I thought the vacuum ISP was supposed to be 380 seconds.
Quote from: llanitedave on 01/21/2015 10:29 pmI thought the vacuum ISP was supposed to be 380 seconds.Maybe?as of http://www.nasaspaceflight.com/2014/03/spacex-advances-drive-mars-rocket-raptor-power/ it was 363/321 at 1Mlbf, but later..."A June 2014 talk by Mueller provided more specific engine performance target specifications indicating 6,900 kN (705 tonnes-force) of sea-level thrust, 8,200 kN (840 tonnes-force) of vacuum thrust, and a specific impulse of 380 s for a vacuum version.[1] Earlier information had estimated the design Isp under vacuum conditions as only 363 s.[2]"
MCT will have meaningfully higher specific impulse engines: 380 vs 345 vac Isp. For those unfamiliar, in the rocket world, that is a super gigantic difference for stages of roughly equivalent mass ratio (mass full to mass empty).
Quote from: Burninate on 01/21/2015 10:34 pmQuote from: llanitedave on 01/21/2015 10:29 pmI thought the vacuum ISP was supposed to be 380 seconds.Maybe?as of http://www.nasaspaceflight.com/2014/03/spacex-advances-drive-mars-rocket-raptor-power/ it was 363/321 at 1Mlbf, but later..."A June 2014 talk by Mueller provided more specific engine performance target specifications indicating 6,900 kN (705 tonnes-force) of sea-level thrust, 8,200 kN (840 tonnes-force) of vacuum thrust, and a specific impulse of 380 s for a vacuum version.[1] Earlier information had estimated the design Isp under vacuum conditions as only 363 s.[2]"From the Redditt AMA Elon Musk:QuoteMCT will have meaningfully higher specific impulse engines: 380 vs 345 vac Isp. For those unfamiliar, in the rocket world, that is a super gigantic difference for stages of roughly equivalent mass ratio (mass full to mass empty).
Okay, thanks, that mystery is solved - must have missed that comment. What nozzle diameter is required for a practical reusable FFSC CH4-LOX 500klbf engine, in vac and at SL?
Okay, thanks, that mystery is solved - must have missed that comment.
Quote from: Burninate on 01/21/2015 11:41 pmOkay, thanks, that mystery is solved - must have missed that comment. What nozzle diameter is required for a practical reusable FFSC CH4-LOX 500klbf engine, in vac and at SL?No idea. I don't think that anybody outside of SpaceX knows.As a crude estimate, the M1C had a SL thrust of 161klbf and a bell diameter of about 1m. The Raptor:M1D thrust ratio would be 3.1056:1, and given an equivalent thrust:area ratio the SL Raptor engine bell would be about 1.76m in diameter.
Quote from: malu5531 on 12/12/2013 12:37 pmWhile trying to understand Merlin 1D and in particular "Merlin 1D+"* in depth, I've iterated my calculations a few times and have reached internal coherence and good balance with reality using the following characteristics/specs. Merlin 1D..Merlin 1D Vac..Merlin 1D+..Merlin 1D+ VacNozzle diameter, m1.073.031.073.03I've tried to draw the Falcon 9 first stage with your diameter, but I can't get it to fit. A circle of 8 engines with a diameter of 1,07 m each is going to have an outer diameter of at least 3,8 m as drawn in my CAD program, and that's with the engine nozzles touching each other. But if you look at images of the launch, the engines do not protrude outside the first stage diameter. And there's a gap between the engines. If I limit the outer diameter of the 8 engines to 3,66 m and allow some spacing between them, the nozzle diameter is around 96,5 cm (my drawing was in 1:144). I've also tried measuring the diameter from the second photo. Ignoring the distortion, the space between the center engine and the outer engines is 0.452 times the diameter of the center engine. So the total diameter of the ring of 8 engines is (3 + (2*0.452)) times the diameter of one engine nozzle. If the total ring diameter is 366 cm, then one engine must be 93,5 cm in diameter.
While trying to understand Merlin 1D and in particular "Merlin 1D+"* in depth, I've iterated my calculations a few times and have reached internal coherence and good balance with reality using the following characteristics/specs. Merlin 1D..Merlin 1D Vac..Merlin 1D+..Merlin 1D+ VacNozzle diameter, m1.073.031.073.03
Quote from: malu5531 on 12/31/2013 01:39 amI fully agree, nozzle diameter should be something like ~93-94 cm. I'm not fully there with the model. I have a similar issue with the RD-0162, not sure yet if those are related or there is a more trivial problem with the 1D model (such as adjusting the chamber pressure a bit, since that info might be old). Quote from: malu5531 on 12/31/2013 01:39 amI fully agree, nozzle diameter should be something like ~93-94 cm. I'm not fully there with the model. I have a similar issue with the RD-0162, not sure yet if those are related or there is a more trivial problem with the 1D model (such as adjusting the chamber pressure a bit, since that info might be old).Chamber pressure need not enter into it In the 1D model*. Thrust is F = q ve + Ae (pe - pa) ,where pa is the ambient pressure. Therefore the difference between sea-level (pa = pSL = 1 atm) thrust and vacuum (pa = 0) thrust is Fvac - FSL = Ae pSL ,so Ae = (Fvac - FSL) / pSL .If we take the total thrust of the Falcon 9's first stage at sea level and in vacuo from the Falcon 9 web page and divide by nine, we get single-engine thrusts of 653.9 and 741.3 kN, respectively. For pSL = 101.325 kPa, we get Ae = 0.863 m2 and hence a diameter of 1.048 m, which agrees closely with the value you've calculated. On the other hand, if we look at the web page for the Merlin engine itself, we're told that the engine's vacuum thrust is just 716 kN. This lower thrust gives an exit area of 0.613 m2 and a diameter of just 0.883 m. This fits within the geometric limit found by Hobbes-22.Now, it could be that this simple analysis violates some constraints imposed by your more extensive model. I would think, though (and please correct me if I'm wrong), that the the other constraints have are pretty loose, given SpaceX's reluctance to give engineering specifics.* Actually, I suppose that's not strictly true. If we're going to assume that flow separation occurs once the pressure drops more than a certain amount below ambient, then the effective nozzle area would depend on chamber pressure. Thus far, though, we've been assuming there's no flow separation. As far as I know (which isn't very far), flow separation is usually avoided these days. (The sustainer of the classic Atlas was over-expanded at sea level to the point that flow separation did occur, but that was back when men were men .) Anyway, allowing for flow separation would tend to increase our estimate of the nozzle's size. Since we're pretty close to the size allowed by geometry already, this suggests that separation does not occur.
