ISP.
Same engine for Kerosene, Methane, Hydrogen give ISP of: 355,368,456s. For Hydrogen you always need bigger tanks with isolation, so for same weight of fuel you have much more weight, but the ISP (+24%) here is making big big difference. In kerosene VS methane the difference is really small and you probably end up in something symilar. But in the particular case of Falcon rockets there is no discussion as the tank size is the big constraint that can not grow. But even if it could grow I don't see any advantage.
If they do a mini Raptor upper stage, they'll probably move the cores later. They'll have the performance margin for upper stage reuse and validate everything for MCT for a lot less money than a whole new development.
But this is no case for SpaceX. They are going from RP-1 Gas Generator to Methane Full Flow. Big difference. And actual RSC Energyia engineer has calculated that a methane Falcon 9 with the same dimensions as v1.1, but engine performance as Raptor, would get 25tonnes to LEO and 8tonnes to GTO. That covers 100% of current market. And that's without the 10% tank lengething and propellant densification that SpaceX is implementing on the enhanced Falcon 9, nor 2050 aluminum tanks and other "cheap" enhancements. Probably could hit 30/10 with that.
If they do a mini Raptor upper stage, they'll probably move the cores later. They'll have the performance margin for upper stage reuse and validate everything for MCT for a lot less money than a whole new development.
If Raptor is 500k lb thrust and Merlin vacuum is 200k lb thrust, that is a lot more thrust. Would the upper be widened to hold more fuel, or stretched? Unless the Raptor vacuum can be throttled down.
Well, the question was about a methalox Merlin (gas generator) , not about a kerolox mini Raptor (staged combustion) .
I reckon that was assuming the first option is easier to develop.
Thanks for your answer -
"the slight improvement on ISP (3,8%) is more than lost in the higher volume (22%)."
As I get it, since the rocket equation is exponential for isp, the 3,6% isp increase can't compare with the 22% volume reduction. Some math has to be made in order to know if the payload increases or decreases with these changes and the same total volume of the F9.
If you keep the same tanks you have 22% less fuel, no way it compensates throught ISP. Otherwise Musk would have gone for methane from the beginning.
In case you make bigger tanks (22%) then you increase the tank weight in small percentage and is here where you loose the small advantage in ISP.
Same engine for Kerosene, Methane, Hydrogen give ISP of: 355,368,456s. For Hydrogen you always need bigger tanks with isolation, so for same weight of fuel you have much more weight, but the ISP (+24%) here is making big big difference. In kerosene VS methane the difference is really small and you probably end up in something symilar. But in the particular case of Falcon rockets there is no discussion as the tank size is the big constraint that can not grow. But even if it could grow I don't see any advantage.
Do you think Spacex will do a min-raptor upper stage or just use the BO engine to replace the current upper stage?
Do you think Spacex will do a min-raptor upper stage or just use the BO engine to replace the current upper stage?
SpaceX is not going to use a non-SpaceX engine in any of its launch vehicles - period. Whatever they use will be built in-house.
Should be able to throttle lower than that.If Raptor is 500k lb thrust and Merlin vacuum is 200k lb thrust, that is a lot more thrust. Would the upper be widened to hold more fuel, or stretched? Unless the Raptor vacuum can be throttled down.
If you assume 550 klbf for the vac version (~250 tf), that would need a minimum ~20t payload for 6g burnout @ 50% throttle.
Fine for LEO with FHR, but not for GTO / escape unless they go with FHE (maybe recover the boosters?)
Good for a prop tanker for MCT, though.
Cheers, Martin
If Raptor is 500k lb thrust and Merlin vacuum is 200k lb thrust, that is a lot more thrust. Would the upper be widened to hold more fuel, or stretched? Unless the Raptor vacuum can be throttled down.
If you assume 550 klbf for the vac version (~250 tf), that would need a minimum ~20t payload for 6g burnout @ 50% throttle.
Fine for LEO with FHR, but not for GTO / escape unless they go with FHE (maybe recover the boosters?)
Good for a prop tanker for MCT, though.
Cheers, Martin
I think some of your figures might be a little off. The estimate I've seen several times for the ISP of a vacuum Raptor is 380, and the best estimate I've seen for the ISP of the present upper stage Falcon is 345. So this would be more like a 10% ISP improvement rather than 3.8%. And with the exponential nature of the rocket equation, this is quite significant.
You're missing the fact that part of the reason Raptor's engine cycle is feasible is because it's using methane instead of kerosene. So methane enables higher Isp by enabling a better engine cycle.I think some of your figures might be a little off. The estimate I've seen several times for the ISP of a vacuum Raptor is 380, and the best estimate I've seen for the ISP of the present upper stage Falcon is 345. So this would be more like a 10% ISP improvement rather than 3.8%. And with the exponential nature of the rocket equation, this is quite significant.
This ISP are between different fuels for same X engine. The big increase in ISP between merlin and raptor is mainly because of the full flow engine not because of the fuel. I think it would make a lot of sense to have a full flow engine for the upper stage and improve the ISP.
What I don't see is the switch to methane. A mini raptor methane upper stage would fit and maybe in 5 years we will see it in case they have a mini-raptor, but FMPOV a kerosene mini raptor would fit better. The switch between kerosene and methane "is not a big deal", so I don't see that crazy that once they have a mini raptor they adapt it to kerosene. But thats lot of speculation.
You're missing the fact that part of the reason Raptor's engine cycle is feasible is because it's using methane instead of kerosene. So methane enables higher Isp by enabling a better engine cycle.I think some of your figures might be a little off. The estimate I've seen several times for the ISP of a vacuum Raptor is 380, and the best estimate I've seen for the ISP of the present upper stage Falcon is 345. So this would be more like a 10% ISP improvement rather than 3.8%. And with the exponential nature of the rocket equation, this is quite significant.
This ISP are between different fuels for same X engine. The big increase in ISP between merlin and raptor is mainly because of the full flow engine not because of the fuel. I think it would make a lot of sense to have a full flow engine for the upper stage and improve the ISP.
What I don't see is the switch to methane. A mini raptor methane upper stage would fit and maybe in 5 years we will see it in case they have a mini-raptor, but FMPOV a kerosene mini raptor would fit better. The switch between kerosene and methane "is not a big deal", so I don't see that crazy that once they have a mini raptor they adapt it to kerosene. But thats lot of speculation.
If there was to be a higher(er) energy FH upper stage with an in-house engine, I'd expect it to be something like this> 5.2m barrel size (the same diameter as the PLF) and the same length as the kerosene-fuelled U/S. Whilst a 'Merlin-M' has not evidently in development, I'd certainly consider having it in a 'paper only' development phase to minimise delays if extra beyond-LEO performance is needed and Raptor is further off than a Methane conversion of the Merlin-VAC.
There is a lot of 'ifs' in that and I doubt it would happen unless Musk were certain of a customer or two who needed the performance. That said, a 360-380s U/S would be an interesting addition to the vehicle.
For starters, F9 v1.1 actually does about 16.5/17tonnes to LEO. Second, the Zenit-2 has a smaller upper stage in relationship. And third, the F9 v1.1 is around 95% of pmf in both first and second, while the Zenit-2 is 92% and 90%, respectively. Zenit-2 was a compromise with the Energyia boosters and has too much T/W and is just too heavy.But this is no case for SpaceX. They are going from RP-1 Gas Generator to Methane Full Flow. Big difference. And actual RSC Energyia engineer has calculated that a methane Falcon 9 with the same dimensions as v1.1, but engine performance as Raptor, would get 25tonnes to LEO and 8tonnes to GTO. That covers 100% of current market. And that's without the 10% tank lengething and propellant densification that SpaceX is implementing on the enhanced Falcon 9, nor 2050 aluminum tanks and other "cheap" enhancements. Probably could hit 30/10 with that.
If they do a mini Raptor upper stage, they'll probably move the cores later. They'll have the performance margin for upper stage reuse and validate everything for MCT for a lot less money than a whole new development.
Humm I cannot argue because I have no data but 25 tones losing 22% fuel mass is hard to believe. I would expect something similar to what you have now (13-14t to LEO). Zenit rocket with slightly less mass than F9 V1.1 and best full flow kerosene engine works 13500kg to LEO. Maybe could be slightly better, but 25 tones... anyway, time will tell.
You're missing the fact that part of the reason Raptor's engine cycle is feasible is because it's using methane instead of kerosene. So methane enables higher Isp by enabling a better engine cycle.
You're missing the fact that part of the reason Raptor's engine cycle is feasible is because it's using methane instead of kerosene. So methane enables higher Isp by enabling a better engine cycle.
You mean that the production of a methane stage combustion engine is easier because methane produce less corrossion in plumbing and combustion chamber than the dirty kerosene?
For starters, F9 v1.1 actually does about 16.5/17tonnes to LEO. Second, the Zenit-2 has a smaller upper stage in relationship. And third, the F9 v1.1 is around 95% of pmf in both first and second, while the Zenit-2 is 92% and 90%, respectively. Zenit-2 was a compromise with the Energyia boosters and has too much T/W and is just too heavy.Yep, let me re-read the SpaceX methalox forum I forgot that the data from Dimitry gave that high numbers.
More or less. The Raptor engine will be full flow staged combustion, so it will have both an oxygen-rich and methane-rich preburner. Kerosene-rich preburners just aren't done because they cause horrible coking problems in the turbine. There may be more to it to that, but that's my basic understanding of the issue.
