I think the operations costs would kill you on that. Falcon 9 is only 3.66 meters in diameter – even if you stuck a Bigelow Hab on some of them, it’d still be quite cramped for a trip to Mars, and you’d have to launch up a lot of them to assemble the craft, assemble the return vehicle, and launch up the fuel for both. And you’d have to deploy everything in advance on Mars’ surface, since your lander would be tiny.
Falcon Heavy reduces the number of launches, but it’s got the same narrow diameter. I’d feel a lot more comfortable with something like the Saturn V’s 6-10 meters to work with – that would give you more space for a habitat that 4-6 people have to live in for three years.
Someone (an actual rocket scientist no doubt) here pointed out recently that the ratio of length to width on a rocket has a finite limit and that the recently stretched F9 was very close to that limit.
I imagine it would make it harder to aerobrake around Mars, too, if you’ve got a long, narrow spacecraft behind the heat shield. Or you’d just have to jettison most of it along the way, so that only the capsule and lander does direct entry on to Mars’ surface and hopefully lands near the habitat and supplies you’ve already sent.
Yes, that’s called the fineness ratio, the ratio of length to width. The Falcon 9’s fineness ratio is 19.13 to 1 (229.6 ft long to 12 ft wide), which is very high for such a large rocket. Von Braun preferred to design rockets with a fineness ratio of less than 10-1, but most rockets today use higher strength materials than Von Braun had to work with, and are around 14-1.
The longer a rocket is, the more bending forces it is subjected to, particularly if it is flying through areas with a high wind shear. This, by the way, is the reason there are launch commit critera for wind speeds both at ground level and through the flight path. I personally would reject any design for a large rocket with a fineness ratio much higher than the Falcon 9’s. 20-1 is probably going to bend under even minor wind load.
I suppose there are lots of issues that make stiffness important, not the least being the gimbaled engines pointing in the right direction:-)
Reading your post, Doug, I imagined large buildings on earth, all designed to move with the wind rather than resisting. Similarly, earthquake zone construction requires improvements to be able to ‘wiggle’ or tolerate modest deformation without failing.
I’m guessing that the deformation tolerance on a rocket is mighty damn tight, especially given the fleetest of time that it is even subject to earth’s atmosphere (something lie ten minutes or so).
It’s more the fact that the rocket is moving and the building (hopefully) is not. A small bend in the structure changes how the rocket is pointed relative to the center of mass, the aerodynamics of the whole vehicle, etc. If it starts oscillating, that’s even worse. All that can lead to control problems and instabilities, and there’s only so much the guidance system can cope with.
Using smaller launch vehicles with narrower payload shrouds would definitely be less efficient. But the claimed cost savings is a factor of four. If that is correct, even a factor of two lower efficiency is still a clear winner.
FWIW, the Endurance in Interstellar is a good example of a MLV/MHLV-launched mission spacecraft: Very Modular and clearly assembled in LEO.
The big kicker is the assembly costs. For these not to make the concept more expensive than nearly all-in-one, as with NASA’s DRM-5, then rendezvous and assembly must be as close to automated as possible. It would likely require some manner of reusable orbital tug vehicle.
13 meters for the payload fairing would be fantastic. That’d be more than enough space for a good-sized hab for 4-6 people for three years, especially if they have multiple levels.
I’ve wondered about that narrow fairing. Why can’t the pointy-end of the rocket be much much wider? Like the modified aircraft used around the world with bulbous noses? Sure, there’s the non-trivial aerodynamic issues. But with the cost delta so gargantuan surely smart people are thinking the same thing.
I’d guess it’s because putting extra size and weight in the second or third stage means you’d need an even bigger second or first stage to push it up, so it would just make the rocket itself larger. But there are rockets that are top-heavy, like Atlas V.
Ben R-G mentioned up-thread that you might be able to make the payload launch fairing up to twice as wide as the stage below it, so a 7 meter wide payload versus 3.66. That’s still tight for both two six-month transits plus an extended stay on the Martian surface (or in Mars orbit), but more doable.
I take your point and should have said, since most readers here are apparently of the fussy engineering/ scientific variety and thus more appreciative of exactitude, that what I meant to say is wider but not necessarily more mass.
Lots of launch vehicles have payload fairings wider than the second stage. I suspect SpaceX designed the current one with their customers in mind: There aren’t many revenue-generating payloads in the Falcon 9 mass range which require a larger volume. That doesn’t mean SpaceX couldn’t design a larger fairing if there were customers, or when the Falcon Heavy is up and launching more massive payloads.
Musk once said, in a interview, that the Falcon Heavy payload fairing can be made larger. It would require strengthening the core vehicle. Then he went on to say that the current payload fairing space haven’t been maxed out yet. And if needed they will make a larger payload fairing.
Reusable Falcon Heavies for Mars? Actually, now that you ask, Google the 2015 Miller grant/study, the Evolvable Lunar Architecture, which evolves into a lot of propellant for Mars missions, using Falcon Heavy or some similar emerging heavy (ULA, if the price is right for Vulcan’s). The lower cost expected there, from SpaceX as an example, opens many possibilities.
Also, Google the 2011 NASA Propellant Depot Requirements study, which offers an alternative to Lunar ISRU using the orbital propellant depot concept.
Something tells me SpaceX will eventually enlist elements of either propellant depots in LEO, or Lunar resources, or some proving ground missions, especially lunar landings, into Mars plans. None of these elements are currently in NASA plans, studies and trades showing their feasibility, lower cost and value to the contrary be damned.
o Reuseability is THE Key to Low Cost Operations o Refuelability is Key to Reuseability (must refuel transport) o In Space Propellant Delivery, Storage, and Transfer Holds the Key to the future, and likely creates new Market(s)
Re-launching and taking risks with dirt cheap class D propellant is a game changer. Boeing Agrees (‘amplification factor”), as does original Shuttle Designer Max Faget-“we really need to get behind a first stage that is recoverable and piggyback off of that event”, and also 1960s Apollo hero John Houbolt, “rendezvous and leave weight in orbit”.
Most lunar resources arrived via asteroids–most are between Mars and Jupiter, so ISRU is not just limited to the moon.
So are larger rockets and payload fairings required? Falcon has a 5.2m fairing that is built and tested, but even ULA does not have a 7m fairing, which would require doable facility and LV modifications, *if* there was a need.
So what is the architecture fuel, LH2 or methane? For LOx and methane, 7m not needed, even for one year long stays, because its dense, and subcooling, demonstrated by Falcon helps. LH2 is volume limited, so a 7m fairing could reduce the number of LH2 launches-trade its development cost vs kg/yr.
Methane however, increases the Initial Mass to LEO by ~25% vs LH2 depending on the deltaV. One can tip the trade one way or the other depending on assumptions. A key driver however is trip time to Mars, as is cost, and it will be difficult to displace LH2 prepositioned by Electric propulsion.