I fully agree, nozzle diameter should be something like ~93-94 cm. I'm not fully there with the model. I have a similar issue with the RD-0162, not sure yet if those are related or there is a more trivial problem with the 1D model (such as adjusting the chamber pressure a bit, since that info might be old). Quote from: malu5531 on 12/31/2013 01:39 amI fully agree, nozzle diameter should be something like ~93-94 cm. I'm not fully there with the model. I have a similar issue with the RD-0162, not sure yet if those are related or there is a more trivial problem with the 1D model (such as adjusting the chamber pressure a bit, since that info might be old).
I fully agree, nozzle diameter should be something like ~93-94 cm. I'm not fully there with the model. I have a similar issue with the RD-0162, not sure yet if those are related or there is a more trivial problem with the 1D model (such as adjusting the chamber pressure a bit, since that info might be old).
At 1.76m engine bells, looks like you can fit 19 of them comfortably on a 10m diameter stage. Think that would be enough for a start?
As a crude estimate, the M1C had a SL thrust of 161klbf and a bell diameter of about 1m. The Raptor:M1D thrust ratio would be 3.1056:1, and given an equivalent thrust:area ratio the SL Raptor engine bell would be about 1.76m in diameter.
15m would be the limit
You need gimballing space:
Quote from: BobCarver on 01/22/2015 02:59 amYou need gimballing space:Do the outer 8 engines gimbal in 2 dimensions each, or only one?
Quote from: BobCarver on 01/22/2015 02:59 amYou need gimballing space:With that many engines, some could be fixed and others able to gimbal. With fast computers and advanced electronic accelerometers and gyros, all the thrust vectoring could be done by only a few of the engines.You also have the option of skirt fairings.
They could also try it like the russians/sovjets did with the N1
Ah yes.Given the stellar fail/success rate of the N1, we can assume that this technique to be robust and reliable, right?
Quote from: Pete on 01/22/2015 07:30 amAh yes.Given the stellar fail/success rate of the N1, we can assume that this technique to be robust and reliable, right?Weren't N1 failures related to plumbing, not gimballing?
Yes, but it was still a terrible - truly terrible - idea. For every engine you lose, you need to turn off the opposite number.
It's still not a great system. It limits the engine-out options because once again you have only a limited number of steering engines. Also as has already been said, throttle response is slower than actuators.
Quote from: malenfant on 01/22/2015 10:57 amIt's still not a great system. It limits the engine-out options because once again you have only a limited number of steering engines. Also as has already been said, throttle response is slower than actuators. I'm not saying it's great, but that there are no indications that it failed, AFAIK. So there is no reason to dismiss it based on the experience of N1.
They must not trust in space assembly and docking much at SpaceX.
Quote from: Hotblack Desiato on 01/22/2015 06:46 amThey could also try it like the russians/sovjets did with the N1Ah yes.Given the stellar fail/success rate of the N1, we can assume that this technique to be robust and reliable, right?Besides, one of the functions of gimballing is to provide roll control. That is a bit difficult to do, with fixed engines with only throttle tweaks.
Quote from: Pete on 01/22/2015 07:30 amQuote from: Hotblack Desiato on 01/22/2015 06:46 amThey could also try it like the russians/sovjets did with the N1Ah yes.Given the stellar fail/success rate of the N1, we can assume that this technique to be robust and reliable, right?Besides, one of the functions of gimballing is to provide roll control. That is a bit difficult to do, with fixed engines with only throttle tweaks.Considering that the Antares rocket and the N-1 used the same engines, I'm gonna agree with you on this.
We've given you lots of reasons, are you ignoring them?
Quote from: Eerie on 01/22/2015 11:57 amQuote from: malenfant on 01/22/2015 10:57 amIt's still not a great system. It limits the engine-out options because once again you have only a limited number of steering engines. Also as has already been said, throttle response is slower than actuators. I'm not saying it's great, but that there are no indications that it failed, AFAIK. So there is no reason to dismiss it based on the experience of N1.FWIW the third flight failed when the stage broke up due to excessive roll. Of course everything is fixable and I understand that roll control was increased after that flight.Point is it's not a solution you would choose if you had options. SpaceX have options, the most obvious of which is to increase the stage diameter.My personal opinion is that Raptor will be bigger anyway. Elon Said T/W optimised around this number. He never said they were settled on building it to this number. Trades are clearly ongoing and there are other factors beside T/W.
The most recent comments on BFR indicate that they are optimized at 500klbf per engine, and simply using lots of engines to generate sufficient thrust. I am trying to model this concept for "100 tons useful cargo to the surface of Mars", and reusable return of the lander, which is likely to represent a launch vehicle of roughly 200 tons to LEO, and a bunch of refueling missions.Raptor will be a full-flow staged combustion engine producing 500klbf per engine with (based on older comments that may no longer be reliable) a vacuum Isp of 363 seconds and a sea-level Isp of 321 seconds.What I would like to know, is what are the dimensions we should expect? What I'm having trouble with is actually fitting all those engines into the rear of the rocket; I have no context to understand how closely they should be spaced without triggering cascading failures if one explodes, or how big the engine bells should be. 10-15m seems to be the consensus on fairing diameter, but my intuitive guesses of scale (2m bell diameter with ~3-4m centerline spacing) are clearly wrong, because they indicate 15m is not big enough for this launch vehicle.*Please distinguish between centerline-to-centerline distance, diameter of the bells, and airgap distance between the outer edges of the bells, or this will get confusing.**Please distinguish between two different engines: A sea-level optimized Raptor, and a vacuum-optimized Raptor. My packing problem is with the SL raptor, the VAC raptor I have not begun modelling***Would the fact that this is a tightly packed array of engines allow for smaller nozzles?