You're missing the fact that part of the reason Raptor's engine cycle is feasible is because it's using methane instead of kerosene. So methane enables higher Isp by enabling a better engine cycle.
You mean that the production of a methane stage combustion engine is easier because methane produce less corrossion in plumbing and combustion chamber than the dirty kerosene?
More or less. The Raptor engine will be full flow staged combustion, so it will have both an oxygen-rich and methane-rich preburner. Kerosene-rich preburners just aren't done because they cause horrible coking problems in the turbine. There may be more to it to that, but that's my basic understanding of the issue.
If Raptor is 500k lb thrust and Merlin vacuum is 200k lb thrust, that is a lot more thrust. Would the upper be widened to hold more fuel, or stretched? Unless the Raptor vacuum can be throttled down.
If you assume 550 klbf for the vac version (~250 tf), that would need a minimum ~20t payload for 6g burnout @ 50% throttle.
Fine for LEO with FHR, but not for GTO / escape unless they go with FHE (maybe recover the boosters?)
Good for a prop tanker for MCT, though.
Cheers, Martin
Wouldn't this be 6g only if the upper stage itself was massless? I think the present one is about 4 mt empty. Switching from Merlin to Raptor probably will add 1 mt. If it's to be reusable, then you also would be looking at adding TPS, legs, grid fins + hydrolics, etc., not to mention the landing fuel. I can easily see this approaching 10 mt total. So at 50% throttle and a 20 mt payload, you'd only be looking at about 4 g at burnout.
You're missing the fact that part of the reason Raptor's engine cycle is feasible is because it's using methane instead of kerosene. So methane enables higher Isp by enabling a better engine cycle.
You mean that the production of a methane stage combustion engine is easier because methane produce less corrossion in plumbing and combustion chamber than the dirty kerosene?
More or less. The Raptor engine will be full flow staged combustion, so it will have both an oxygen-rich and methane-rich preburner. Kerosene-rich preburners just aren't done because they cause horrible coking problems in the turbine. There may be more to it to that, but that's my basic understanding of the issue.
Well I can imagine that. But in this case I don't understand how the normal gas generator cycle can work.
The gas generator is also kerosene rich and the gas gets pushed through the turbine just fine, before being dumped as one big black sooty stream of evil. So maybe the coking problem is with the injectors into the burning chamber instead? Those are fairly small orifices and should be much more vulnerable to coking.
If there was to be a higher(er) energy FH upper stage with an in-house engine, I'd expect it to be something like this> 5.2m barrel size (the same diameter as the PLF) and the same length as the kerosene-fuelled U/S. Whilst a 'Merlin-M' has not evidently in development, I'd certainly consider having it in a 'paper only' development phase to minimise delays if extra beyond-LEO performance is needed and Raptor is further off than a Methane conversion of the Merlin-VAC.A few days ago in another thread I worked out some very crude numbers for a Methane powered Falcon Heavy upper stage. Well not really, I used a fuel with methane's ISP and RP-1's density for a best case, unrealistic, yet super easy to calculate figure. Even if one could just up the ISP to 370s the payload goes up only by 3-4 mt. Falcon Heavy would not see even that much improvement to its LEO payload since the tanks would get bigger and heavier. One might be able to make the upper stage deliver a bigger portion of the total Delta V of the rocket by changing the stage sizes and do other tricks. However methane is not likely to to increase the payload to the point where replacing the upper stage and its engine is worth the headache.
There is a lot of 'ifs' in that and I doubt it would happen unless Musk were certain of a customer or two who needed the performance. That said, a 360-380s U/S would be an interesting addition to the vehicle.
COLORADO SPRINGS, Colo. — SpaceX says it sent the U.S. Air Force an updated letter of intent April 14 outlining a certification process for its Falcon Heavy rocket to launch national security satellites.
SpaceX hopes to have its Falcon Heavy rocket certified by 2017, Gwynne Shotwell, the company’s president and chief operating officer, told SpaceNews in an April 14 interview.
- See more at: http://spacenews.com/spacex-sends-air-force-an-outline-for-falcon-heavy-certification/#sthash.0Wzwy5m7.dpuf
SpaceX Sends Air Force an Outline for Falcon Heavy Certification - See more at: http://spacenews.com/spacex-sends-air-force-an-outline-for-falcon-heavy-certification/#sthash.0Wzwy5m7.dpufInteresting. It is after all exactly what the last Assured Access Congressional Hearing was requesting. I believe it was something to the effect of, "Please have that FH ready as soon as possible." A 2017 certification, IMO, puts real pressure on ULA in that it makes needing to extend the RD-180 past 2019 a mute point. (Except for the VI issue) This is going to get very interesting. Not that it isn't already.QuoteCOLORADO SPRINGS, Colo. — SpaceX says it sent the U.S. Air Force an updated letter of intent April 14 outlining a certification process for its Falcon Heavy rocket to launch national security satellites.
SpaceX hopes to have its Falcon Heavy rocket certified by 2017, Gwynne Shotwell, the company’s president and chief operating officer, told SpaceNews in an April 14 interview.
- See more at: http://spacenews.com/spacex-sends-air-force-an-outline-for-falcon-heavy-certification/#sthash.0Wzwy5m7.dpuf
Interesting. It is after all exactly what the last Assured Access Congressional Hearing was requesting. I believe it was something to the effect of, "Please have that FH ready as soon as possible." A 2017 certification, IMO, puts real pressure on ULA in that it makes needing to extend the RD-180 past 2019 a mute point. (Except for the VI issue) This is going to get very interesting. Not that it isn't already.
Shotwell said she expects the Falcon Heavy rocket to fly once this year, three times in 2016 and three to five times in 2017.
“The market is huge,” she said. “The market is bigger in the commercial marketplace than it is for the single stick Falcon 9.”
Surprising...Quote from: Gwynne ShotwellShotwell said she expects the Falcon Heavy rocket to fly once this year, three times in 2016 and three to five times in 2017.
“The market is huge,” she said. “The market is bigger in the commercial marketplace than it is for the single stick Falcon 9.”
Surprising...Very. I'm not saying that she's misstating but I do wonder who the potential customers in the 50t IMLEO/17t GTO market might be.I think the better way of thinking about it is the number of customers in the >6t to GTO market at a price of roughly $100m. That's about half Ariane, and competitive with Proton without the issues of working with Russians. Even without employing the ful capacity, Falcon Heavy can be the cheaper $/kg option. The customer doesn't have to be using the full capacity to choose FH, they just have to be looking for (1) more than F9R can lift and (2) not seeing a competing launcher in the class that can match that price.
competitive with Proton without the issues of working with Russians.
Do you guys think they will develop a methane Merlin?
I thought they were working on a methane Merlin at one time.
From the café on the mezzanine I can see twin enormous nose cones sitting in the next room, waiting to be used on the demo flight of the Falcon Heavy [...]
During the tour I also saw dozens of people working on the various components of the Falcons 9 and Heavy.
Any info about planned performance of FH to Sun Earth L2 point?
thx
(I tried to search for it, but didnt find it)
Any info about planned performance of FH to Sun Earth L2 point?
Arabsat said it will sign a contract with SpaceX to launch the Arabsat 6A satellite. Arabsat 6A is sized to launch on SpaceX’s Falcon Heavy rocket, according to industry officials.http://spaceflightnow.com/2015/04/29/arabsat-contracts-go-to-lockheed-martin-arianespace-and-spacex/
The company did not identify a launch site for Arabsat 6A, which Lockheed Martin officials said would be ready to fly in 2018, when SpaceX plans to have launch pads ready at Kennedy Space Center in Florida and near Brownsville, Texas, to support Falcon Heavy launches.
The Falcon Heavy’s inaugural test flight is scheduled later this year.
Do you guys think they will develop a methane Merlin? If so, what performance increase to GTO would that bring to Falcon Heavy with a properly sized upper stage, (stretched if need be)?I have read about it here several times.
Do you guys think they will develop a methane Merlin? If so, what performance increase to GTO would that bring to Falcon Heavy with a properly sized upper stage, (stretched if need be)?
Rocket | Falcon 9-M | Falcon Heavy-M (cross-feed) |
Payload to LEO | 24.93 mt | 78.16 mt |
Gross Mass | 446 mt | 1262.99 mt |
Diameter | 3.66 m | 3.66 m x 3 |
SI Gross Mass | 358.47 mt | 763.77 mt |
SI Propellant Mass | 335.06 mt | 716.94 mt |
SI Engines | 9xMini-Raptor | 27xMini-Raptor |
SI SL Thrust | 576 tf | 1728 tf |
SI Vac Thrust | 651.4 tf | 1954.2 tf |
SI Engine Isp | 321/363 | 321/363 |
SII Gross Mass | 60.90 mt | 358.47 mt |
SII Propellant Mass | 56.80 mt | 335.06 mt |
SII SL Thrust | N/A | 576 tf |
SII Vac Thrust | 70 tf | 651.4 tf |
SII Engine Isp | 380 | 321/363 |
SIII Gross Mass | N/A | 60.90 mt |
SIII Propellant Mass | N/A | 56.80 mt |
SIII Vac Thrust | N/A | 70 tf |
SIII Isp | N/A | 380 |
PLF Mass | 1.70 mt | 1.70 mt |
PLF separation time (sec.) | 220 | 220 |
Please review this pictures of the Angara-3 cross-feed preliminary design. Regrattably I don't speak Russian. It is different because this system is proposed at the tank level, while FH would be at the manifold level. I would also suggest that you look at the Booster Engine separation details, which is a bit more similar.