On evolvable lunar architectures, unfortunately, destinations (e.g. lunar or any 1/6th g or greater gravity well) divert all resources away from reusability, infrastructure, and the long term space travel technology to head ‘to Mars’. Simply stated, one can head to Mars with expendable, smaller rockets (HLV need not apply).
Even Griffin, tasked with casting aside the depot centric architecture for HLV, agrees on avoiding gravity wells until the infrastructure is in place. “An interplanetary vehicle at L2 is reused for many trips to many destinations and the energy savings at such a node are significant. Note that human landings on the moon or Mars are not included, although landings on the Martian moons could be made, as they have negligible gravitational attraction and no atmosphere..both are safer and more cost effective than going directly to the planetary surfaces.”
Its not NASA’s plans….look no further than Congress.
Or, keep using Falcon 9 and Falcon Heavy to refine operations while developing the much bigger, fully reusable, LOX/methane launch vehicle that Musk has in mind for just such a mission. Why? Because this is truly a “first generation” vehicle and isn’t at all optimal. The lessons have not all been learned yet, since this is only the first landing of the first stage.
Part of the reason NASA launch vehicles are so expensive is that they get locked into a “first generation” vehicle design. The mistakes of the shuttle design weren’t fully known until it had already flown maybe a dozen times. But, by then it was “too late” to fix them, since many were fundamental to the design. So, NASA had to live with the far too low launch rate and far too high costs for the life of the program. Now, NASA has completely abandoned that design, so we won’t see a second generation (i.e. much improved) space shuttle design.
I read somewhere that the maximum fairing diameter is ~2x the stage diameter on which it is perched. So, theoretically, could a Falcon Heavy fly with a 7m PLF, maybe extended down to the top of the interstage to reduce bending?
Congrats to SpaceX getting the landing to work and it is certainly a big deal, but the real trick is going to be whether or not they can really achieve their projected cost savings. Remember that the Space Shuttle managed to achieve an even higher percentage of vehicle recovery right from its first flight 34 years ago, yet never was able to achieve any of its projected cost savings from reusability even after three decades of operations. We’ll see where this goes. (Not to mention the mismatching of development versus operational costs, the ignoring of the ISS lessons about the expense of breaking up unitary payloads into small launch packages, the reliability hole that Falcon is still climbing out of, etc. that are implicitly ignored in your contrast to SLS.)
While SLS has proven subsystems, the entire vehicle doesn’t have any reliability record yet. There’s nothing to ignore.
Dividing payloads in to smaller units definitely has an interfacing cost. (Even Mars reference missions require multiple SLS launches.)
Do the interfaces between smaller units increase the aggregate payload cost by a factor of 20-30? Doubtful.
That’s one of the strategic exploration architecture trade-offs.
The time convenience of *quickly* heavy lifting “all in one”, or a few, isn’t feasibly in SLS’s advantage either. NASA has to be very patient to accumulate the SLS inventory required to launch the Martian fleet.
According to NASA, SLS is limited to a production rate of 1 or 2 per year where as SpaceX is currently finishing upgraded Falcon 9s about once every 2-3 weeks.
Which means SpaceX is manufacturing SLS-equivalent LEO lift capacity every 10-15 weeks. (ULA is building SLS-equivalent launch capacity every few months as well.)
In current expendable lunar and Mars missions over 70% of the mass is propellant, so 120mt or up to 450 mT * 30% is 36 mT to 135 mT is hardware and supplies. All of the current hardware elements are less than 20mT, and the only exception may be the Mars aerobrake, which has not been developed.
So that leaves one with at most 2-6 launches of hardware filled with supplies to reach gravity well missions, much less for many other destinations. The hardware and supplies can be launched to LEO and checked out prior to departure, even repaired by crew. ISS demonstrated this is not an issue and clearly way less than the 3B/yr fixed costs of the production lines. Meanwhile, the gas station is being filled at a zero boiloff depot independent of the hardware, supplies and propellant are being pre-positioned, waiting for the crew and to fill the departure stages just prior to blast off.
SLS is based on solids which have ignored the fixed costs of the production lines and the performance impact of staging too soon (the diameter is too small due to shipping limitations). As a result, the SSME has to work harder and is heavier in the compromise, and now the SSME becomes expendable, which removes the post flight inspection to find flaws. Splashing Orion offers little hope of reuse. The costs of refurbishment to the solids, main engines, the tiles.. resulted in a LV that cost about billion per launch, and with lower launch rate, more for SLS/Orion.
A more realistic commercial mars mission would probably use more than just the FH. It would probably be staged in LEO(it could depart from either LEO or much further out). Faring diameter could prove to be pointless as inflatable heat shields, in space assemble or even some type of folding could be developed. It could be launched at an higher rate than the shuttle due to being able to process payloads in parallel. The commercial crew and commercial cargo craft could provide both crew for assembly work and supplies for that work.
This is just so blindingly obvious- the advantages of FH, I mean, over SLS- that at some point even Congress will notice.
I predict that Congress will follow the leadership of the military, which showed keen interest in SpaceX even very early. Those guys might have lots of money to spend but they also have large appetites, and they do appreciate a bigger bang for the buck.
It isn’t a jobs program for.everyone in congress. There are plenty of Congressional districts that don’t get a cent from it. Assuming their representatives are interested in jobs programs, their interest is in having SLS supporters.return the favor (by supporting other jobs programs) and in avoiding public embarrassment. That puts some limits on how blatant and useless a jobs program can become.
What you say is true of all pork programs which send money back to a few (or one) state or even congressional district. Congressmen and Senators tolerate this in a very “I’ll scratch your back while you scratch mine” sort of way.
One of the things that was slipped into the U.S. Commercial Space Launch Competitiveness Act that has been over looked was the little gem of Title I, Sec. 117.
Basically under it the Space Launch System replaces the Space Shuttle in U.S. Code 51, Chapter 701, the section of U.S.Code that basically made the use of the Space Shuttle free of commercial competition. The link to the section is here.
In plain English it means if the NASA Administrator decides it is in the ‘national interest’ that something should fly on the SLS it will fly on it, regardless of cheaper commercial alternatives.
So talk about Falcon Heavy as a replacement for the SLS may be nice, and logical in a sane world, but in the world of NASA, NASA is free to use the SLS to go to Mars regardless of costs associated with it or even, in theory, stiff CCP systems and use the SLS if it is the “national interest” as determined by the NASA Administrator.
That will not change until the NON space state Reps and Senators see the first blood in the water (once commercial is viable to take over i.e. Falcon heavy starts flying) and then what always happens the poster boy for a particular federal agency’s pork disappears and they no longer fund the old and the new gets all the pork.
Sorry I did not see this discussion last night, as we were still celebrating after decades of waiting. The two key issues are time frame and aero-braking area.
– ‘We’ will not be ready to go to Mars until the SpaceX super-booster is ready, with its 10 meter diameter. Any lunar missions can be conducted with Falcon Heavy since there is no atmosphere to use for braking. I am sure that the BFR will be ready to use before ‘we’ are ready to go to Mars.