The most recent comments on BFR indicate that they are optimized at 500klbf per engine, and simply using lots of engines to generate sufficient thrust. I am trying to model this concept for "100 tons useful cargo to the surface of Mars", and reusable return of the lander, which is likely to represent a launch vehicle of roughly 200 tons to LEO, and a bunch of refueling missions.Raptor will be a full-flow staged combustion engine producing 500klbf per engine with (based on older comments that may no longer be reliable) a vacuum Isp of 363 seconds and a sea-level Isp of 321 seconds.What I would like to know, is what are the dimensions we should expect? What I'm having trouble with is actually fitting all those engines into the rear of the rocket; I have no context to understand how closely they should be spaced without triggering cascading failures if one explodes, or how big the engine bells should be. 10-15m seems to be the consensus on fairing diameter, but my intuitive guesses of scale (2m bell diameter with ~3-4m centerline spacing) are clearly wrong, because they indicate 15m is not big enough for this launch vehicle.
*Please distinguish between centerline-to-centerline distance, diameter of the bells, and airgap distance between the outer edges of the bells, or this will get confusing.**Please distinguish between two different engines: A sea-level optimized Raptor, and a vacuum-optimized Raptor. My packing problem is with the SL raptor, the VAC raptor I have not begun modelling***Would the fact that this is a tightly packed array of engines allow for smaller nozzles?
I think you're thinking about it the wrong way. Not that long ago Elon (or someone else and SpaceX) made a comment about in orbit refueling. That coupled with the recent comments about Raptor only being 500klbs makes me thing SpaceX could be changing paradigm. They may be going away from a ginormous 12-15m LV with 15Mlbs of thrust or more, to something smaller and easier to handle. (By "smaller" I mean more Saturn V size). Engines that are smaller and easier to handle, cores that are smaller and easier to handle, etc. And then they can utilize the volume that helps to make reusability beneficial. More smaller launches with full reusability rather than fewer larger ones. At the end of the day, BFR only needs to be large enough to get a dry MCT to LEO. It can be fueled and crewed in orbit with subsequent BFR launches while prepping for a MArs mission.
...People who argue that only a few engines should gimbal seem to forget the scenario where a gimballing engine shuts down...
Why is it terrible? Russian engines are pretty reliable, aren't they?
I don't believe this for a second. MCT needs to get to orbit with crew or 100t payload. The operational complexity of transfering cargo, including possible large pieces of equipment would be huge. The only thing that will be transfered is fuel. That is my firm opinion. It does not need that huge a launch vehicle. With the second stage being MCT its weight is not fully part of what would usually be counted as payload. The engines and fuel tanks are "free".
I do realize that if you do come to a mid-air full-stop hover that your grid fins can no longer contribute to stability and control. Still, the ability to throttle down closer and closer to a hover as you approach the pad is helpful.
I keep wondering how deeply Raptor will throttle.
Quote from: TomH on 01/22/2015 08:10 pmI do realize that if you do come to a mid-air full-stop hover that your grid fins can no longer contribute to stability and control. Still, the ability to throttle down closer and closer to a hover as you approach the pad is helpful.Why? This seems to get repeated over several threads. A computer monitors all the inputs and can 1000 (or 10,000, or even 100,000) times a second have a model of exactly how far from centre its current course with current thrust takes it. It can that many times a second adjust its commands that control grid fins, gymballing, thrust, and RCS. If it stops and hovers it risks having to make larger and larger corrections (oscillating) to actually stay in place then slowly land. Hover and land slowly will be far more difficult to program, and will have lower tolerances for wind (in aviation it would be called cross wind, but for a rocket landing vertically all wind is cross wind).
Quote from: nadreck on 01/22/2015 08:22 pmQuote from: TomH on 01/22/2015 08:10 pmI do realize that if you do come to a mid-air full-stop hover that your grid fins can no longer contribute to stability and control. Still, the ability to throttle down closer and closer to a hover as you approach the pad is helpful.Why? This seems to get repeated over several threads. A computer monitors all the inputs and can 1000 (or 10,000, or even 100,000) times a second have a model of exactly how far from centre its current course with current thrust takes it. It can that many times a second adjust its commands that control grid fins, gymballing, thrust, and RCS. If it stops and hovers it risks having to make larger and larger corrections (oscillating) to actually stay in place then slowly land. Hover and land slowly will be far more difficult to program, and will have lower tolerances for wind (in aviation it would be called cross wind, but for a rocket landing vertically all wind is cross wind).What TomH is referring to is that when then stage is hovering there is no airflow through the grid fins. Without airflow the grid fins don't work.It's another reason not to hover, just come in and land.
What TomH is referring to is that when then stage is hovering there is no airflow through the grid fins. Without airflow the grid fins don't work.It's another reason not to hover, just come in and land.
Quote from: RonM on 01/22/2015 08:39 pmWhat TomH is referring to is that when then stage is hovering there is no airflow through the grid fins. Without airflow the grid fins don't work.It's another reason not to hover, just come in and land.Not at all. During the landing burn the engine will give all control authority needed. The big reason to come in and land is saving fuel.
Quote from: guckyfan on 01/22/2015 07:15 pmI don't believe this for a second. MCT needs to get to orbit with crew or 100t payload. The operational complexity of transfering cargo, including possible large pieces of equipment would be huge. The only thing that will be transfered is fuel. That is my firm opinion. It does not need that huge a launch vehicle. With the second stage being MCT its weight is not fully part of what would usually be counted as payload. The engines and fuel tanks are "free".Won't a single rocket capable of sending 100t to the surface of Mars have to lift like 500+t to LEO? Even with 15m diameter BFR, is it plausible? Sea Dragon was supposed to lift 550t and be 23m in diameter.
The Apollo LEM was about 10mt dry (descent stage), but could land around 6mt of "cargo" in a fully fueled ascent stage, astros, provisions, cargo and equipment. Could a dryer MCT weight about 50mt, and carry 100mt of cargo, and then another X tonnes of propellant when fully fueled? That would put the minimum required single BFR to loft about 150mt. I think it could be around Saturn V size and do that.
Quote from: Lobo on 01/22/2015 09:50 pmThe Apollo LEM was about 10mt dry (descent stage), but could land around 6mt of "cargo" in a fully fueled ascent stage, astros, provisions, cargo and equipment. Could a dryer MCT weight about 50mt, and carry 100mt of cargo, and then another X tonnes of propellant when fully fueled? That would put the minimum required single BFR to loft about 150mt. I think it could be around Saturn V size and do that.As a wild uneducated guess I thought of 60t for the empty MCT so we are in the same ballbpark. Yes 150 to 160t is needed with fuel to spare for first stage RTLS. This leaves no margin for a escape pod or escape engines. Likely at least one, maybe two engines extra for engine loss capability during the long flight and to add ability to separate from a failing first stage if not enough to speed away from a fireball. Early small crews may board separately. Later, when 100 colonists go to Mars there will be a long and hopefully positive history behind the vehicle that allows the confidence to launch them in MCT.