One issue you have is that you interrupt the flow of propellant with the spool valve. You need a a system that give priority to the propellant from the boosters, but won't interrupt the flow on separation, and then you need a one-way valve to avoid losing the propellant at separation. You'd also need a one way valve on the incoming pipes from the the core tank, since the boosters would probably need to supply a higher head pressure and thus you need to avoid a retro flow.
Please review this pictures of the Angara-3 cross-feed preliminary design.
Do you guys think they will develop a methane Merlin? If so, what performance increase to GTO would that bring to Falcon Heavy with a properly sized upper stage, (stretched if need be)?
They might, though if you're going to go "all-methalox", you generally want a big increase in performance to compensate for the rocket's increased dry mass and production costs. That's why I would suggest something like the following:
The only reason you'd prefer RP1 is bulk density, but CH4/O2 is pretty nearly the same bulk density, actually. ~830kg/m^3 vs 1030 for kerolox and 360 for hydrolox.Do you guys think they will develop a methane Merlin? If so, what performance increase to GTO would that bring to Falcon Heavy with a properly sized upper stage, (stretched if need be)?
They might, though if you're going to go "all-methalox", you generally want a big increase in performance to compensate for the rocket's increased dry mass and production costs. That's why I would suggest something like the following:
FWIW, I've long favoured a Falcon Hybrid, with a kerolox core (because RP1 performs better as an atmospheric fuel) and LCH4 in the upper stage (to get better ISP for BLEO insertion burns).
IF, big if, they did make a Methane Merlin and directly replaced the kerosene on the Falcon 9, (of course they would have to adjust tank size via the common bulkhead), what kind of performance would the Falcon 9 and Falcon H have?
IF, big if, they did make a Methane Merlin and directly replaced the kerosene on the Falcon 9, (of course they would have to adjust tank size via the common bulkhead), what kind of performance would the Falcon 9 and Falcon H have?
Please see Hyperion5's post here (http://forum.nasaspaceflight.com/index.php?topic=36806.msg1372522#msg1372522).
Instead of a spool 3-way valve, you should use two on/off valves (one for in-core tank, one for booster incoming flow), a one way valve after the booster incoming flow valve to close the circuit after separation. And another one-way valve before the in core flow valve to avoid back flow. And somehow design the piping in such a way that it leaves no bubbles (or ad a heavy gas trap). Also, you should make sure that the process of opening one valve and close the other can comply with two requirements: no head pressure drop below the minimum engine specified AND generates no turbulence that may end up in cavitation.Please review this pictures of the Angara-3 cross-feed preliminary design.
Well I disagree with your comments about shortcomings with my design, I am glad you posted the Angara pictures. Without understanding Russian, it is clear that tank-to-tank cross flow is viable (which I didn't think it was). Clearly by pressurizing the outboard booster tanks and not the central ones, sufficient fuel will flow from the outboard tanks to the central ones to power the central engine. The diagram shows the central tanks remaining completely full.
IF, big if, they did make a Methane Merlin and directly replaced the kerosene on the Falcon 9, (of course they would have to adjust tank size via the common bulkhead), what kind of performance would the Falcon 9 and Falcon H have?
Please see Hyperion5's post here (http://forum.nasaspaceflight.com/index.php?topic=36806.msg1372522#msg1372522).
That assumes replacing the gas generator Merlin with a "mini Raptor" methane staged combustion engine that has equal thrust to Merlin. Replacing the current Merlin with an equivalent gas generator methane engine would probably give you a small decrease in performance, since methane's ~20% lower bulk density will have a slightly larger impact than the ~4% increase in ISP you get. There are some advantages to switching to methane, but you really need an entirely new high performance engine to make a big difference.
Have Dimitry done any simulation an an Falcon 9 v1.2 M? I suspect it might cover Delta IV Heavy payloads in expendable mode.Do you guys think they will develop a methane Merlin? If so, what performance increase to GTO would that bring to Falcon Heavy with a properly sized upper stage, (stretched if need be)?
They might, though if you're going to go "all-methalox", you generally want a big increase in performance to compensate for the rocket's increased dry mass and production costs. That's why I would suggest something like the following:
Detailed figures on both Falcon 9-M & Falcon Heavy-M below:
Rocket Falcon 9-M Falcon Heavy-M (cross-feed) Payload to LEO 24.93 mt 78.16 mt Gross Mass 446 mt 1262.99 mt Diameter 3.66 m 3.66 m x 3 SI Gross Mass 358.47 mt 763.77 mt SI Propellant Mass 335.06 mt 716.94 mt SI Engines 9xMini-Raptor 27xMini-Raptor SI SL Thrust 576 tf 1728 tf SI Vac Thrust 651.4 tf 1954.2 tf SI Engine Isp 321/363 321/363 SII Gross Mass 60.90 mt 358.47 mt SII Propellant Mass 56.80 mt 335.06 mt SII SL Thrust N/A 576 tf SII Vac Thrust 70 tf 651.4 tf SII Engine Isp 380 321/363 SIII Gross Mass N/A 60.90 mt SIII Propellant Mass N/A 56.80 mt SIII Vac Thrust N/A 70 tf SIII Isp N/A 380 PLF Mass 1.70 mt 1.70 mt PLF separation time (sec.) 220 220
Instead of a spool 3-way valve, you should use two on/off valves (one for in-core tank, one for booster incoming flow), a one way valve after the booster incoming flow valve to close the circuit after separation. And another one-way valve before the in core flow valve to avoid back flow. And somehow design the piping in such a way that it leaves no bubbles (or ad a heavy gas trap). Also, you should make sure that the process of opening one valve and close the other can comply with two requirements: no head pressure drop below the minimum engine specified AND generates no turbulence that may end up in cavitation.Please review this pictures of the Angara-3 cross-feed preliminary design.
Well I disagree with your comments about shortcomings with my design, I am glad you posted the Angara pictures. Without understanding Russian, it is clear that tank-to-tank cross flow is viable (which I didn't think it was). Clearly by pressurizing the outboard booster tanks and not the central ones, sufficient fuel will flow from the outboard tanks to the central ones to power the central engine. The diagram shows the central tanks remaining completely full.
I don't know what's your experience with cryo liquids, but it's not that easy (but totally doable). Just not possible with a 3-way valve (which are specially heavy for big diameter piping.
In this video from the Planetary Society ...
Bill Nye and others are discussing the launch of Light Sail on the Falcon Heavy, Bill Nye states that it will be the first flight of the FH. Does TPS know something we don't, given that they are customers (i.e. the first flight is now 2016) or is Mr. Nye mistaken?
During the Light Sail reveal event last year it was stated that PROX- 1 would be a secondary payload. I was under the impression that SpaceX was still aiming to fly the demo mission by the end of 2015, but Mr. Nye states that Light Sail will fly on the first flight of the FH in 2016. I'm just wondering if anyone here has heard if the demo mission has shifted to the right or if Mr. Nye mispoke.In this video from the Planetary Society ...
Bill Nye and others are discussing the launch of Light Sail on the Falcon Heavy, Bill Nye states that it will be the first flight of the FH. Does TPS know something we don't, given that they are customers (i.e. the first flight is now 2016) or is Mr. Nye mistaken?
FWIW, references to FH start at about the 3:30 mark, and include CG video. More details on the mission are at the kickstarter page:
https://www.kickstarter.com/projects/theplanetarysociety/lightsail-a-revolutionary-solar-sailing-spacecraft
which mentions that the sail will be carried by yet another experimental spacecraft (PROX-1 from Georgia Tech); all are pretty specific about FH being the launch vehicle.
One presumes that these are secondary payloads; nothing here says much specific about the primary, though.
He didn't misspeak, he just left out "operational" (as stated in the Kickstarter campaign description and likely elsewhere).During the Light Sail reveal event last year it was stated that PROX- 1 would be a secondary payload. I was under the impression that SpaceX was still aiming to fly the demo mission by the end of 2015, but Mr. Nye states that Light Sail will fly on the first flight of the FH in 2016. I'm just wondering if anyone here has heard if the demo mission has shifted to the right or if Mr. Nye mispoke.In this video from the Planetary Society ...
Bill Nye and others are discussing the launch of Light Sail on the Falcon Heavy, Bill Nye states that it will be the first flight of the FH. Does TPS know something we don't, given that they are customers (i.e. the first flight is now 2016) or is Mr. Nye mistaken?
FWIW, references to FH start at about the 3:30 mark, and include CG video. More details on the mission are at the kickstarter page:
https://www.kickstarter.com/projects/theplanetarysociety/lightsail-a-revolutionary-solar-sailing-spacecraft
which mentions that the sail will be carried by yet another experimental spacecraft (PROX-1 from Georgia Tech); all are pretty specific about FH being the launch vehicle.
One presumes that these are secondary payloads; nothing here says much specific about the primary, though.
Thanks, I must have missed that in the Kickstarter description. I figured that that was what he meant, but I wasn't sure.He didn't misspeak, he just left out "operational" (as stated in the Kickstarter campaign description and likely elsewhere).During the Light Sail reveal event last year it was stated that PROX- 1 would be a secondary payload. I was under the impression that SpaceX was still aiming to fly the demo mission by the end of 2015, but Mr. Nye states that Light Sail will fly on the first flight of the FH in 2016. I'm just wondering if anyone here has heard if the demo mission has shifted to the right or if Mr. Nye mispoke.In this video from the Planetary Society ...