– Mars has an atmosphere that lets a vehicle slow down for essentially free when we arrive. For that method to be most efficient, we want a really WIDE lander/ferry vehicle, not a narrow one. The reason for this is that the larger the base area of the ferry, the more it slows down during entry and the less propellant you need to land. You can easily land more cargo than propellant with each trip down. Thus the Mars EDL “problem” is not a problem any more, it is a landing aide!
To use that vehicle at Mars, we first must put it into Earth orbit. A 10 meter dia. booster will allow the 15 meter dia. ferry vehicles to be launched into LEO, so they can be used at Mars. With the BFR, we will also be free of the “Apollo Mission Model” thinking and able to take advantage of high mass mission thinking. With a fleet of Mars vehicles to enhance crew safety, we could land over 600 tons of cargo at a Mars base with just the first expedition. (That mass total does not count the mass of the ferries, since they are used for multiple trips.) Just think of what a crew could do at a base with that much mass.
NASA has been flight testing inflatable heat shields for possible use on a Mars mission. So, (with a bit more R&D) it might just be possible to have a much bigger diameter heat shield than your launch fairing diameter.
The discussion has become focused on Mars strategies based on lessons learned, and in that frame I’ve always wondered why a true ‘space ship’ isn’t used between Earth and Mars/Moon/wherever, a machine that lives on solar orbit cleverly arranged between destinations, never landing.
Fuel is one issue, clearly. And the idea isn’t new; a novel by Buzz Aldrin had a ‘cycler’ concept, a spacecraft that never went into orbit but relied on rendezvous at each end. Tricky and possibly dangerous no doubt, but by now rendezvous is routine (and I do recall Gemini missions).
The whole thing would rely on depots, and in situ tech, but has the advantage of not throwing away so much damn hardware.
It is possible to design orbit so that the spaceship can travel between earth and mars without much extra energy. However, I am not sure there is a clear energy advantage with such a ship. To rendezvous with the ship on earth side, your shuttle needs to be as fast as the ship itself. In other words, your ship has enough kinetic energy to go to mars all by itself. Of course you can build some large real estate while in the orbit with this method.
The benefit of a cycler is you can launch a small crew ship and a separate supply capsule from Earth to rendezvous with the cycler. Yes, they’ll need to be put on a speed and trajectory to Mars, but the size of the rocket needed to launch people and supplies from Earth to the cycler is smaller than if you were launching crew, habitat, and supplies from Earth to Mars all at one time.
Good points, all of which are covered in the aforementioned novel, which I need to read again as the details aren’t fresh.
More than anything the concept has a purity about it, a sensibility that we are a space faring race, the certainty that wasteful splashing or disposal of very very expensive one-use command or other modules is just inefficient and most significantly not conducive to sustainability.
In many ways the ‘throwaway’ attitude contributed to the fall of the Apollo program, although admittedly Apollo was so cutting edge it was hard to see if any of the approaches would be viable.
Over at centaurs-dreams Paul Gilster has been discussing habitability at red dwarf stars and posted a link to the Apollo 10 transcripts. Some very fun reading and quite enlightening.
“Trying to picture what it would cost to use reusable Falcon rockets to send humans to Mars Vs multiple SLS rockets at $1billion+ each.“
NASA has contracted Boeing to supply the first two engine-less cores for $2.8 billion. So the cores alone, excluding the solid boosters, the upper-stage, refurbishing the remaining recycled Shuttle engines, and all the integration and ops cost, cost at least $1.4b each.
Working out the total price is pretty difficult. NASA is spending about $3b/yr on SLS and Orion, recently higher but it’s an election year. They’ve spent billions up to now, and will spend another $21b between 2016 and 2022 (inclusive) and the first EM-2 manned flight. So that’s two flights (EM-1 and 2) for $21b.
At some point in the 2020’s they’ll produce a 105 tonne cargo variant, and by 2032 they are expected to fly a 130 tonne cargo version. At $3b/yr, that’s around $50 billion between now and 2032.
They are planning a mission every 2 years or so. And some of the missions will be double launches (one cargo, one manned), although none of those are funded yet. So we’re looking at maybe ten missions in twenty years, between 2020 and 2040. (I’m counting the 2018 launch of EM-1 as the equivalent of the 2020 mission.) Let’s say half of those are double launches. And half those double launches are 130 tonnes.
So about 15 SLS launches between now and 2040-ish. 8-10 SLS-70’s. 3 SLS-105’s. 3 SLS-130’s. Plus 6-10 Orion capsules.
Cost is $50b + what has been spent up ’til now + whatever proportion of the current $3b/yr will need to be spent beyond 2032. Say $70 billion all up, just to give us a concrete number.
That’s around $4.7b per launch. Or $7b per “mission”.
Total theoretical capacity between now and 2040 is therefore just under 800 tonnes to LEO (or equivalent.)
Double it to 1600, divide by 40 tonnes to take a worst case for FH, and you’ve got 40 launches to allow FH to substitute for SLS. SpaceX wants to charge around $100m for non-reusable FH launches, and typically charges NASA/DoD a 50-100% bump for the extra paperwork. But let’s say $250m per NASA launch, so a total of $10 billion to launch similar mass into LEO. Leaving NASA an extra $40 billion over the next 15 years to develop mission hardware, habitats, etc.
Obviously reusability lowers the price further.
While SLS can launch single payloads larger than FH’s 40 (or 50) tonnes, none of the proposed architectures have a single element massing much more than 40 tonnes. The extra capacity is used to “throw” that 20, 30 or 40 tonne payload directly into lunar orbit or BEO.
With FH, you’d instead launch the 20, 30, or 40 tonne payload into LEO on one flight, and launch a special kick stage on another flight which would dock with the mission modules and boost them into their mission trajectory.
Such a docking orbital booster stage would be a generally useful technology (and is similar to LM’s proposed Jupiter refuelable cargo tug.) It should only cost a few billion to develop, and at least three, maybe four, US companies are capable of developing it. (As would ESA and Japan.) So a COTS/CC style competition would be ideal, keeping the price down.
Such a modular docking-boost-stage gives you a flexibility that SLS’s “direct launch” architecture doesn’t have. For example, a double-launch SLS mission would be limited to one SLS-70 carrying only Orion, it’s service module, and pushing it out to LEO in order to carry a crew to the 40 tonne module put into lunar orbit by the SLS-130.
By contrast, with a modular approach, you can launch as many modules as you want for a specific mission. And as many crew as you need.
And with tens of billions freed up from cancelling SLS/Orion development, you have the funding to actually build those habs and landers and so on. Whereas with SLS, no mission hardware can be developed until extra funding becomes available (on top of the $3b for SLS/Orion.) Since Congress is unlike to significantly increase NASA’s funding. So you are waiting until ISS is deorbited to free up its $3b to allow you to develop habs/etc starting in 2028 or so.
Killing SLS/Orion and investing immediately in multipurpose, adaptable mission hardware means that when ISS comes to an end, its $3b/yr can be directed to a brand new program. A new program in addition to whatever NASA was going to do with SLS.