AlsoQuote from: llanitedave on 01/22/2015 02:33 amAt 1.76m engine bells, looks like you can fit 19 of them comfortably on a 10m diameter stage. Think that would be enough for a start?No By the way, what do you use for modelling?
In my thought, all missions would require staging in LEO with several of these "smaller" BFR's to fuel them up, and then some sort of LEO taxi, probably a modified MCT for that purpose, takes the crew up. This LEO taxi could have some sort of LAS system. The one going to Mars doesn't need it. The taxi would take up, up to 100 colonists (eventually) with their personal gear.
Is the fixation on Mars mission profiles making people lose sight that the most important thing necessary to transform space travel is cheap access to low earth orbit? Once cheap access to LEO becomes possible then many other possibilities open up, including slow shipment of non-human freight to Mars and other places by ion propulsion (how the payload gets there doesn't have to decided by SpaceX), not to mention the fact that cheap launches could inspire lower value payloads. Perhaps this kind of thinking is starting to affect the design of SpaceX's next vehicle.I'm sure this has been discussed many times, so could someone please point me at a thread on this forum about optimising cost to LEO, without any other consideration being involved?
So, can anyone give a good reason for us to expect that the BFR will be greater than 10 meters
Its not 19 engines on a 10m stage but 25 and here is why:On an F9 to fit exactly 8 engines in the outer ring requires the bell to be .93m and the engine to engine seperation to be 1.07m.The thrust is related to the square of the bell diameter. So taking the sqrt of the 500/147 Raptor/M1D and multiply by the center to center spacing gives 1.83m Using this spacing value you get exactly 14 engins in the outer ring, 10 in the middle ring and the 1 center engine = 25 engines.Now for estimated payload capability of such a LV take the total thrust of 25 Raptor engins / total thrust of 9 M1Ds both at sL thrust levels and you get the payload increase factor for a scalled paylod capability for a reusable 1st stage. Now multiply by the LEO capability of a reusable F9 of 13.5mt and you get 128mt. Now calculate the ISP advantage by using the Raptor ISP and dividing by yhe M1D ISP 325/280 and multiplying the payload times this advantage number which gives an estimate of 149mt for the 10m BFR using the Raptor 500klbf engines.
The thrust is related to the square of the bell diameter.
So, can anyone give a good reason for us to expect that the BFR will be greater than 10 meters in diameter and have greater than 19 Raptors on the 1st stage?
Sounds like good reasoning to me as well. One question, though.Quote from: oldAtlas_Eguy on 01/23/2015 03:51 pmThe thrust is related to the square of the bell diameter. Doesn't that assume equal chamber pressures? Could a higher pressure allow a still smaller bell diameter?
On an F9 to fit exactly 8 engines in the outer ring requires the bell to be .93m and the engine to engine seperation to be 1.07m.The thrust is related to the square of the bell diameter. So taking the sqrt of the 500/147 Raptor/M1D and multiply by the center to center spacing gives 1.83m Using this spacing value you get exactly 14 engins in the outer ring, 10 in the middle ring and the 1 center engine = 25 engines.
Quote from: GORDAP on 01/23/2015 01:37 pmSo, can anyone give a good reason for us to expect that the BFR will be greater than 10 meters in diameter and have greater than 19 Raptors on the 1st stage?This photo was posted before, the overall height of 10m stage is 120' higher than 12m stage which 10m stage need much larger, longer and more expensive infrastructure (applied to FSS, lighting towers, HIF, storages & Transporter Erectors) but maybe SpaceX have already figured it out...
Now lets look at the number of engines from the exact oposite direction. The BFR min payload goal is 180mt. So starting with that and reversing through all the calculation in order to get nuymber of engines of 31 and then using the engine number you get a stage diameter of ~12.33m.
Quote from: oldAtlas_Eguy on 01/23/2015 03:51 pmOn an F9 to fit exactly 8 engines in the outer ring requires the bell to be .93m and the engine to engine seperation to be 1.07m.The thrust is related to the square of the bell diameter. So taking the sqrt of the 500/147 Raptor/M1D and multiply by the center to center spacing gives 1.83m Using this spacing value you get exactly 14 engins in the outer ring, 10 in the middle ring and the 1 center engine = 25 engines.oldAtlas, I'm having trouble following your math. is the 1.83m the bell diameter of the Raptor under your scenario, or the center to center engine spacing along the outer ring? In either case, how'd you get there: If I multiply either .93m or 1.07m (your F9 figures) by the sqrt of 500/147, I don't get 1.83.
Quote from: Jdeshetler on 01/23/2015 04:30 pmQuote from: GORDAP on 01/23/2015 01:37 pmSo, can anyone give a good reason for us to expect that the BFR will be greater than 10 meters in diameter and have greater than 19 Raptors on the 1st stage?This photo was posted before, the overall height of 10m stage is 120' higher than 12m stage which 10m stage need much larger, longer and more expensive infrastructure (applied to FSS, lighting towers, HIF, storages & Transporter Erectors) but maybe SpaceX have already figured it out...Jdeshetler, I think those photos/designs were done before there was any any SpaceX announced target of how many tons of payload were to be landed on Mars. And the person doing them was speculating on some maximum possible value of payload, perhaps in excess of 200 tons to Mars surface. Now that we have heard Musk say the target is 100 tons, I think the target designs could all be scaled back correspondingly, and a 10 meter core would not be nearly as tall as previously believed. But I could certainly be wrong about my assumptions. Does anyone know more firmly what assumptions were used in the design that produced these pics?