Bill Nye and others are discussing the launch of Light Sail on the Falcon Heavy, Bill Nye states that it will be the first flight of the FH. Does TPS know something we don't, given that they are customers (i.e. the first flight is now 2016) or is Mr. Nye mistaken?
FWIW, references to FH start at about the 3:30 mark, and include CG video. More details on the mission are at the kickstarter page:
https://www.kickstarter.com/projects/theplanetarysociety/lightsail-a-revolutionary-solar-sailing-spacecraft
which mentions that the sail will be carried by yet another experimental spacecraft (PROX-1 from Georgia Tech); all are pretty specific about FH being the launch vehicle.
One presumes that these are secondary payloads; nothing here says much specific about the primary, though.
Love Bill Nye, but I don't hang on every word he says when it's about something very technical.I would agree with that, but I do love his enthusiasm
The recent F9H video showed black legs and leg nacelles...
Is this extra thermal protection than we have had to date?
Does anyone know when we might see this leg configuration for the first time? Perhaps on a "regular" F9 launch soon?
Thanks,
Paul.
The recent F9H video showed black legs and leg nacelles...
Is this extra thermal protection than we have had to date?
Does anyone know when we might see this leg configuration for the first time? Perhaps on a "regular" F9 launch soon?
Thanks,
Paul.
I always assumed that SpaceX intentionally used this "irregular" coloring scheme to easily differentiate between actual pictures and renderings/models. Though I like cscott's notion that it might represent composites, too.
This is pure speculation on my part, but I was wondering if SpaceX might take advantage of the demo launch of Falcon Heavy to launch a Dragon 2 (mostly complete but without support for humans) to try a propulsive landing at Vandenberg?
This is pure speculation on my part, but I was wondering if SpaceX might take advantage of the demo launch of Falcon Heavy to launch a Dragon 2 (mostly complete but without support for humans) to try a propulsive landing at Vandenberg?
A retro fitted, previously flown Dragon 1, would make more sense, than a Dragon 2. This is assuming, the Dragonfly tests, have progressed far enough, to give it a try.
This is pure speculation on my part, but I was wondering if SpaceX might take advantage of the demo launch of Falcon Heavy to launch a Dragon 2 (mostly complete but without support for humans) to try a propulsive landing at Vandenberg?
A retro fitted, previously flown Dragon 1, would make more sense, than a Dragon 2. This is assuming, the Dragonfly tests, have progressed far enough, to give it a try.
Hmm, the Dragon 1 don't have Super-Dracos or do propulsive landings.
...
Changing to a black color scheme for the legs and aft end of the stage might make sense once re-use becomes standard procedure. We've seen how the aft end gets sooted/scorched during retro burn, and black paint might cover scorch marks up for reuse better than white paint.
Changing to a black color scheme has the possible benefit not painting the composite legs which could be carbon fiber, saving production time and weight. Does this increase performance and the payload to orbit?In theory, any weight savings increases maximum performance, but not necesarily meaningfully. If there's 100 kg of paint on the legs (and we know they only mass about 2000 kg total, so it's probably far less), then the improvement in maximum payload would be about a tenth of that--the rule of thumb is something like ten kg off the first stage for one extra kg of payload. Thus, even in this crazy "hundred kg of paint" case, you're only gaining 10kg of maximum payload out of tons, when many of F9 v1.1's launches so far have left significant room on the table below the maximum payload.
Then the cost factor of not having to paint the legs, inspecting the paint, repair scratches, rework, clean or re-paint the legs on reuse, etc may be the biggest benefit if they choose to 'go commando'.Changing to a black color scheme has the possible benefit not painting the composite legs which could be carbon fiber, saving production time and weight. Does this increase performance and the payload to orbit?In theory, any weight savings increases maximum performance, but not necesarily meaningfully. If there's 100 kg of paint on the legs (and we know they only mass about 2000 kg total, so it's probably far less), then the improvement in maximum payload would be about a tenth of that--the rule of thumb is something like ten kg off the first stage for one extra kg of payload. Thus, even in this crazy "hundred kg of paint" case, you're only gaining 10kg of maximum payload out of tons, when many of F9 v1.1's launches so far have left significant room on the table below the maximum payload.
They could use the opportunity to show off the capabilities of the Science Dragon initiative they've been pushing. I don't know if they have a spare Dragonv2 laying around that they could throw on top but I would assume that you could easily load one with cameras, radiation sensors, etc. and make a trip around the moon without much issue and make SpaceX a household name for a week.This is pure speculation on my part, but I was wondering if SpaceX might take advantage of the demo launch of Falcon Heavy to launch a Dragon 2 (mostly complete but without support for humans) to try a propulsive landing at Vandenberg?
I'm sure they've thought of it. Could re-fly one of their once-used Cargo Dragons, maybe around the Moon? That would inspire a lot of conversations...
I am not sure about modern carbon composites, but as recently as 15 years ago dark colors were not allowed on composite structures to keep them from over heating in the harshest solar heating situations.
Matthew
With the Falcon Heavy, which we plan to launch later this year, fly three times next year and certify soon thereafter, SpaceX will be able to launch 100 percent of the DOD's manifest.
Sounds to me as if they are still on track for a FH flight this year. It might slop, of course, but IMHO they wouldn't have said this year if they knew it was false.As Dilbert's boss said (http://dilbert.com/strip/2001-11-19), optimism is not crime.
Sounds to me as if they are still on track for a FH flight this year. It might slop, of course, but IMHO they wouldn't have said this year if they knew it was false.As Dilbert's boss said (http://dilbert.com/strip/2001-11-19), optimism is not crime.
I will believe about FH launch in 2015 when I will see it, not picosecond earlier.
Sounds to me as if they are still on track for a FH flight this year. It might slop, of course, but IMHO they wouldn't have said this year if they knew it was false.As Dilbert's boss said (http://dilbert.com/strip/2001-11-19), optimism is not crime.
I will believe about FH launch in 2015 when I will see it, not picosecond earlier.
Agreed, if it's even vertical in Texas or Florida this year I'll be happy.
Given the F9 flight rate are we even sure they can build cores quick enough?
Sounds to me as if they are still on track for a FH flight this year. It might slop, of course, but IMHO they wouldn't have said this year if they knew it was false.As Dilbert's boss said (http://dilbert.com/strip/2001-11-19), optimism is not crime.
I will believe about FH launch in 2015 when I will see it, not picosecond earlier.
Agreed, if it's even vertical in Texas or Florida this year I'll be happy.
Given the F9 flight rate are we even sure they can build cores quick enough?
From a SpaceX talk by upper-level management I was at last week, they said their current capacity is 18 cores/year, but they're ramping up for 40/year.
Sounds to me as if they are still on track for a FH flight this year. It might slop, of course, but IMHO they wouldn't have said this year if they knew it was false.As Dilbert's boss said (http://dilbert.com/strip/2001-11-19), optimism is not crime.
I will believe about FH launch in 2015 when I will see it, not picosecond earlier.
Agreed, if it's even vertical in Texas or Florida this year I'll be happy.
Given the F9 flight rate are we even sure they can build cores quick enough?
From a SpaceX talk by upper-level management I was at last week, they said their current capacity is 18 cores/year, but they're ramping up for 40/year.
Since the max usage for this year looks to be 14 cores (11 F9 and 1 FH) If they build at the max current rate for 2 years at 18 each year that means that there would be 22 cores at least available for next year or 3 FH and 13 F9 launches. Now add in eventual F9 reuse and launch rate could be higher.
I do get giddy at the idea of 360 Merlin 1D's per year. Amazing.400 if you think stage two counts ... and it should.
Four scheduled launches of the world's soon-to-be largest rocket in first year (or two)... that will be something to watch! Did I mention reusable? Three landing attempts each launch!
Delta IV Heavy (half the payload, three times the price) has only launched eight times in its eleven year lifetime.
Mr. Thornburg confirmed that they would self-fund the first launch.
We still don't know the payload/destination, do we?
Three landing attempts each launch!
Delta IV Heavy (half the payload, three times the price) has only launched eight times in its eleven year lifetime.
Three landing attempts each launch!
Delta IV Heavy (half the payload, three times the price) has only launched eight times in its eleven year lifetime.
A. the payload is greater to GSO which is was designed for and not LEO
b. That should be a hint at the small market for it
c. Three times the cores to blow up
Sounds to me as if they are still on track for a FH flight this year. It might slop, of course, but IMHO they wouldn't have said this year if they knew it was false.
FH Spring next year.
NET Spring 2016FH Spring next year.
http://forum.nasaspaceflight.com/index.php?topic=37739.340
NET Spring 2016FH Spring next year.
http://forum.nasaspaceflight.com/index.php?topic=37739.340
He said Spring and then added maybe April 2016.
Re: CRS-7 - it's mind boggling to think that the arrival of the first humans on Mars was likely pushed back further by 6 more months because a few billion atoms of iron, carbon, chromium and whatnot were in the wrong place at the wrong time.
My 0.02$ : Space X is keeping a low profile because it would seem too audacious to announce FH after the CRS 7 kaboom.
42:15 Stephen Clark: And the Falcon Heavy?
42:22 Elon Musk: Given our focus on Falcon 9, we've deprioritized the Falcon Heavy to probably launch in the spring next year. So maybe April or so.
SpaceX has updated the renderings on their F9/FH pages with the "v1.2" renderings, which are more detailed, and presumably accurate. It does seem to indicate a ~6ft stretch as well in the upper stage.