Kudos to Paul for his good cost analysis of using the SLS.
“I imagine it would make it harder to aerobrake around Mars, too, if you’ve got a long, narrow spacecraft behind the heat shield.” (theBrett) Yes, the obvious solution is for entry vehicles (ferries) to be large, capsule shaped vehicles with ~15 meter dia. bases. They can use their built-in heat shields for the single pass aero-capture, then later for entry during a landing. All of the other vehicles, carrying cargo or modules that would stay in Low Mars orbit, would be placed behind large external heat shields up to 30 meters across. The ferries can then carry all the cargo down to the surface in successive round trip missions. The large shields can be assembled in space since they are not pressure vehicles and have no “insides”. Using long, narrow spacecraft for entry make no sense given the usefulness of the Mars atmosphere for aero-braking during entry. With poor use of the atmosphere, you might need over 2000 meters per second to enter and land. (Entry velocity from orbit is about 3800 meters per second.) With a cone-shaped capsule with a wide base, you only need about 800 meters per second.
The longer a vehicle is, the more mass it tends to have compared to its volume. The heavier your vehicle is, the less cargo it can carry.
Imagine an umbrella point first going along in a gale, then imagine what happens when you open the umbrella in the gale. The Mars atmosphere is a free “air brake” for your vehicle. A short cone presents the maximum area to the air stream with the minimum mass, giving you the maximum payload.
A long narrow vehicle also requires active stabilization during entry, or it could tumble out of control and be destroyed. A short cone with its base facing down is hard to tip over in the last phase of landing, while a long narrow vehicle can easily tip over, like the Falcon 9’s that tried to land on the ship. It is also easier to unload cargo from the side hatch of a cone. How do you unload cargo easily from the top of a tall rocket?
A biconic or lifting body design increases the cross-sectional area facing the entry interface, without needing a wide base (and hence oversized launch vehicle.)
(Also, capsules don’t scale well. One of the reasons Orion is struggling.)
It is also easier to unload cargo from the side hatch of a cone. How do you unload cargo easily from the top of a tall rocket?
How is putting a hatch in the side of a tall vehicle harder than putting a hatch in the side of a capsule?
A sphere has the smallest surface area for its volume and mass but it would not be stable during entry. Robot probes have often used biconic entry designs, but with most of these the entry shell halves (heat shield and back-shell) are discarded before landing. If we want a re-usable vehicle, the capsule shape is a very good compromise. A cargo compartment can be located just above the engines, with the propellant tanks above that. Having the cargo bay and side hatch closer to the ground makes it much easier to unload. If you have a tall vehicle, you will need a tall crane on Mars to unload it. How does the crane unload itself. An unloader vehicle for a capsule-shaped ferry can be designed to first unload itself.
Why worry about scaling when the SpaceX plans to build a booster of the right size (10 meters dia.) are in the works. That is what the Raptor engines are for. Reverse fairings were used for the Constellation design upper stage that were 50% wider than the first stage. They would work fine for a 15 meter payload.
This is a completely valid question. The two big issues for me are (1) that you cannot get the maximum braking area out of a more cylindrical vehicle. To do this the vehicle would need to fly at 90 degrees to its long axis during entry and it would not be stable in that position (and would sustain thermal damage if it was). The other issue is (2) how do you land such a vehicle on Mars. A horizontal landing would require a prepared runway and extremely high landing speed, and a vertical landing would take us back to the tall, narrow vehicle that could tip over during a vertical landing attempt. The reason they gave the lunar module such a wide, low profile is to reduce the danger of tipping over. So building such a vehicle is physically possible, but riskier to land and would land less payload compared to its mass due to lower passive drag (entry) efficiency.
The other issue is reusability. Can the narrow vehicle take off again and return to Low Mars orbit for another load? I see no landing legs or wheels on either of those vehicles.
I admit I have no clue as to what kind of passenger vehicle types SpaceX will come up with for use to LEO and for use from Low Mars orbit to the surface. I suspect that the vehicles will be wider and larger than you think they will. I have heard (strictly) rumors of a 15 meter wide version of the BFR, which would allow a 20+ meter wide vehicle to be launched. it will take a really wide vehicle of some kind to hold the 100 passengers Musk is aiming at. This would be the exact opposite of the ultra-slim Falcon 9 rockets he is currently flying.
Love the discussion, but the one thing I want to see before planning the future is 27 engines lighting at the same time, without destroying the stack, the side mounts, or the launch pad. That, along with the first launch of a refurb’d first stage are going to be the next SpaceX big moments.
Well if we could send up everything in a F9 why not use a Falcon 1 since it is cheaper or maybe even a bottle rocket (those are super cheap)?
Mr. Kavanaugh is being factious. Cost is not the only consideration. Capability is a major concern and it affects all other operations. Sure you could launch 3 Falcon Heavies to make up for one SLS launch in terms of cargo but you have massively increased the complexity and cost of the overall mission. Increasing the amount of orbital assembly is not something you want to do (especially on BEO missions).
Also the ability to stage from somewhere other than LEO is very important. SLS offers a distinct advantage there with its optimized BEO upper stage.
Falcon Heavy will be an asset to BEO missions but demanding that everything must be put on is shoulders is a mistake in my view. Instead lets develop all the tools we are working on right now and get them to work together
First of all the post posited re-usable Falcon Heavies (which would get 36mt or so to LEO vs. 53mt for the expendable version). SLS can get 86mt to LEO in Block I and around 100mt in Block IB. Also even if expendable Falcon Heavies are assumed in order to stage from somewhere other than LEO you need 3 FH’s to replace 1 SLS Block IB
So BOTH are totally non flying examples? A reusable Falcon Heavy is not available as is the SLS of ANY block.. neither have flown so both numbers are power point at this time.
Nope. F9 was modified so that it enables the re-usable versions to take more cargo up, and still return. IOW, F9R still does some 13+ tonnes, while the FHR does over 53+ tonnes.
I think the operations costs would kill you on that. Falcon 9 is only 3.66 meters in diameter – even if you stuck a Bigelow Hab on some of them, it’d still be quite cramped for a trip to Mars, and you’d have to launch up a lot of them to assemble the craft, assemble the return vehicle, and launch up the fuel for both. And you’d have to deploy everything in advance on Mars’ surface, since your lander would be tiny.
Falcon Heavy reduces the number of launches, but it’s got the same narrow diameter. I’d feel a lot more comfortable with something like the Saturn V’s 6-10 meters to work with – that would give you more space for a habitat that 4-6 people have to live in for three years.
So make it longer.
Someone (an actual rocket scientist no doubt) here pointed out recently that the ratio of length to width on a rocket has a finite limit and that the recently stretched F9 was very close to that limit.
I imagine it would make it harder to aerobrake around Mars, too, if you’ve got a long, narrow spacecraft behind the heat shield. Or you’d just have to jettison most of it along the way, so that only the capsule and lander does direct entry on to Mars’ surface and hopefully lands near the habitat and supplies you’ve already sent.