Quote from: GORDAP on 01/23/2015 04:46 pmQuote from: Jdeshetler on 01/23/2015 04:30 pmQuote from: GORDAP on 01/23/2015 01:37 pmSo, can anyone give a good reason for us to expect that the BFR will be greater than 10 meters in diameter and have greater than 19 Raptors on the 1st stage?This photo was posted before, the overall height of 10m stage is 120' higher than 12m stage which 10m stage need much larger, longer and more expensive infrastructure (applied to FSS, lighting towers, HIF, storages & Transporter Erectors) but maybe SpaceX have already figured it out...Jdeshetler, I think those photos/designs were done before there was any any SpaceX announced target of how many tons of payload were to be landed on Mars. And the person doing them was speculating on some maximum possible value of payload, perhaps in excess of 200 tons to Mars surface. Now that we have heard Musk say the target is 100 tons, I think the target designs could all be scaled back correspondingly, and a 10 meter core would not be nearly as tall as previously believed. But I could certainly be wrong about my assumptions. Does anyone know more firmly what assumptions were used in the design that produced these pics?Indeed - and with LEO refueling being in the picture now, it changes things further. The stack does not need to have 3 stages (or MCT on *top* of a 2-stage BFR). It makes the most sense to let MCT be its own 2nd stage - since it needs to be capable of that Delta-V anyway. This would shorten the stack.
Quote from: oldAtlas_Eguy on 01/23/2015 04:32 pmNow lets look at the number of engines from the exact oposite direction. The BFR min payload goal is 180mt. So starting with that and reversing through all the calculation in order to get nuymber of engines of 31 and then using the engine number you get a stage diameter of ~12.33m.Where in the world do you get a BFR min payload of 180mt? Please explain your assumptions.
Quote from: Lars-J on 01/23/2015 05:07 pmQuote from: oldAtlas_Eguy on 01/23/2015 04:32 pmNow lets look at the number of engines from the exact oposite direction. The BFR min payload goal is 180mt. So starting with that and reversing through all the calculation in order to get nuymber of engines of 31 and then using the engine number you get a stage diameter of ~12.33m.Where in the world do you get a BFR min payload of 180mt? Please explain your assumptions.Back when the BFR/MCT was first talked about by SpaceX, Ms Shotwell gave the payload values of between 180-210mt. Those goals for payload could have changed or the goal for BFR is to eventially get to those values with a engine thrust upgrade of 24%. Or it could be some other LV improvement giiving a higher payload such as the US design propelant factors being significantly better than the F9 US. If that is the case then the assumptions about the number of engines or stage diameter could be significantly different.
However, are we certain the math works out for a 19 engine BFR of roughly Saturn 5 scale for landing a 60 tonne MTC + 100 tonnes of cargo + all the fuel + consumables that it's going to use in flight down on the surface of mars? Does a 10m diameter give enough internal volume for 100 people? Nobody wants to stuck in an area the size of a people carrier for months and months in space.Edit: to be succinct; are we sure 19 engines is truly enough engines? It's going to need to lift more than just dry mass + fuel + cargo.
I read somewhere with continuous 1g acceleration and deceleration to and from Mars, that would solve the time and gravity problems using ion propulsion
Quote from: GORDAP on 01/23/2015 01:37 pmSo, can anyone give a good reason for us to expect that the BFR will be greater than 10 meters10m diameter rocket will be too high, perhaps?
Quote from: oldAtlas_Eguy on 01/23/2015 03:51 pmIts not 19 engines on a 10m stage but 25 and here is why:On an F9 to fit exactly 8 engines in the outer ring requires the bell to be .93m and the engine to engine seperation to be 1.07m.The thrust is related to the square of the bell diameter. So taking the sqrt of the 500/147 Raptor/M1D and multiply by the center to center spacing gives 1.83m Using this spacing value you get exactly 14 engins in the outer ring, 10 in the middle ring and the 1 center engine = 25 engines.Now for estimated payload capability of such a LV take the total thrust of 25 Raptor engins / total thrust of 9 M1Ds both at sL thrust levels and you get the payload increase factor for a scalled paylod capability for a reusable 1st stage. Now multiply by the LEO capability of a reusable F9 of 13.5mt and you get 128mt. Now calculate the ISP advantage by using the Raptor ISP and dividing by yhe M1D ISP 325/280 and multiplying the payload times this advantage number which gives an estimate of 149mt for the 10m BFR using the Raptor 500klbf engines.That cut through a lot of discussions, were have you been when we needed you Anyone who can confirm this? Sound so simple put this way that I wonder that no one said it before... (I do not doubt you Atlas, but I do not know enough so more input is good.)Edit: Also when SpaceX change the trust to 165 they probably will not change bell diameter so with that you will get more than 149mt.
Quote from: oldAtlas_Eguy on 01/23/2015 05:18 pmQuote from: Lars-J on 01/23/2015 05:07 pmQuote from: oldAtlas_Eguy on 01/23/2015 04:32 pmNow lets look at the number of engines from the exact oposite direction. The BFR min payload goal is 180mt. So starting with that and reversing through all the calculation in order to get nuymber of engines of 31 and then using the engine number you get a stage diameter of ~12.33m.Where in the world do you get a BFR min payload of 180mt? Please explain your assumptions.Back when the BFR/MCT was first talked about by SpaceX, Ms Shotwell gave the payload values of between 180-210mt. Those goals for payload could have changed or the goal for BFR is to eventially get to those values with a engine thrust upgrade of 24%. Or it could be some other LV improvement giiving a higher payload such as the US design propelant factors being significantly better than the F9 US. If that is the case then the assumptions about the number of engines or stage diameter could be significantly different.Ok, that makes sense. But I do think some of those assumptions are no longer valid, now that LEO depots/refueling is in the picture. The question - of course - is how to make everything work with Musk's "100 tons of cargo to Mars" statement. It could be interpreted a number of ways.
Elon said that BFR would dwarf Saturn 5I speculate since FH has already 27 engines that BFR will have a lot more.Whole lotta little raptors speculation thread ? I say 27 x 2 = 54 And like the M1D very high thrust to weight ratio... 150 or more ??
There is more to a rocket than engine count. Unless you think the F9 dwarfs the Saturn V.
About the diameter of the engines..Raptor will be staged combustion with much higher pressure than Merlin 1d which is gas generator.Doesn't this higher pressure allow smallr nozzle for same thrust?So the nozzle area will not grow by factor of 3 but much less?RD-191 has about 87% of the rumoured thrust of thrust as raptor (2 MJ) and diameter of 1.45m. Pressure is expected to be in same range. Extrapolating the area from rd-191 would give diameter 1.55m for raptor.