Here is the old (left) FH rendering compared to the new one (right), but it is difficult to match them exactly:
Changes:
- stretched upper stage
- more detail on booster attachment
- grid fins
SpaceX has updated the renderings on their F9/FH pages with the "v1.2" renderings, which are more detailed, and presumably accurate. It does seem to indicate a ~6ft stretch as well in the upper stage.
Here is the old (left) FH rendering compared to the new one (right), but it is difficult to match them exactly:
Changes:
- stretched upper stage
- more detail on booster attachment
- grid fins
And a larger interstage.
Cheers, Martin
SpaceX has updated the renderings on their F9/FH pages with the "v1.2" renderings, which are more detailed, and presumably accurate. It does seem to indicate a ~6ft stretch as well in the upper stage.
Here is the old (left) FH rendering compared to the new one (right), but it is difficult to match them exactly:
Changes:
- stretched upper stage
- more detail on booster attachment
- grid fins
And a larger interstage.
Cheers, Martin
SpaceX has updated the renderings on their F9/FH pages with the "v1.2" renderings, which are more detailed, and presumably accurate. It does seem to indicate a ~6ft stretch as well in the upper stage.
Here is the old (left) FH rendering compared to the new one (right), but it is difficult to match them exactly:
Changes:
- stretched upper stage
- more detail on booster attachment
- grid fins
And a larger interstage.
Cheers, Martin
Also, the leg's seem shorter and the center engine alignment is different. This could be just the rendering ofcourse.
SpaceX has updated the renderings on their F9/FH pages with the "v1.2" renderings, which are more detailed, and presumably accurate. It does seem to indicate a ~6ft stretch as well in the upper stage.
Here is the old (left) FH rendering compared to the new one (right), but it is difficult to match them exactly:
Changes:
- stretched upper stage
- more detail on booster attachment
- grid fins
And a larger interstage.
Cheers, Martin
It indeed does looks a bit longer.
Also, the leg's seem shorter and the center engine alignment is different. This could be just the rendering ofcourse.
No C/F With C/F
FHRv1.1 FHv1.1 FHRv1.2 FHv1.2 FHRv1.1 FHv1.1 FHRv1.2 FHv1.2
36 44.3 41 50.5 43 53 49 60.5
These values although very rough estimate numbers shows the relationships between the original FHv1.1 versions capabilities and the newer FHv1.2 versions capabilities. [Versions are [R][C/F:woC/F], four in all for each v.
The R versions assume the same landing of booster RTLS and center core ASDS landing downrange.
The conclusion is that unless you need something with a capability of > 49mt then a reusable configuration would work just fine. Even woC/F you get about 41mt reusable.
No C/F With C/F
FHRv1.1 FHv1.1 FHRv1.2 FHv1.2 FHRv1.1 FHv1.1 FHRv1.2 FHv1.2
36 44.3 41 50.5 43 53 49 60.5
These values although very rough estimate numbers shows the relationships between the original FHv1.1 versions capabilities and the newer FHv1.2 versions capabilities. [Versions are [R][C/F:woC/F], four in all for each v.
The R versions assume the same landing of booster RTLS and center core ASDS landing downrange.
The conclusion is that unless you need something with a capability of > 49mt then a reusable configuration would work just fine. Even woC/F you get about 41mt reusable.
Did you run your centre core expendable but not crossfed models with something like the following centre core firing engine profile (note time in seconds):
L-2 to L+13 9 centre core engines firing at 100% thrust (uses 8% of centre core fuel - 92% remaining)
L+13 to L+30 7 centre core engines firing at 100% thrust (uses 8% of centre core fuel - 83% remaining)
L+30 to L+60 4 centre core engines firing at 100% thrust (uses 7% of centre core fuel - 76% remaining)
L+60 to L+154 4 centre core engines firing at 70% thrust (uses 23% of centre core fuel - 53% remaining)
L+154 side boosters seperate with 20% fuel remaining in each
L+154 to 262 4 centre core engines firing at 100% thrust
L+262 to 479 2 centre core engines firing at 100% thrust
No C/F With C/F
FHRv1.1 FHv1.1 FHRv1.2 FHv1.2 FHRv1.1 FHv1.1 FHRv1.2 FHv1.2
36 44.3 41 50.5 43 53 49 60.5
These values although very rough estimate numbers shows the relationships between the original FHv1.1 versions capabilities and the newer FHv1.2 versions capabilities. [Versions are [R][C/F:woC/F], four in all for each v.
The R versions assume the same landing of booster RTLS and center core ASDS landing downrange.
The conclusion is that unless you need something with a capability of > 49mt then a reusable configuration would work just fine. Even woC/F you get about 41mt reusable.
Did you run your centre core expendable but not crossfed models with something like the following centre core firing engine profile (note time in seconds):
L-2 to L+13 9 centre core engines firing at 100% thrust (uses 8% of centre core fuel - 92% remaining)
L+13 to L+30 7 centre core engines firing at 100% thrust (uses 8% of centre core fuel - 83% remaining)
L+30 to L+60 4 centre core engines firing at 100% thrust (uses 7% of centre core fuel - 76% remaining)
L+60 to L+154 4 centre core engines firing at 70% thrust (uses 23% of centre core fuel - 53% remaining)
L+154 side boosters seperate with 20% fuel remaining in each
L+154 to 262 4 centre core engines firing at 100% thrust
L+262 to 479 2 centre core engines firing at 100% thrust
As you point out optimizing the burn profile can change the payload amounts for the expendable FH. My numbers are nothing more than a set of estimates showing the relationships of payloads capabilities between all of configurations. The values can be +-a few mt. So configurations within about 5% (~+-2.5mt) values are considered equal.
The new choices for payloads related to the configurations required in the FHv1.2 would be
<40 FHRv1.2woC/F
<50 FHRv1.2C/F
<60 FHv1.2
What I am trying to point out is that the goals for FH of 50mt can be done with a FHRv1.2C/F lowering the costs of that payload size class by $75M making flying with C/F almost the same cost as woC/F (each C/F booster costs $10M more to make) such that a FHRv1.2woC/F could be priced at as low as $50M and a FHRv1.2C/F as low as $60M.
SpaceX has updated the renderings on their F9/FH pages with the "v1.2" renderings, which are more detailed, and presumably accurate. It does seem to indicate a ~6ft stretch as well in the upper stage.
Here is the old (left) FH rendering compared to the new one (right), but it is difficult to match them exactly:
Changes:
- stretched upper stage
- more detail on booster attachment
- grid fins
And a larger interstage.
Cheers, Martin
No C/F With C/FGood work OldAtlas.
FHRv1.1FHv1.1FHRv1.2FHv1.2FHRv1.1FHv1.1FHRv1.2FHv1.2
36 44.3 41 50.5 43 53 49 60.5
These values although very rough estimate numbers shows the relationships between the original FHv1.1 versions capabilities and the newer FHv1.2 versions capabilities. [Versions are [R][C/F:woC/F], four in all for each v.
The R versions assume the same landing of booster RTLS and center core ASDS landing downrange.
The conclusion is that unless you need something with a capability of > 49mt then a reusable configuration would work just fine. Even woC/F you get about 41mt reusable.
nadreck
You are probably correct in that there may never be a plain FHC/F. If or when they get RTLS working FH is unlikely to ever fly a non -reusable configuration after that even if only the boosters are recovered. A C/F where the center core is expended would cost $25M more or a price as low as $85M. Those values of $90M in the NASA Moon return study for FH usage and costs starts to look like they may be closer to correct than we think. Maybe they got some probable pricing data from SpaceX for the FH2Rv1.2C/F configuration that could do what they specified for TLI.
nadreck
You are probably correct in that there may never be a plain FHC/F. If or when they get RTLS working FH is unlikely to ever fly a non -reusable configuration after that even if only the boosters are recovered. A C/F where the center core is expended would cost $25M more or a price as low as $85M. Those values of $90M in the NASA Moon return study for FH usage and costs starts to look like they may be closer to correct than we think. Maybe they got some probable pricing data from SpaceX for the FH2Rv1.2C/F configuration that could do what they specified for TLI.
Why would they expend the center core, ever? Barge landings, if they're done without boost-back, cost very little propellant and have to be cheaper than $25M... Or am I reading your logic wrong?
Interesting calculations. I keep wondering how much FH non cf with disposable central core could send to Mars. That might be something SpaceX wants to do and I would like to know if a Red Dragon can be sent to Mars expending only the central core. Red Dragon might be 10t.
nadreck
You are probably correct in that there may never be a plain FHC/F. If or when they get RTLS working FH is unlikely to ever fly a non -reusable configuration after that even if only the boosters are recovered. A C/F where the center core is expended would cost $25M more or a price as low as $85M. Those values of $90M in the NASA Moon return study for FH usage and costs starts to look like they may be closer to correct than we think. Maybe they got some probable pricing data from SpaceX for the FH2Rv1.2C/F configuration that could do what they specified for TLI.
Why would they expend the center core, ever? Barge landings, if they're done without boost-back, cost very little propellant and have to be cheaper than $25M... Or am I reading your logic wrong?
From my perspective I don't think they will ever find a reason to do C/F at a $25M price tag. Throwing away the centre core without C/F represents at most a $40M cost (more likely $10 to $30M based on it being a used core), and it more than doubles GSO and higher energy payloads, and more than a 50% increase to GTO over a FH3R (downrange recovery of centre core). If C/F costs $25M more per flight than non C/F but you try to use it on a centre core that is recovered down range, you are limited to a centre core flight profile that has enough fuel to boost the centre core back from >6km/s to re-enter and land on a very down range ASDS to have as much performance to GTO, GSO and beyond as a disposable centre core NON C/F flight profile and at best it saves you $15M.
the logic is that if you don't recover the cetral core you get more payload to GEO; in fact 50% more, so 50% payload more to GEO is well worthy 20M.
the logic is that if you don't recover the cetral core you get more payload to GEO; in fact 50% more, so 50% payload more to GEO is well worthy 20M.