Yes, that’s called the fineness ratio, the ratio of length to width. The Falcon 9’s fineness ratio is 19.13 to 1 (229.6 ft long to 12 ft wide), which is very high for such a large rocket. Von Braun preferred to design rockets with a fineness ratio of less than 10-1, but most rockets today use higher strength materials than Von Braun had to work with, and are around 14-1.
The longer a rocket is, the more bending forces it is subjected to, particularly if it is flying through areas with a high wind shear. This, by the way, is the reason there are launch commit critera for wind speeds both at ground level and through the flight path. I personally would reject any design for a large rocket with a fineness ratio much higher than the Falcon 9’s. 20-1 is probably going to bend under even minor wind load.
I suppose there are lots of issues that make stiffness important, not the least being the gimbaled engines pointing in the right direction:-)
Reading your post, Doug, I imagined large buildings on earth, all designed to move with the wind rather than resisting. Similarly, earthquake zone construction requires improvements to be able to ‘wiggle’ or tolerate modest deformation without failing.
I’m guessing that the deformation tolerance on a rocket is mighty damn tight, especially given the fleetest of time that it is even subject to earth’s atmosphere (something lie ten minutes or so).
It’s more the fact that the rocket is moving and the building (hopefully) is not. A small bend in the structure changes how the rocket is pointed relative to the center of mass, the aerodynamics of the whole vehicle, etc. If it starts oscillating, that’s even worse. All that can lead to control problems and instabilities, and there’s only so much the guidance system can cope with.
Using smaller launch vehicles with narrower payload shrouds would definitely be less efficient. But the claimed cost savings is a factor of four. If that is correct, even a factor of two lower efficiency is still a clear winner.
FWIW, the Endurance in Interstellar is a good example of a MLV/MHLV-launched mission spacecraft: Very Modular and clearly assembled in LEO.
The big kicker is the assembly costs. For these not to make the concept more expensive than nearly all-in-one, as with NASA’s DRM-5, then rendezvous and assembly must be as close to automated as possible. It would likely require some manner of reusable orbital tug vehicle.
5.2 m/13m with the payload fairing could make it work
13 meters for the payload fairing would be fantastic. That’d be more than enough space for a good-sized hab for 4-6 people for three years, especially if they have multiple levels.
I’ve wondered about that narrow fairing. Why can’t the pointy-end of the rocket be much much wider? Like the modified aircraft used around the world with bulbous noses?
Sure, there’s the non-trivial aerodynamic issues. But with the cost delta so gargantuan surely smart people are thinking the same thing.
I’d guess it’s because putting extra size and weight in the second or third stage means you’d need an even bigger second or first stage to push it up, so it would just make the rocket itself larger. But there are rockets that are top-heavy, like Atlas V.
Ben R-G mentioned up-thread that you might be able to make the payload launch fairing up to twice as wide as the stage below it, so a 7 meter wide payload versus 3.66. That’s still tight for both two six-month transits plus an extended stay on the Martian surface (or in Mars orbit), but more doable.
I take your point and should have said, since most readers here are apparently of the fussy engineering/ scientific variety and thus more appreciative of exactitude, that what I meant to say is wider but not necessarily more mass.
Lots of launch vehicles have payload fairings wider than the second stage. I suspect SpaceX designed the current one with their customers in mind: There aren’t many revenue-generating payloads in the Falcon 9 mass range which require a larger volume. That doesn’t mean SpaceX couldn’t design a larger fairing if there were customers, or when the Falcon Heavy is up and launching more massive payloads.
It is. Falcon 9’s payload faring is 5.2m wide (4.6m internal.)
(By comparison, the upperstage is 3.6m wide.)
http://deepspaceindustries….
http://i2.wp.com/www.univer…
Pictures of rockets on the pad are so deceiving! With no experiential scale it ooks like a toy when it is actually gargantuan.
The payload fairing can be 5m in diameter. What is the concern with the rocket diameter (constrained by truck shipment from factory).
Musk once said, in a interview, that the Falcon Heavy payload fairing can be made larger. It would require strengthening the core vehicle. Then he went on to say that the current payload fairing space haven’t been maxed out yet. And if needed they will make a larger payload fairing.
You underestimate the size of the Falcon 9 payload fairing.
https://upload.wikimedia.or…
http://www.wired.com/images…
And the similar sized Ariane payload fairing,
http://spacenews.com/wp-con…
You’re also underestimating the advantage in being able to launch fifteen of them for the same price as a single year’s funding of SLS/Orion.
Reusable Falcon Heavies for Mars? Actually, now that you ask, Google the 2015 Miller grant/study, the Evolvable Lunar Architecture, which evolves into a lot of propellant for Mars missions, using Falcon Heavy or some similar emerging heavy (ULA, if the price is right for Vulcan’s). The lower cost expected there, from SpaceX as an example, opens many possibilities.
Also, Google the 2011 NASA Propellant Depot Requirements study, which offers an alternative to Lunar ISRU using the orbital propellant depot concept.
Something tells me SpaceX will eventually enlist elements of either propellant depots in LEO, or Lunar resources, or some proving ground missions, especially lunar landings, into Mars plans. None of these elements are currently in NASA plans, studies and trades showing their feasibility, lower cost and value to the contrary be damned.
o Reuseability is THE Key to Low Cost Operations
o Refuelability is Key to Reuseability (must refuel transport)
o In Space Propellant Delivery, Storage, and Transfer Holds the Key to the future, and likely creates new Market(s)
Re-launching and taking risks with dirt cheap class D propellant is a game changer. Boeing Agrees (‘amplification factor”), as does original Shuttle Designer Max Faget-“we really need to get behind a first stage that is recoverable and piggyback off of that event”, and also 1960s Apollo hero John Houbolt, “rendezvous and leave weight in orbit”.
Most lunar resources arrived via asteroids–most are between Mars and Jupiter, so ISRU is not just limited to the moon.
So are larger rockets and payload fairings required? Falcon has a 5.2m fairing that is built and tested, but even ULA does not have a 7m fairing, which would require doable facility and LV modifications, *if* there was a need.
So what is the architecture fuel, LH2 or methane? For LOx and methane, 7m not needed, even for one year long stays, because its dense, and subcooling, demonstrated by Falcon helps. LH2 is volume limited, so a 7m fairing could reduce the number of LH2 launches-trade its development cost vs kg/yr.
Methane however, increases the Initial Mass to LEO by ~25% vs LH2 depending on the deltaV. One can tip the trade one way or the other depending on assumptions. A key driver however is trip time to Mars, as is cost, and it will be difficult to displace LH2 prepositioned by Electric propulsion.
On evolvable lunar architectures, unfortunately, destinations (e.g. lunar or any 1/6th g or greater gravity well) divert all resources away from reusability, infrastructure, and the long term space travel technology to head ‘to Mars’. Simply stated, one can head to Mars with expendable, smaller rockets (HLV need not apply).