Edit: To clarify further, Utrecht's point was as sound as anyone's since we don't actually know the amount of engines the BFR will have. Even the dimensions of the engines talked about are still heavily ambiguous. Since FH has proven to feature a 3x engine count to F9, you can see the line of thought.I agree with you however; there's unlikely to be a linear relationship with any SpaceX rocket we've seen before. However, you have to admit, from F1 to FH, there's been an undeniable tendency to strap together more engines.
Quote from: The Amazing Catstronaut on 01/24/2015 09:02 amEdit: To clarify further, Utrecht's point was as sound as anyone's since we don't actually know the amount of engines the BFR will have. Even the dimensions of the engines talked about are still heavily ambiguous. Since FH has proven to feature a 3x engine count to F9, you can see the line of thought.I agree with you however; there's unlikely to be a linear relationship with any SpaceX rocket we've seen before. However, you have to admit, from F1 to FH, there's been an undeniable tendency to strap together more engines.And you miss to the point that this line of thought only applies if Raptor is the same size as Merlin 1D. And we KNOW that is not the case.
I do not understand why this line of thought only applies when Raptor has the same side as the M1D.On what is that assumption based ?
Quote from: Lars-J on 01/24/2015 10:28 pmQuote from: The Amazing Catstronaut on 01/24/2015 09:02 amEdit: To clarify further, Utrecht's point was as sound as anyone's since we don't actually know the amount of engines the BFR will have. Even the dimensions of the engines talked about are still heavily ambiguous. Since FH has proven to feature a 3x engine count to F9, you can see the line of thought.I agree with you however; there's unlikely to be a linear relationship with any SpaceX rocket we've seen before. However, you have to admit, from F1 to FH, there's been an undeniable tendency to strap together more engines.And you miss to the point that this line of thought only applies if Raptor is the same size as Merlin 1D. And we KNOW that is not the case.Elon Musk said that he first thought of upscaling FH"Goal is 100 metric tons of useful payload to the surface of Mars. This obviously requires a very big spaceship and booster system. ... At first, I was thinking we would just scale up Falcon Heavy, but it looks like ... [the] default plan is to have a sea level and vacuum version of Raptor, much like Merlin. ... [Raptor will boast] a little over 230 metric tons (~500 klbf) of thrust per engine, but we will have a lot of them ".I do not understand why this line of thought only applies when Raptor has the same side as the M1D.On what is that assumption based ?We do not know the diameter of BFR and here are speculations that Raptor can have a diameter from 1.50 meters to 1.85 m1
If he is going to use launch pad 39a, the maximum it was designed for is 11-12 million lbs thrust.
If he is going to use launch pad 39a, the maximum it was designed for is 11-12 million lbs thrust. At 500k lb thrust, they would need around 20 raptors. It could have 18-20 circling two in the middle of a 10-15m wide rocket. This could get 200-225 tons to LEO. If the vehicle refueled in LEO, might get 100 tons to Mars without cargo transfer. Two raptors in the middle should be able to land the first stage. If the Raptor was 1 million lb thrust, they could do the same thing as Falcon 9 on a 12 meter core for 9 million plus lbs thrust. This could work also at pad 39a. A three core heavy version could probably go to Mars and back with one launch but would require a new launch site.
SpaceX have said that the BFR will be too big for LC-39A and they will need a new launch site for it. This indicates that it will have more than 12Mlbf at liftoff. This means that if Raptor for the 1st stage is 500klbf the BFR will need more than 24 of them.
SpaceX have said that the BFR will be too big for LC-39A and they will need a new launch site for it. This indicates that it will have more than 12Mlbf at liftoff. This means that if Raptor for the 1st stage is 500klbf the BFR will need more than 24 of them. BFR could have as many as 37 Raptors under a 15m core with a center engine surrounded by 3 concentric hexagons of 6, 12, and 18 engines. This configuration easily allows all the engines to fit under the core without protruding out in the way of the landing legs while providing the needed thrust.
Honestly I don't believe that SpaceX will put 30 plus engines on their BFR because they won't want to go through the nightmare of the N-1. Too many engines on the 1st stage is potentially a disaster for reliability which was testified by all 4 failures of the N-1. Not to mention all the launch aborts associated with so many engines so even if the launch doesn't fail the launch aborts may cause Mars synods to be missed.I think SpaceX will change their mind on the BFR 1st stage engine and bite the bullet and go for a scaled up Raptor to keep the no. of engines on it's 1st stage to 9 in the proven octaweb configuration of the F9 v1.1. The 500klbf Raptor then be used exclusively on the upper stage/MCT. The 500klbf Raptor is the right size for the MCT spacecraft but way too small for the BFR 1st stage.
From what I understand. The Saturn V started at 7.5 million lbs thrust. The engines were to be upgraded to 1.8 million lbs thrust each or 9 million lbs. I read on this forum, with all the Direct talk that the pads were designed for 11 million lbs thrust. The Saturn V J2 engines on the upper stages were also to be upgraded to 250k thrust each for 1.25 million second stage and 250k 3rd stage. This would have increased the throw weight for an increased LEM size. Also 6-8 Saturn V's were to build a potential Mars exploration vehicle and lander by 1984, before Nasa's budget was cut. So, unless the pads are upgraded or a new pad built at the Texas location, 9-10 million lbs thrust is probably the maximum Raptor will do as well as SLS black knights. Either could get 150 tons to LEO. Which one will be cheaper?
Just a reminder. They are going to fly 27 Merlin engines on a Falcon Heavy. They plan to fly many of those. So is there a limit that prohibits a few more than 27 engines on one launch vehicle if they need the thrust? It gives them a nice production line, even with few cores to be built. It gives them just one engine type for first and upper stage as the larger Raptor could not.My prediction they make the core just big enough that they can launch MCT including payload but without fuel to LEO. And big enough to do RTLS after delivering.
Just a reminder. They are going to fly 27 Merlin engines on a Falcon Heavy. They plan to fly many of those. So is there a limit that prohibits a few more than 27 engines on one launch vehicle if they need the thrust?
No theoretical limit, but certainly a practical one. Remember that they are not flying 27 engines because it is ideal - they are using what they have. But the BFR will be more of a clean sheet design. New size, new fuel, new engines.