For that to be true, you need to show that recovering the center core downrange costs 33% of the payload.
I thought that was the number when considering RTLS, and even that was deemed acceptable.
You need to look at cost, since that's the ultimate driver.
If the other two cores are reused, an expendable center core becomes the most expensive component in the stack, and so you're highly motivated to reuse it too - it's much more than "one core out of three". Luckily, they have down-range recovery capabilities, which has a lower payload penalty than RTLS, so I can't see why they'd refrain from using it.
nadreck
You are probably correct in that there may never be a plain FHC/F. If or when they get RTLS working FH is unlikely to ever fly a non -reusable configuration after that even if only the boosters are recovered. A C/F where the center core is expended would cost $25M more or a price as low as $85M. Those values of $90M in the NASA Moon return study for FH usage and costs starts to look like they may be closer to correct than we think. Maybe they got some probable pricing data from SpaceX for the FH2Rv1.2C/F configuration that could do what they specified for TLI.
Why would they expend the center core, ever? Barge landings, if they're done without boost-back, cost very little propellant and have to be cheaper than $25M... Or am I reading your logic wrong?
From my perspective I don't think they will ever find a reason to do C/F at a $25M price tag. Throwing away the centre core without C/F represents at most a $40M cost (more likely $10 to $30M based on it being a used core), and it more than doubles GSO and higher energy payloads, and more than a 50% increase to GTO over a FH3R (downrange recovery of centre core). If C/F costs $25M more per flight than non C/F but you try to use it on a centre core that is recovered down range, you are limited to a centre core flight profile that has enough fuel to boost the centre core back from >6km/s to re-enter and land on a very down range ASDS to have as much performance to GTO, GSO and beyond as a disposable centre core NON C/F flight profile and at best it saves you $15M.
I'm not following you - I might have missed some logic upthread.
Where is the $25M coming from? I assumed it is the write-off cost of the center core if you expend it.
If you recover down-range, you save the $25M, and your costs are the barge ops and the re-entry propellant, which might be a few km/s dV, but on an otherwise empty stage.
They might not have a customer now, but if they'll use FH for refueling MCTs, it'll make sense to get as much performance out of them as possible.
the logic is that if you don't recover the cetral core you get more payload to GEO; in fact 50% more, so 50% payload more to GEO is well worthy 20M.
For that to be true, you need to show that recovering the center core downrange costs 33% of the payload.
I thought that was the number when considering RTLS, and even that was deemed acceptable.
You need to look at cost, since that's the ultimate driver.
If the other two cores are reused, an expendable center core becomes the most expensive component in the stack, and so you're highly motivated to reuse it too - it's much more than "one core out of three". Luckily, they have down-range recovery capabilities, which has a lower payload penalty than RTLS, so I can't see why they'd refrain from using it.
No it costs more like 50% of the payload beyond GTO, it costs much less against the LEO payload figures. If people have time to wait, I will put up a break down of based on the assumptions I was using to calculate from either tomorrow or sometime over the weekend.nadreck
You are probably correct in that there may never be a plain FHC/F. If or when they get RTLS working FH is unlikely to ever fly a non -reusable configuration after that even if only the boosters are recovered. A C/F where the center core is expended would cost $25M more or a price as low as $85M. Those values of $90M in the NASA Moon return study for FH usage and costs starts to look like they may be closer to correct than we think. Maybe they got some probable pricing data from SpaceX for the FH2Rv1.2C/F configuration that could do what they specified for TLI.
Why would they expend the center core, ever? Barge landings, if they're done without boost-back, cost very little propellant and have to be cheaper than $25M... Or am I reading your logic wrong?
From my perspective I don't think they will ever find a reason to do C/F at a $25M price tag. Throwing away the centre core without C/F represents at most a $40M cost (more likely $10 to $30M based on it being a used core), and it more than doubles GSO and higher energy payloads, and more than a 50% increase to GTO over a FH3R (downrange recovery of centre core). If C/F costs $25M more per flight than non C/F but you try to use it on a centre core that is recovered down range, you are limited to a centre core flight profile that has enough fuel to boost the centre core back from >6km/s to re-enter and land on a very down range ASDS to have as much performance to GTO, GSO and beyond as a disposable centre core NON C/F flight profile and at best it saves you $15M.
I'm not following you - I might have missed some logic upthread.
Where is the $25M coming from? I assumed it is the write-off cost of the center core if you expend it.
If you recover down-range, you save the $25M, and your costs are the barge ops and the re-entry propellant, which might be a few km/s dV, but on an otherwise empty stage.
They might not have a customer now, but if they'll use FH for refueling MCTs, it'll make sense to get as much performance out of them as possible.
I was taking the $25M OldAtlasGuy quoted for making C/F happen on a per flight basis. I consider that a throwaway of a centre core costs at the high end the cost of a centre core new at the high end less the cost of refurbishing and relaunching a used centre core (so about $30-$40M at the top end) but if it was a near end of life core maybe it is only a $10M cost.
Interesting calculations. I keep wondering how much FH non cf with disposable central core could send to Mars. That might be something SpaceX wants to do and I would like to know if a Red Dragon can be sent to Mars expending only the central core. Red Dragon might be 10t.
From their website 13,200kg
SpaceX has updated the renderings on their F9/FH pages with the "v1.2" renderings, which are more detailed, and presumably accurate. It does seem to indicate a ~6ft stretch as well in the upper stage.
Here is the old (left) FH rendering compared to the new one (right), but it is difficult to match them exactly:
Changes:
- stretched upper stage
- more detail on booster attachment
- grid fins
And a larger interstage.
Cheers, Martin
It indeed does looks a bit longer.
Also, the leg's seem shorter and the center engine alignment is different. This could be just the rendering ofcourse.
it looks to me like the previous render was displayed in parallel projection, while the new one is displayed in perspective from a long viewing distance. i think thats whats making the center engine look different.
the logic is that if you don't recover the cetral core you get more payload to GEO; in fact 50% more, so 50% payload more to GEO is well worthy 20M.
For that to be true, you need to show that recovering the center core downrange costs 33% of the payload.
I thought that was the number when considering RTLS, and even that was deemed acceptable.
You need to look at cost, since that's the ultimate driver.
If the other two cores are reused, an expendable center core becomes the most expensive component in the stack, and so you're highly motivated to reuse it too - it's much more than "one core out of three". Luckily, they have down-range recovery capabilities, which has a lower payload penalty than RTLS, so I can't see why they'd refrain from using it.
No it costs more like 50% of the payload beyond GTO, it costs much less against the LEO payload figures. If people have time to wait, I will put up a break down of based on the assumptions I was using to calculate from either tomorrow or sometime over the weekend.
i just thought about this again and perhaps they cut and pasted the engines from one render onto a different render. the stages are too perfectly straight to be perspective. if it was parallel at an angle the horizontal lines between stages would be curved.
So if we presume that we are presuming a 200 KM LEO orbit and that GTO is a 2450m/s impulse past that then here are the following capabilities of an FH either with side boosters RTLS and centre core recovered down range:
25t to LEO with centre core landing down range and side cores RTLS
40t to LEO with centre core expended and side cores landing down range
8t to GTO with centre core landing down range and side cores RTLS
17t to GTO with centre core expended and side cores landing down range
So if we presume that we are presuming a 200 KM LEO orbit and that GTO is a 2450m/s impulse past that then here are the following capabilities of an FH either with side boosters RTLS and centre core recovered down range:
25t to LEO with centre core landing down range and side cores RTLS
40t to LEO with centre core expended and side cores landing down range
8t to GTO with centre core landing down range and side cores RTLS
17t to GTO with centre core expended and side cores landing down range
So first, you're double presuming up there, and I don't know if that's legal.
Less importantly, this is not an apples to apples comparison. You're evaluating the two center core options (expend it or recover downrange) under two different scenarios (side core RTLS and side core down-range recovery)
The question was, suppose we RTLS the side cores, what is the payload hit for recovering the center core deep downrange as opposed to expending it.
You will need to make some assumptions on the amount of slow-down necessary before re-entry, and we don't quite know what that is.
My understanding is that they were able to get by with a rather minimal braking burn on a regular F9 launch, and so I would guestimate that they need to slow down from center-core speed (which we don't quite know either) to just under single-core speed. (which we should have a better guess at)
For Guckyfan's question on TMI presuming that TMI is accomplished with an impulse of 1.3km/s more than GTO:
3t to TMI with centre core landing down range and side cores RTLS
9t to TMI with centre core expended and side cores landing down range
For Guckyfan's question on TMI presuming that TMI is accomplished with an impulse of 1.3km/s more than GTO:
3t to TMI with centre core landing down range and side cores RTLS
9t to TMI with centre core expended and side cores landing down range
Thanks a lot. Sadly this seems to indicate that Red Dragon needs a fully expended FH which is still quite reasonable for a NASA mission but expensive for a mission self funded by SpaceX.
Note: side cores downrange, is that even an option? It would need two ASDS or possibly a new landing site in Florida assuming launch from Texas.