Even Griffin, tasked with casting aside the depot centric architecture for HLV, agrees on avoiding gravity wells until the infrastructure is in place. “An interplanetary vehicle at L2 is reused for many trips to many destinations and the energy savings at such a node are significant. Note that human landings on the moon or Mars are not included, although landings on the Martian moons could be made, as they have negligible gravitational attraction and no atmosphere..both are safer and more cost effective than going directly to the planetary surfaces.”
Its not NASA’s plans….look no further than Congress.
http://nextbigfuture.com/20…
http://www.jsc.nasa.gov/his…
http://www.lpi.usra.edu/lun…
Or, keep using Falcon 9 and Falcon Heavy to refine operations while developing the much bigger, fully reusable, LOX/methane launch vehicle that Musk has in mind for just such a mission. Why? Because this is truly a “first generation” vehicle and isn’t at all optimal. The lessons have not all been learned yet, since this is only the first landing of the first stage.
Part of the reason NASA launch vehicles are so expensive is that they get locked into a “first generation” vehicle design. The mistakes of the shuttle design weren’t fully known until it had already flown maybe a dozen times. But, by then it was “too late” to fix them, since many were fundamental to the design. So, NASA had to live with the far too low launch rate and far too high costs for the life of the program. Now, NASA has completely abandoned that design, so we won’t see a second generation (i.e. much improved) space shuttle design.
I read somewhere that the maximum fairing diameter is ~2x the stage diameter on which it is perched. So, theoretically, could a Falcon Heavy fly with a 7m PLF, maybe extended down to the top of the interstage to reduce bending?
Congrats to SpaceX getting the landing to work and it is certainly a big deal, but the real trick is going to be whether or not they can really achieve their projected cost savings. Remember that the Space Shuttle managed to achieve an even higher percentage of vehicle recovery right from its first flight 34 years ago, yet never was able to achieve any of its projected cost savings from reusability even after three decades of operations. We’ll see where this goes.
(Not to mention the mismatching of development versus operational costs, the ignoring of the ISS lessons about the expense of breaking up unitary payloads into small launch packages, the reliability hole that Falcon is still climbing out of, etc. that are implicitly ignored in your contrast to SLS.)
While SLS has proven subsystems, the entire vehicle doesn’t have any reliability record yet. There’s nothing to ignore.
Dividing payloads in to smaller units definitely has an interfacing cost. (Even Mars reference missions require multiple SLS launches.)
Do the interfaces between smaller units increase the aggregate payload cost by a factor of 20-30? Doubtful.
That’s one of the strategic exploration architecture trade-offs.
The time convenience of *quickly* heavy lifting “all in one”, or a few, isn’t feasibly in SLS’s advantage either. NASA has to be very patient to accumulate the SLS inventory required to launch the Martian fleet.
According to NASA, SLS is limited to a production rate of 1 or 2 per year where as SpaceX is currently finishing upgraded Falcon 9s about once every 2-3 weeks.
Which means SpaceX is manufacturing SLS-equivalent LEO lift capacity every 10-15 weeks. (ULA is building SLS-equivalent launch capacity every few months as well.)
In current expendable lunar and Mars missions over 70% of the mass is propellant, so 120mt or up to 450 mT * 30% is 36 mT to 135 mT is hardware and supplies. All of the current hardware elements are less than 20mT, and the only exception may be the Mars aerobrake, which has not been developed.
So that leaves one with at most 2-6 launches of hardware filled with supplies to reach gravity well missions, much less for many other destinations. The hardware and supplies can be launched to LEO and checked out prior to departure, even repaired by crew. ISS demonstrated this is not an issue and clearly way less than the 3B/yr fixed costs of the production lines. Meanwhile, the gas station is being filled at a zero boiloff depot independent of the hardware, supplies and propellant are being pre-positioned, waiting for the crew and to fill the departure stages just prior to blast off.
SLS is based on solids which have ignored the fixed costs of the production lines and the performance impact of staging too soon (the diameter is too small due to shipping limitations). As a result, the SSME has to work harder and is heavier in the compromise, and now the SSME becomes expendable, which removes the post flight inspection to find flaws. Splashing Orion offers little hope of reuse. The costs of refurbishment to the solids, main engines, the tiles.. resulted in a LV that cost about billion per launch, and with lower launch rate, more for SLS/Orion.
A more realistic commercial mars mission would probably use more than just the FH. It would probably be staged in LEO(it could depart from either LEO or much further out). Faring diameter could prove to be pointless as inflatable heat shields, in space assemble or even some type of folding could be developed. It could be launched at an higher rate than the shuttle due to being able to process payloads in parallel. The commercial crew and commercial cargo craft could provide both crew for assembly work and supplies for that work.
This is just so blindingly obvious- the advantages of FH, I mean, over SLS- that at some point even Congress will notice.
I predict that Congress will follow the leadership of the military, which showed keen interest in SpaceX even very early. Those guys might have lots of money to spend but they also have large appetites, and they do appreciate a bigger bang for the buck.
So to speak.
SLS is primarily a jobs program for Congress. Do you really think they care if it is cost effective?
It isn’t a jobs program for.everyone in congress. There are plenty of Congressional districts that don’t get a cent from it. Assuming their representatives are interested in jobs programs, their interest is in having SLS supporters.return the favor (by supporting other jobs programs) and in avoiding public embarrassment. That puts some limits on how blatant and useless a jobs program can become.
What you say is true of all pork programs which send money back to a few (or one) state or even congressional district. Congressmen and Senators tolerate this in a very “I’ll scratch your back while you scratch mine” sort of way.
One of the things that was slipped into the U.S. Commercial Space Launch Competitiveness Act that has been over looked was the little gem of Title I, Sec. 117.
http://thomas.loc.gov/cgi-b…
Basically under it the Space Launch System replaces the Space Shuttle in U.S. Code 51, Chapter 701, the section of U.S.Code that basically made the use of the Space Shuttle free of commercial competition. The link to the section is here.
https://www.law.cornell.edu…
In plain English it means if the NASA Administrator decides it is in the ‘national interest’ that something should fly on the SLS it will fly on it, regardless of cheaper commercial alternatives.
So talk about Falcon Heavy as a replacement for the SLS may be nice, and logical in a sane world, but in the world of NASA, NASA is free to use the SLS to go to Mars regardless of costs associated with it or even, in theory, stiff CCP systems and use the SLS if it is the “national interest” as determined by the NASA Administrator.
That will not change until the NON space state Reps and Senators see the first blood in the water (once commercial is viable to take over i.e. Falcon heavy starts flying) and then what always happens the poster boy for a particular federal agency’s pork disappears and they no longer fund the old and the new gets all the pork.
Sorry I did not see this discussion last night, as we were still celebrating after decades of waiting. The two key issues are time frame and aero-braking area.