Quote from: Lars-J on 01/27/2015 06:08 pmNo theoretical limit, but certainly a practical one. Remember that they are not flying 27 engines because it is ideal - they are using what they have. But the BFR will be more of a clean sheet design. New size, new fuel, new engines.Yes, and knowing that they say they have decided on a large number of engines instead of few very big ones. I suppose they have conciously made that decision. I like others believe they may upgrade thrust per engine later when they need more launches. And I expect they will design their rocket in a way that they can upgrade without major redesign. Making the core and engine spacing wide enough that upgrade will be not much more than stretching the tank.You know, I am one who does not suppose they are fools.
Elon Musk: Thrust to weight is optimizing for a surprisingly low thrust level, even when accounting for the added mass of plumbing and structure for many engines. Looks like a little over 230 metric tons (~500 klbf) of thrust per engine, but we will have a lot of them - See more at: http://www.parabolicarc.com/2015/01/06/highlights-elon-musks-reddit-session/#sthash.Eyy4Y5Vl.dpuf
There's no evidence that the rocket got smaller, at least no reference I am aware of. My "guess" is they'll have 27 of them just because it keeps the rocket the same size. And yes, the plumbing required for that does blow my mind.And of course, it's SpaceX, so they aren't committed until they start bending metal. And even then they aren't totally committed to a specific solution, so I could be wrong. (And to be clear, I like that about them.)
Quote from: guckyfan on 01/27/2015 05:43 pmJust a reminder. They are going to fly 27 Merlin engines on a Falcon Heavy. They plan to fly many of those. So is there a limit that prohibits a few more than 27 engines on one launch vehicle if they need the thrust? It gives them a nice production line, even with few cores to be built. It gives them just one engine type for first and upper stage as the larger Raptor could not.My prediction they make the core just big enough that they can launch MCT including payload but without fuel to LEO. And big enough to do RTLS after delivering.This is my thinking exactly. In fact, I would think 27 engines on one vehicle would be easier than 27 engines on three cores with cross-feed.
N-1 cannot be equivocated with a high Raptor count. All the other variables are different. Thirty plus engines does give a higher probability of having one engine out. OTOH, the probability of a single engine out causing LOM falls to nill. Modern controllers shut an engine (liquid) down long before it explodes. Lose one engine and lose only about 3% of thrust. You just throttle up the other engines a hair, run them longer, or both.
Lobo, I think all your numbers on the pad, Nova, F-1A are correct. I remember them all exactly the same way.
I think the new info about a 500klb Raptor also indicate that BFR is going to be smaller than origainlly envisioned, and flown and reused more per Mars mission. (which makes a lot of sense if you are using reusability as part of your paradigm, you want as high flight rate as possible. Otherwise you have a ginormous BFR that you are only launching once or twice per Mars Launch window for quite some time. SpaceX then would have SLS).
IIRC the ISP for CO2 is only 290 seconds while CH4 is in the high 360+ second range, so the answer is CO2's low efficiency makes it less practical.
Quote from: Lobo on 01/27/2015 05:29 pm I think the new info about a 500klb Raptor also indicate that BFR is going to be smaller than origainlly envisioned, and flown and reused more per Mars mission. (which makes a lot of sense if you are using reusability as part of your paradigm, you want as high flight rate as possible. Otherwise you have a ginormous BFR that you are only launching once or twice per Mars Launch window for quite some time. SpaceX then would have SLS).There is obviously a penalty for full reuse, IE including the upper stage. Even more so for GTO payloads. It seems possible that the goal is to up the flight rate of BFR by eliminating Falcon Heavy, and fully-reusing BFR instead. Without having run any numbers, ISTM that a less-ginourmous BFR could be more economical. Cheers, Martin
Quote from: MP99 on 01/30/2015 08:56 amQuote from: Lobo on 01/27/2015 05:29 pm I think the new info about a 500klb Raptor also indicate that BFR is going to be smaller than origainlly envisioned, and flown and reused more per Mars mission. (which makes a lot of sense if you are using reusability as part of your paradigm, you want as high flight rate as possible. Otherwise you have a ginormous BFR that you are only launching once or twice per Mars Launch window for quite some time. SpaceX then would have SLS).There is obviously a penalty for full reuse, IE including the upper stage. Even more so for GTO payloads. It seems possible that the goal is to up the flight rate of BFR by eliminating Falcon Heavy, and fully-reusing BFR instead. Without having run any numbers, ISTM that a less-ginourmous BFR could be more economical. Cheers, MartinAgreed on the less-ginormous BFR being more economical for heavier commercial/government paylaods. Ditto that for multi-launch Mars Missions. Coupled with something less ginormous with relatively "small" engines being inherently cheaper and easier to process and handle.In my mind, I keep thinking around 150mt to LEO. A 50mt dry MCT and 100mt of stowed cargo. A dry MCT could be heavier than that, that's just a nice round number for speculation at this level. :-)I don't know that I agree on such a less ginormous BFR eliminating FH though. I could see it eliminating a fully expendable FH, or maybe FH that needs the central core expended. Although those payloads would probably be rare. But I just don't see how even a less ginormous BFR could launch from LC-40, SLC-4E, or Boca Chica. And I think FH-R with all 3 cores RTLS will be of a performance level (23mt to LEO, 6-7mt to GTO) that it will probably have a relatively high flight rate. Part of the advantage of F9/FH is the ability for it's stages to be road transportated around, where any sort of BFR will be limited to barge ocean transport.
where any sort of BFR will be limited to barge ocean transport.
Quote from: Burninate on 01/30/2015 01:32 pm Destroying the interior of the chamber with deposits that cannot be removed except in an Earthside factory, is.
Destroying the interior of the chamber with deposits that cannot be removed except in an Earthside factory, is.
If your Mars launch window is limited, then wouldn't it mean that bigger rocket is better, to carry as much as possible thru limited window?
That window is for Trans-Martian Injection. You have two yeas to assemble and transfer fuel to a modular MCT somewhere between LEO and L1 before sending the thing on its way to Mars. Is it better to have a smaller rocket that flies a number of times each cycle to the rendezvous point, or a gigantic rocket that sits in a hangar for two years at a time between short flights? The argument could be made that you could send a much larger assembly through TMI this way than via a single shot on a VBFR.