Have you made some calculations for a payload to the moon? I'm asking this because I was thinking about something I read a wile ago about the possibility of SpaceX launching a Dragon capsule (like CRS-5 one) for a free ride around the moon and then to retrieve it in the Pacific Ocean on the inaugural launch of the Falcon Heavy next spring.For Guckyfan's question on TMI presuming that TMI is accomplished with an impulse of 1.3km/s more than GTO:
3t to TMI with centre core landing down range and side cores RTLS
9t to TMI with centre core expended and side cores landing down range
Thanks a lot. Sadly this seems to indicate that Red Dragon needs a fully expended FH which is still quite reasonable for a NASA mission but expensive for a mission self funded by SpaceX.
Note: side cores downrange, is that even an option? It would need two ASDS or possibly a new landing site in Florida assuming launch from Texas.
Side cores won't have gone all that far, and given that there seem to be more than 2 ASDS's in the Atlanatic (at least as far as I recall from the very long winded ASDS thread) I think they could be recovered handily. Also expending the side cores would add between 400 and 500m/s which doesn't sound like much but would push the TMI number up past 11, however Full Thrust and densification might do as much, put the two together maybe you get 13. Remember I used vanilla V1.1 stats to build these tables yet after Jason-3 there will be no more vanilla V1.1.
The question was, suppose we RTLS the side cores, what is the payload hit for recovering the center core deep downrange as opposed to expending it.
So if we presume that we are presuming a 200 KM LEO orbit and that GTO is a 2450m/s impulse past that then here are the following capabilities of an FH either with side boosters RTLS and centre core recovered down range:
25t to LEO with centre core landing down range and side cores RTLS
40t to LEO with centre core expended and side cores landing down range
8t to GTO with centre core landing down range and side cores RTLS
17t to GTO with centre core expended and side cores landing down range
So first, you're double presuming up there, and I don't know if that's legal.
Less importantly, this is not an apples to apples comparison. You're evaluating the two center core options (expend it or recover downrange) under two different scenarios (side core RTLS and side core down-range recovery)
The question was, suppose we RTLS the side cores, what is the payload hit for recovering the center core deep downrange as opposed to expending it.
You will need to make some assumptions on the amount of slow-down necessary before re-entry, and we don't quite know what that is.
My understanding is that they were able to get by with a rather minimal braking burn on a regular F9 launch, and so I would guestimate that they need to slow down from center-core speed (which we don't quite know either) to just under single-core speed. (which we should have a better guess at)
I assumed slowing down to 900m/s and using 400m/s of delta V to land that is all in my spread sheets, if you don't read my spreadsheets and assumptions then you are simply left with taking my conclusions with out any reasoning behind it and can't criticize it.
You're supposed to compare expend/recover of center core under identical assumptions.
Instead, you compared:
- recover-center-core with side-core-RTLS, to
- expend-center-core with side-core-downrange
Which inflates the difference in payload to orbit (any orbit) - the "expend" version benefits from not having to RTLS the side cores.
So here is the updated spreadsheet.
I added a 2nd column on the tables in the {Centre core performance} worksheet for 100% remaining propellant for the cross feed case, and I added another table allowing for 20% propellant reserve for recovering the centre core in that case. I also added another worksheet for the 3 booster boost phase in the cross feed scenario.
So with the cross feed and side boosters RTLS'ing and centre core recovery down range
The payload to GTO would be 12t
with cross feed and side boosters RTLS'ing and centre core expending the payload to GTO is 20t
GTO performance with side cores recovered down range and centre core expended is 25t
TMI under those same 3 cases is, respectively, 5t, 12t and 14t
REMEMBER this is V1.1 legacy specs not the new full thrust.
ALSO note that my spreadsheet does not calculated several factors, that while some cancel others, it is still only an approximation and a better indication of relative performance rather than absolute. On the underestimating side it presumes sea level ISP for the 3 core boost phase, also when coming up with my delta V budgets for the scenarios I ignored the benefit of launching east from the cape. On the other side of the ledger I didn't allow for air resistance and while I accounted for gravity loss in the 3 core boost phase I ignored it after that and while it is minimal after that it is not zero.
Aviation Week has a new item on Falcon Heavy. Lists launch customers. First launch is SpaceX funded test.
http://aviationweek.com/space/spacex-introduce-falcon-heavy-early-2016-0?NL=AW-19&Issue=AW-19_20150909_AW-19_329&sfvc4enews=42&cl=article_4&utm_rid=CPEN1000000903672&utm_campaign=3734&utm_medium=email&elq2=0bc134d29fc64b68921dade1f49f5f80
Aviation Week has a new item on Falcon Heavy. Lists launch customers. First launch is SpaceX funded test.
http://aviationweek.com/space/spacex-introduce-falcon-heavy-early-2016-0?NL=AW-19&Issue=AW-19_20150909_AW-19_329&sfvc4enews=42&cl=article_4&utm_rid=CPEN1000000903672&utm_campaign=3734&utm_medium=email&elq2=0bc134d29fc64b68921dade1f49f5f80
Peter B. de Selding @pbdes
ViaSat: Our ViaSat-2 sat to launch on Falcon Heavy in Q4 2016, after 1st Falcon Heavy ~ May. Plan B, an Ariane 5, would cost time & money.
https://twitter.com/pbdes/status/641654264918634496
That ASDS thread was really long winded. :)
For a short time it was thought there would be 3 ASDS. But we know now that Marmac 300 was decomissioned as an ASDS so there are two, one Atlantic, one Pacific. They would have to build a new one for sidebooster recovery.
I missed, that the numbers were 1.1. So payload to Mars with booster recovery is still possible. According to the Red Dragon video it would require ~11t to TMI for a full 2t payload to the surface. Which is not a bad payload fraction. It is similar to the payload fraction of Curiosity.
For Guckyfan's question on TMI presuming that TMI is accomplished with an impulse of 1.3km/s more than GTO:
3t to TMI with centre core landing down range and side cores RTLS
9t to TMI with centre core expended and side cores landing down range
Thanks a lot. Sadly this seems to indicate that Red Dragon needs a fully expended FH which is still quite reasonable for a NASA mission but expensive for a mission self funded by SpaceX.
Note: side cores downrange, is that even an option? It would need two ASDS or possibly a new landing site in Florida assuming launch from Texas.
Side cores won't have gone all that far, and given that there seem to be more than 2 ASDS's in the Atlanatic (at least as far as I recall from the very long winded ASDS thread) I think they could be recovered handily. Also expending the side cores would add between 400 and 500m/s which doesn't sound like much but would push the TMI number up past 11, however Full Thrust and densification might do as much, put the two together maybe you get 13. Remember I used vanilla V1.1 stats to build these tables yet after Jason-3 there will be no more vanilla V1.1.
Methinks a new, fully reusable, methane fueled launcher powered by Raptors or Raptor variants will replace both F9 and FH. Larger payloads migrate to BFR.
Methinks a new, fully reusable, methane fueled launcher powered by Raptors or Raptor variants will replace both F9 and FH. Larger payloads migrate to BFR.
Where does that leave the loads of progressively smaller satellites they would need to launch to make money?
Methinks a new, fully reusable, methane fueled launcher powered by Raptors or Raptor variants will replace both F9 and FH. Larger payloads migrate to BFR.
Where does that leave the loads of progressively smaller satellites they would need to launch to make money?
You launch in bulk. To use an Earth based delivery metaphor: Does it make more sense to ship a lot of small packages in a single truck, or using individual deliveries with motorcycles?
Methinks a new, fully reusable, methane fueled launcher powered by Raptors or Raptor variants will replace both F9 and FH. Larger payloads migrate to BFR.
Where does that leave the loads of progressively smaller satellites they would need to launch to make money?
Feel stupid asking this on an image about guessing about future rockets, but has SX stated plans for a Super Falcon (quad core) and calling all BFRs Eagles?
A quad core will mean new or massively changed assembly buildings, strong backs, and launch sites.
Methinks a new, fully reusable, methane fueled launcher powered by Raptors or Raptor variants will replace both F9 and FH. Larger payloads migrate to BFR.
I used to think the same, an intermediate sized LV between the FHR and the BFR. The only problem from my perspective, is the delays and costs involved. How Quickly does SpaceX want to get moving on it's Mars Project? Some early launches of an upgraded FH around 2018-20 is possible for cargo and exploration satellites and practice at launching to Mars; but then there is the cash flow to follow on with the BFR lite or BFR full. How soon can Musk get his Internet Constellation paying for itself? What are the trade offs and the benefits in the short term launch goals for 2020-25? Seems to me we are missing a few pieces of the puzzle.
Would anyone suggest a Raptor powered FHR Shell and replace Dracos/Super Dracos on the Dragon with suitably sized mini-raptors. Simply as a proof of concept test vehicle. Seems crazy just thinking about it. The real question is, how crazy are Musk and the engineers at SpaceX. What are they working on behind those curtains?
YES I know (Jim) that LV/SC are not Lego Blocks. So don't disrupt the electrons with a retort ;-)
No, that image I posted is total fiction.Probably shouldn't put the SpaceX logo on it, then, lest some hapless journalist mistakes it for a SpaceX publication.
I saw an interesting discussion elsewhere on the internet and I'd like to ask you guys a few things.
I'll start off with this premise: In the future, SpaceX is replacing the Falcon Heavy. What do you think is most likely, and why?
1. Falcon Heavy with improved Merlins, even methalox Merlins of some design
2. Falcon Heavy with Raptor engines
3. Single core stick in between whatever Falcon 9 they are using at the time and the BFR in size
Which is the most economical to use? Which is the easiest to build? Which is the best performer?