– ‘We’ will not be ready to go to Mars until the SpaceX super-booster is ready, with its 10 meter diameter. Any lunar missions can be conducted with Falcon Heavy since there is no atmosphere to use for braking. I am sure that the BFR will be ready to use before ‘we’ are ready to go to Mars.
– Mars has an atmosphere that lets a vehicle slow down for essentially free when we arrive. For that method to be most efficient, we want a really WIDE lander/ferry vehicle, not a narrow one. The reason for this is that the larger the base area of the ferry, the more it slows down during entry and the less propellant you need to land. You can easily land more cargo than propellant with each trip down. Thus the Mars EDL “problem” is not a problem any more, it is a landing aide!
To use that vehicle at Mars, we first must put it into Earth orbit. A 10 meter dia. booster will allow the 15 meter dia. ferry vehicles to be launched into LEO, so they can be used at Mars. With the BFR, we will also be free of the “Apollo Mission Model” thinking and able to take advantage of high mass mission thinking. With a fleet of Mars vehicles to enhance crew safety, we could land over 600 tons of cargo at a Mars base with just the first expedition. (That mass total does not count the mass of the ferries, since they are used for multiple trips.) Just think of what a crew could do at a base with that much mass.
John Strickland
NASA has been flight testing inflatable heat shields for possible use on a Mars mission. So, (with a bit more R&D) it might just be possible to have a much bigger diameter heat shield than your launch fairing diameter.
The discussion has become focused on Mars strategies based on lessons learned, and in that frame I’ve always wondered why a true ‘space ship’ isn’t used between Earth and Mars/Moon/wherever, a machine that lives on solar orbit cleverly arranged between destinations, never landing.
Fuel is one issue, clearly. And the idea isn’t new; a novel by Buzz Aldrin had a ‘cycler’ concept, a spacecraft that never went into orbit but relied on rendezvous at each end. Tricky and possibly dangerous no doubt, but by now rendezvous is routine (and I do recall Gemini missions).
The whole thing would rely on depots, and in situ tech, but has the advantage of not throwing away so much damn hardware.
It is possible to design orbit so that the spaceship can travel between earth and mars without much extra energy. However, I am not sure there is a clear energy advantage with such a ship. To rendezvous with the ship on earth side, your shuttle needs to be as fast as the ship itself. In other words, your ship has enough kinetic energy to go to mars all by itself. Of course you can build some large real estate while in the orbit with this method.
The benefit of a cycler is you can launch a small crew ship and a separate supply capsule from Earth to rendezvous with the cycler. Yes, they’ll need to be put on a speed and trajectory to Mars, but the size of the rocket needed to launch people and supplies from Earth to the cycler is smaller than if you were launching crew, habitat, and supplies from Earth to Mars all at one time.
Good points, all of which are covered in the aforementioned novel, which I need to read again as the details aren’t fresh.
More than anything the concept has a purity about it, a sensibility that we are a space faring race, the certainty that wasteful splashing or disposal of very very expensive one-use command or other modules is just inefficient and most significantly not conducive to sustainability.
In many ways the ‘throwaway’ attitude contributed to the fall of the Apollo program, although admittedly Apollo was so cutting edge it was hard to see if any of the approaches would be viable.
Over at centaurs-dreams Paul Gilster has been discussing habitability at red dwarf stars and posted a link to the Apollo 10 transcripts. Some very fun reading and quite enlightening.
NASA has contracted Boeing to supply the first two engine-less cores for $2.8 billion. So the cores alone, excluding the solid boosters, the upper-stage, refurbishing the remaining recycled Shuttle engines, and all the integration and ops cost, cost at least $1.4b each.
Working out the total price is pretty difficult. NASA is spending about $3b/yr on SLS and Orion, recently higher but it’s an election year. They’ve spent billions up to now, and will spend another $21b between 2016 and 2022 (inclusive) and the first EM-2 manned flight. So that’s two flights (EM-1 and 2) for $21b.
At some point in the 2020’s they’ll produce a 105 tonne cargo variant, and by 2032 they are expected to fly a 130 tonne cargo version. At $3b/yr, that’s around $50 billion between now and 2032.
They are planning a mission every 2 years or so. And some of the missions will be double launches (one cargo, one manned), although none of those are funded yet. So we’re looking at maybe ten missions in twenty years, between 2020 and 2040. (I’m counting the 2018 launch of EM-1 as the equivalent of the 2020 mission.) Let’s say half of those are double launches. And half those double launches are 130 tonnes.
So about 15 SLS launches between now and 2040-ish. 8-10 SLS-70’s. 3 SLS-105’s. 3 SLS-130’s. Plus 6-10 Orion capsules.
Cost is $50b + what has been spent up ’til now + whatever proportion of the current $3b/yr will need to be spent beyond 2032. Say $70 billion all up, just to give us a concrete number.
That’s around $4.7b per launch. Or $7b per “mission”.
Total theoretical capacity between now and 2040 is therefore just under 800 tonnes to LEO (or equivalent.)
Double it to 1600, divide by 40 tonnes to take a worst case for FH, and you’ve got 40 launches to allow FH to substitute for SLS. SpaceX wants to charge around $100m for non-reusable FH launches, and typically charges NASA/DoD a 50-100% bump for the extra paperwork. But let’s say $250m per NASA launch, so a total of $10 billion to launch similar mass into LEO. Leaving NASA an extra $40 billion over the next 15 years to develop mission hardware, habitats, etc.
Obviously reusability lowers the price further.
While SLS can launch single payloads larger than FH’s 40 (or 50) tonnes, none of the proposed architectures have a single element massing much more than 40 tonnes. The extra capacity is used to “throw” that 20, 30 or 40 tonne payload directly into lunar orbit or BEO.
With FH, you’d instead launch the 20, 30, or 40 tonne payload into LEO on one flight, and launch a special kick stage on another flight which would dock with the mission modules and boost them into their mission trajectory.
Such a docking orbital booster stage would be a generally useful technology (and is similar to LM’s proposed Jupiter refuelable cargo tug.) It should only cost a few billion to develop, and at least three, maybe four, US companies are capable of developing it. (As would ESA and Japan.) So a COTS/CC style competition would be ideal, keeping the price down.
Such a modular docking-boost-stage gives you a flexibility that SLS’s “direct launch” architecture doesn’t have. For example, a double-launch SLS mission would be limited to one SLS-70 carrying only Orion, it’s service module, and pushing it out to LEO in order to carry a crew to the 40 tonne module put into lunar orbit by the SLS-130.
By contrast, with a modular approach, you can launch as many modules as you want for a specific mission. And as many crew as you need.
And with tens of billions freed up from cancelling SLS/Orion development, you have the funding to actually build those habs and landers and so on. Whereas with SLS, no mission hardware can be developed until extra funding becomes available (on top of the $3b for SLS/Orion.) Since Congress is unlike to significantly increase NASA’s funding. So you are waiting until ISS is deorbited to free up its $3b to allow you to develop habs/etc starting in 2028 or so.