Quote from: sanman on 01/31/2015 04:21 amIf your Mars launch window is limited, then wouldn't it mean that bigger rocket is better, to carry as much as possible thru limited window?That window is for Trans-Martian Injection. You have two yeas to assemble and transfer fuel to a modular MCT somewhere between LEO and L1 before sending the thing on its way to Mars. Is it better to have a smaller rocket that flies a number of times each cycle to the rendezvous point, or a gigantic rocket that sits in a hangar for two years at a time between short flights? The argument could be made that you could send a much larger assembly through TMI this way than via a single shot on a VBFR.
I don't know that I agree on such a less ginormous BFR eliminating FH though. I could see it eliminating a fully expendable FH, or maybe FH that needs the central core expended. Although those payloads would probably be rare. But I just don't see how even a less ginormous BFR could launch from LC-40, SLC-4E, or Boca Chica. And I think FH-R with all 3 cores RTLS will be of a performance level (23mt to LEO, 6-7mt to GTO) that it will probably have a relatively high flight rate. Part of the advantage of F9/FH is the ability for it's stages to be road transportated around, where any sort of BFR will be limited to barge ocean transport.
The Mars transport system will be a completely new architecture. Am hoping to present that towards the end of this year. Good thing we didn’t do it sooner, as we have learned a huge amount from Falcon and Dragon.
Actually, we could make the 2nd stage of Falcon reusable and still have significant payload on Falcon Heavy, but I think our engineering resources are better spent moving on to the Mars system.
Quote from: TomH on 01/31/2015 07:32 amThat window is for Trans-Martian Injection. You have two yeas to assemble and transfer fuel to a modular MCT somewhere between LEO and L1 before sending the thing on its way to Mars. Is it better to have a smaller rocket that flies a number of times each cycle to the rendezvous point, or a gigantic rocket that sits in a hangar for two years at a time between short flights? The argument could be made that you could send a much larger assembly through TMI this way than via a single shot on a VBFR.Well, I kind of said something not so dissimilar to that in the MCT thread, and they said the problem is that you can't do it for $500K per person.http://forum.nasaspaceflight.com/index.php?topic=35424.msg1323147#msg1323147If you had a Mars-cycler (as per Aldrin's suggestion), and you shuttle people between it and the ground, then your interplanetary vehicle never has to be designed for EDL.Have a purpose-built vehicle that travels from ground to orbit and back. Have another purpose-built vehicle that travels between the planets. That interplanetary vehicle probably might experience less wear-and-tear, given that it only lives in the vacuum of space, and so that helps its reusability lifetime.Meanwhile, the ones that go from Earth-to-orbit get more servicing, and some of them even hitch a ride to Mars to be used for ground-to-orbit over there, and then eventually get sent back to Earth again for more servicing.
Quote from: sanman on 01/31/2015 07:57 amQuote from: TomH on 01/31/2015 07:32 amThat window is for Trans-Martian Injection. You have two yeas to assemble and transfer fuel to a modular MCT somewhere between LEO and L1 before sending the thing on its way to Mars. Is it better to have a smaller rocket that flies a number of times each cycle to the rendezvous point, or a gigantic rocket that sits in a hangar for two years at a time between short flights? The argument could be made that you could send a much larger assembly through TMI this way than via a single shot on a VBFR.Well, I kind of said something not so dissimilar to that in the MCT thread, and they said the problem is that you can't do it for $500K per person.http://forum.nasaspaceflight.com/index.php?topic=35424.msg1323147#msg1323147If you had a Mars-cycler (as per Aldrin's suggestion), and you shuttle people between it and the ground, then your interplanetary vehicle never has to be designed for EDL.Have a purpose-built vehicle that travels from ground to orbit and back. Have another purpose-built vehicle that travels between the planets. That interplanetary vehicle probably might experience less wear-and-tear, given that it only lives in the vacuum of space, and so that helps its reusability lifetime.Meanwhile, the ones that go from Earth-to-orbit get more servicing, and some of them even hitch a ride to Mars to be used for ground-to-orbit over there, and then eventually get sent back to Earth again for more servicing.I'll never understand the fascination with 'cyclers'. They really only give you ONE thing. More living space during the transit. That's all. You still have to bring ALL the supplies with you (no mass savings), you still have to accelerate to TMI (no propellant savings), and for what? To save on habitable volume? Remember that this is living space volume AND cargo volume that will be needed on Mars when you get there.And as a cherry on top, you have a very narrow launch window, and add another failure more if you fail to rendezvous.Cyclers are just a bad idea, and Aldrin arguing for them doesn't make it any better.
Isn’t the fact that the Upper Stage could be part of the MCT a relevant point for calculations?If it was designed for in-space reuse, the current F9 could put 13.5 mT of cargo + the full US (empty), correct? The US can reach orbit?
I believe that BFR will have a basic Upper-Stage structure, which will then be customised into three variants:-1) A basic Upper-Stage which carries a standard payload under a standard PLF. Elon has said that he will be deferring Upper-Stage reuse to BFR, and I believe this will be used for GTO payloads while carrying a large mass penalty for reuse. 2) The same basic structure, but integrated into MCT. This may involve a barrel stretch or shrink of that U/S structure. The reentry of variant (1) at Earth will allow the design to be optimised, which will increase the maturity / safety of EDL at Mars, and is just one of the justifications for (1).3) A stretched version of (1) which will be used as a tanker. No payload. No PLF. Just some sort of ogive-like tank top, like Shuttle's ET.
I'll never understand the fascination with 'cyclers'. They really only give you ONE thing. More living space during the transit. That's all. You still have to bring ALL the supplies with you (no mass savings), you still have to accelerate to TMI (no propellant savings), and for what? To save on habitable volume? Remember that this is living space volume AND cargo volume that will be needed on Mars when you get there.And as a cherry on top, you have a very narrow launch window, and add another failure more if you fail to rendezvous.Cyclers are just a bad idea, and Aldrin arguing for them doesn't make it any better.
Note: Your comment re restricted launch windows gives me pause. Hmm, maybe I'm misunderstanding 'cyclers'? I envision a craft that's constructed in LEO, and transits to Mars orbit where it offloads cargo and people to a lander that came up from the surface, then it transits back to Earth orbit where it gets it's next batch of cargo and people. Repeat ad infinitum. Guess I'll read up on Aldrin's version.