No, that image I posted is total fiction.Probably shouldn't put the SpaceX logo on it, then, lest some hapless journalist mistakes it for a SpaceX publication.
You launch in bulk. To use an Earth based delivery metaphor: Does it make more sense to ship a lot of small packages in a single truck, or using individual deliveries with motorcycles?
You launch in bulk. To use an Earth based delivery metaphor: Does it make more sense to ship a lot of small packages in a single truck, or using individual deliveries with motorcycles?
Well, Bezos wants to deliver via small flying drones.
Would anyone suggest a Raptor powered FHR Shell and replace Dracos/Super Dracos on the Dragon with suitably sized mini-raptors. Simply as a proof of concept test vehicle. Seems crazy just thinking about it. The real question is, how crazy are Musk and the engineers at SpaceX. What are they working on behind those curtains?
YES I know (Jim) that LV/SC are not Lego Blocks. So don't disrupt the electrons with a retort ;-)
An you need a special ignition system with some another type of fuel to start the methane engine. You cannot restart that hundreds of times without huge amount of the starting fuel. And then you don't need the methane anymore/does not make sense to have the complexity to also to be able to burn also methane.
Methane is natural gas. 50% of homes in America have it for heating, cooking, water heating, and cloths drying. It is lit and relit 100's if not thousands of time to preform these functions. Just because it is under pressure going into a combustion chamber doesn't mean it cant be lit or relit. It doesn't clog up fuel lines and such like kerosene might over time. It doesn't coke, or very, very little. It is abundant, and is liquefied all over the country for storage during the summer to be released in winter when it is used most. It is stored in what looks like giant ground level water tanks or oil tanks. They are built like a vacuum bottle with walls about 3' or 1m apart. The space in between is vacuumed out to keep from having heat cause boil off. The giant tanks are about 200-300' in diameter.
Methinks a new, fully reusable, methane fueled launcher powered by Raptors or Raptor variants will replace both F9 and FH. Larger payloads migrate to BFR.
Methinks a new, fully reusable, methane fueled launcher powered by Raptors or Raptor variants will replace both F9 and FH. Larger payloads migrate to BFR.
Where does that leave the loads of progressively smaller satellites they would need to launch to make money?
You launch in bulk. To use an Earth based delivery metaphor: Does it make more sense to ship a lot of small packages in a single truck, or using individual deliveries with motorcycles?
1% is odor put in to identify leaks because pure methane is odorless, another 1 or 2 percent is various other flammable gasses like butane and ethane. It would be very hard and expensive to get it pure, with centrifuges. Probably not worth the effort. Making it liquid will probably filter out a lot of the butane and ethane.
Even K1 is probably not 100% K1, it is a refined fuel. At some point cost to make fuel completely pure outweigh 98%
I was an engineer with a natural gas utility for 39 years. We maintained a .9-1.2% odor concentration, not parts per million. Sorry 1% is required by the Department of Transportation, Office of Pipeline Safety rules in parts 192 and parts 193 that we maintain .9-1/2% odorant in distribution natural gas.
It is just not economically feasible to get everything out. We maintained 1000 btu's per 1 cubic foot as much as possible at 6" water column pressure. Some was less, some more, but not by much. Adjustments were made in billing to compensate. Most fracked natural gas, natural gas in wells without oil and in coal seams is not so hard to separate. When it all boils down, most natural gas nationwide is about 96-97% pure methane.
Another thing, methane is liquid -161 C, while nitrogen is -196 C, so when making liquid methane from natural gas, the nitrogen will be released that was in the natural gas. I think people are splitting hairs trying to get from 97 or 98 percent methane any further. They will liquefy the gas straight from the nearest pipeline at the Launchpad. It will not be trucked in, or a liquification tanker ship if the BFR Launchpad is at or near a seacoast.
It would be opportunistic of SpaceX and all others interested in CH4 for ISRU considerations, to source part of the methane from providers who can demonstrate technical and economic competence at using the Sabatier reaction to produce pure methane.
I realize it would probably be cheaper to just buy O2 and liquify existing natural gas. But where is the fun and coolness in that?
An you need a special ignition system with some another type of fuel to start the methane engine. You cannot restart that hundreds of times without huge amount of the starting fuel. And then you don't need the methane anymore/does not make sense to have the complexity to also to be able to burn also methane.
While tricky, I think you are VASTLY overstating the combustion problem. It is in fact possible to ignite hydrocarbon fuels hundreds, thousands, or more times without special ignition fuel. Doesn't your car do it every time you start it?
Reliable microgravity start is of course very difficult, but not an unsolvable problem.
If on the day SpX closed the deal on LC-39A (14 April 2014) they had installed a half million dollar photovoltaic Sabatier reactor, and ran it non stop at full speed, would they have enough propellant for a BFR launch in 2020?
While tricky, I think you are VASTLY overstating the combustion problem. It is in fact possible to ignite hydrocarbon fuels hundreds, thousands, or more times without special ignition fuel. Doesn't your car do it every time you start it?
Reliable microgravity start is of course very difficult, but not an unsolvable problem.
My car takes about on second to start, and coupl of seconds more until I get full thrust or precise thrust control from it.
That's unacceptable for thruster or emergency exit engine that either is only used for ~0.1 second at a time for precise manouvering(like docking to a space station), or one that has to reach full thrust instantly to escape exploding rocket.
If on the day SpX closed the deal on LC-39A (14 April 2014) they had installed a half million dollar photovoltaic Sabatier reactor, and ran it non stop at full speed, would they have enough propellant for a BFR launch in 2020?
Fortunately a BFR is not needed on Mars. MCT is big but miniscule compared to BFR so fuel requirement is miniscule too.
Producing methane at the pad feels like a stunt to me. Only the sabatier reactor and electrolysis for hydrogen would be similar. Solar panels, water production and CO2 production are the critical technologies and those would be very different on Mars.
It would be a fun exercise to go through the process and energy flow diagrams for such a design.
A question. After the latest known info about the merlin engine upgrades, has one tried to calculate what would the FH payload capability be for LEO? The site still lists the older numbers.
Is there anywhere on NSF or elsewhere that outlines what a Mars mission might look like using FHs to assemble something in orbit/send supplies?
Is there anywhere on NSF or elsewhere that outlines what a Mars mission might look like using FHs to assemble something in orbit/send supplies?
Start here: http://forum.nasaspaceflight.com/index.php?topic=38506.0 (http://forum.nasaspaceflight.com/index.php?topic=38506.0)
Does anyone know if the FH Boosters will be interchangeable or are they right and left.
Same goes for the center core is it always going to be a center core and can any FH cores be used on an F9?
Does anyone know if the FH Boosters will be interchangeable or are they right and left.Note that that engine layout schematic is inconsistent with the model they use for their renders and video where the right booster is rotated 180 degrees so that all attachment points are in the same locations (and presumably identical) for both boosters.
Same goes for the center core is it always going to be a center core and can any FH cores be used on an F9?
AFAIK Center Cores are unique specific to heavy (extra structural reinforcement to handle all that oomph, I guess) while side boosters are effectively standard F9s with a pretty nose cap.
Does anyone know if the FH Boosters will be interchangeable or are they right and left.
Same goes for the center core is it always going to be a center core and can any FH cores be used on an F9?
AFAIK Center Cores are unique specific to heavy (extra structural reinforcement to handle all that oomph, I guess) while side boosters are effectively standard F9s with a pretty nose cap.
If the FH boosters are not mirror imaged, how do they handle umbilicals and fueling?
The graphic on the SpaceX website shows boosters that are anything but mirrored.
If they were mirrored we would see identical exterior features on each booster.
Since they are different-- all we can say is they are not mirrored. They may be identical but we don't know.
Is there anywhere on NSF or elsewhere that outlines what a Mars mission might look like using FHs to assemble something in orbit/send supplies?
Start here: http://forum.nasaspaceflight.com/index.php?topic=38506.0 (http://forum.nasaspaceflight.com/index.php?topic=38506.0)
I wasn't expecting that much of a discussion, that is fantastic! Thanks!
Want to bring this thread back for the new year, perhaps the year of FH.
Let's say there are a handful of landed stages by the time FH is ready to fly (likely), and at least one has been reflown successfully (maybe). Do you think recovered first stages can undergo whatever modifications necessary for use as side boosters (minor, as claimed)? If so, will FH demo flight use reflown boosters for cost savings?
One argument against is that part of developing FH is building at least one, including side boosters, from scratch. But that's irrelevant if they're nearly identical to first stages. Another is that perhaps reflight will have been shown, but not enough times to prove reliable for the high-profile FH demo.
Want to bring this thread back for the new year, perhaps the year of FH.
[1] You'll save money by just using the recovered stages unaltered for more F9 flights.
[2] they frankly can't afford a potential new problem to show up on this already-complicated vehicle.
Methinks a new, fully reusable, methane fueled launcher powered by Raptors or Raptor variants will replace both F9 and FH. Larger payloads migrate to BFR.
Where does that leave the loads of progressively smaller satellites they would need to launch to make money?
The bread and butter of the launch market is still GTO and ISS destinations. GTO launches can be aggregated easily, can't them ? Plus perhaps hurried launches to GTO could be delivered directly to GEO or near GEO ? How much would that be worth in extra launch value ?
Something I've been puzzling over: Are there any differences, structurally, between the F9 v.1.2 core and the FH outboards? I'm wondering if one of the markets for reused cores might be to reduce FH production costs.