Killing SLS/Orion and investing immediately in multipurpose, adaptable mission hardware means that when ISS comes to an end, its $3b/yr can be directed to a brand new program. A new program in addition to whatever NASA was going to do with SLS.
Thanks for running this down.
Kudos to Paul for his good cost analysis of using the SLS.
“I imagine it would make it harder to aerobrake around Mars, too, if you’ve got a long, narrow spacecraft behind the heat shield.” (theBrett)
Yes, the obvious solution is for entry vehicles (ferries) to be large, capsule shaped vehicles with ~15 meter dia. bases. They can use their built-in heat shields for the single pass aero-capture, then later for entry during a landing. All of the other vehicles, carrying cargo or modules that would stay in Low Mars orbit, would be placed behind large external heat shields up to 30 meters across. The ferries can then carry all the cargo down to the surface in successive round trip missions. The large shields can be assembled in space since they are not pressure vehicles and have no “insides”.
Using long, narrow spacecraft for entry make no sense given the usefulness of the Mars atmosphere for aero-braking during entry. With poor use of the atmosphere, you might need over 2000 meters per second to enter and land. (Entry velocity from orbit is about 3800 meters per second.) With a cone-shaped capsule with a wide base, you only need about 800 meters per second.
Er… …why can’t a long spacecraft enter the atmosphere sideways? A bit like the famous ‘breadbox’ shuttle design..
The longer a vehicle is, the more mass it tends to have compared to its volume. The heavier your vehicle is, the less cargo it can carry.
Imagine an umbrella point first going along in a gale, then imagine what happens when you open the umbrella in the gale. The Mars atmosphere is a free “air brake” for your vehicle. A short cone presents the maximum area to the air stream with the minimum mass, giving you the maximum payload.
A long narrow vehicle also requires active stabilization during entry, or it could tumble out of control and be destroyed. A short cone with its base facing down is hard to tip over in the last phase of landing, while a long narrow vehicle can easily tip over, like the Falcon 9’s that tried to land on the ship. It is also easier to unload cargo from the side hatch of a cone. How do you unload cargo easily from the top of a tall rocket?
A biconic or lifting body design increases the cross-sectional area facing the entry interface, without needing a wide base (and hence oversized launch vehicle.)
(Also, capsules don’t scale well. One of the reasons Orion is struggling.)
How is putting a hatch in the side of a tall vehicle harder than putting a hatch in the side of a capsule?
A sphere has the smallest surface area for its volume and mass but it would not be stable during entry. Robot probes have often used biconic entry designs, but with most of these the entry shell halves (heat shield and back-shell) are discarded before landing. If we want a re-usable vehicle, the capsule shape is a very good compromise.
A cargo compartment can be located just above the engines, with the propellant tanks above that. Having the cargo bay and side hatch closer to the ground makes it much easier to unload. If you have a tall vehicle, you will need a tall crane on Mars to unload it. How does the crane unload itself. An unloader vehicle for a capsule-shaped ferry can be designed to first unload itself.
Why worry about scaling when the SpaceX plans to build a booster of the right size (10 meters dia.) are in the works. That is what the Raptor engines are for. Reverse fairings were used for the Constellation design upper stage that were 50% wider than the first stage. They would work fine for a 15 meter payload.
Not sure why you made this comment.
You misunderstand the EDL profile I’m talking about.
Not this:
https://upload.wikimedia.or…
This:
https://www.researchitaly.i…
Or this:
http://www.buran.fr/kliper-…
Increasing surface area for the heat shield beyond diameter of the launch vehicle, by changing the angle of attack, is pretty standard.
https://encrypted-tbn0.gsta…
Even then, no-one is really expecting SpaceX to use a 10 or 15m capsule design.
http://forum.nasaspacefligh…
http://planete-mars.com/wp-…
I’m not sure why you are insisting that the heat shield is limited by the cross-section of the launch vehicle.
This is a completely valid question. The two big issues for me are (1) that you cannot get the maximum braking area out of a more cylindrical vehicle. To do this the vehicle would need to fly at 90 degrees to its long axis during entry and it would not be stable in that position (and would sustain thermal damage if it was). The other issue is (2) how do you land such a vehicle on Mars. A horizontal landing would require a prepared runway and extremely high landing speed, and a vertical landing would take us back to the tall, narrow vehicle that could tip over during a vertical landing attempt. The reason they gave the lunar module such a wide, low profile is to reduce the danger of tipping over. So building such a vehicle is physically possible, but riskier to land and would land less payload compared to its mass due to lower passive drag (entry) efficiency.
The other issue is reusability. Can the narrow vehicle take off again and return to Low Mars orbit for another load? I see no landing legs or wheels on either of those vehicles.
I admit I have no clue as to what kind of passenger vehicle types SpaceX will come up with for use to LEO and for use from Low Mars orbit to the surface. I suspect that the vehicles will be wider and larger than you think they will. I have heard (strictly) rumors of a 15 meter wide version of the BFR, which would allow a 20+ meter wide vehicle to be launched. it will take a really wide vehicle of some kind to hold the 100 passengers Musk is aiming at. This would be the exact opposite of the ultra-slim Falcon 9 rockets he is currently flying.
Love the discussion, but the one thing I want to see before planning the future is 27 engines lighting at the same time, without destroying the stack, the side mounts, or the launch pad. That, along with the first launch of a refurb’d first stage are going to be the next SpaceX big moments.
Well if we could send up everything in a F9 why not use a Falcon 1 since it is cheaper or maybe even a bottle rocket (those are super cheap)?
Mr. Kavanaugh is being factious. Cost is not the only consideration. Capability is a major concern and it affects all other operations.
Sure you could launch 3 Falcon Heavies to make up for one SLS launch in terms of cargo but you have massively increased the complexity and cost of the overall mission. Increasing the amount of orbital assembly is not something you want to do (especially on BEO missions).
Also the ability to stage from somewhere other than LEO is very important. SLS offers a distinct advantage there with its optimized BEO upper stage.
Falcon Heavy will be an asset to BEO missions but demanding that everything must be put on is shoulders is a mistake in my view. Instead lets develop all the tools we are working on right now and get them to work together
If the Falcon Heavy can push 53 metric tons in to LEO and the SLS can do 77 metric tons to LEO why would you have to launch 3 to 1 ?
First of all the post posited re-usable Falcon Heavies (which would get 36mt or so to LEO vs. 53mt for the expendable version). SLS can get 86mt to LEO in Block I and around 100mt in Block IB. Also even if expendable Falcon Heavies are assumed in order to stage from somewhere other than LEO you need 3 FH’s to replace 1 SLS Block IB
So BOTH are totally non flying examples? A reusable Falcon Heavy is not available as is the SLS of ANY block.. neither have flown so both numbers are power point at this time.
Nope. F9 was modified so that it enables the re-usable versions to take more cargo up, and still return.
IOW, F9R still does some 13+ tonnes, while the FHR does over 53+ tonnes.
Which optimized second stage the one that just received funding?
I was referring to the EUS which has received funding in the latest omnibus bill.