Ok, two weeks ago, I mentioned that the “pre-depot” dual EELV launch concept my friend had passed to me could be adapted to do manned lunar missions. Lunar missions are a lot easier to close if you assume a depot in LEO (and even easier if there’s also a small depot at L2). But it turns out that if you use a couple of tricks, you can actually make a pre-depot concept close as well. This wouldn’t be my optimal approach, but it at least illustrates the point.
The mission uses the following tricks to make things work:
- Dual Engine Centaur for this mission is stretched by 50% and includes an “Extended Mission Kit” to allow for it to function for the ~5 days necessary for the mission (normal DEC dry mass is ~5400lb, and the EMK is ~1750lb and includes stuff like extra hydrazine bottles, more batteries, deep space navigation upgrades to avionics, sunshields, etc)
- Command module does a powered lunar swingby to go to L2, thus cutting down on overall dV requirements (~750m/s total required, 335m/s per leg), thus allowing for a much smaller CSM (possibly with the service module integrated into the command module).
- The Stretched Centaur and the Lander break into lunar orbit and descend to the surface instead of continuing to L2. I’m not positive if this allows you to land anywhere on the lunar surface or not (this is one of the few big questions for this mission mode). This avoids the extra dV requirements you normally get for stopping everything at L2 first.
- Upper stage performs part of the landing burn (between LOI and the descent burn it provides about 1950m/s out of the total 3050m/s needed for LOI and landing).
- RS-68A Upgraded Delta-IVH. This upgrade is already in engine testing and is badly needed by the DoD, so there’s a good chance this will work out. Expected payload capacity I’ve heard is 27mT for the system.
- Instead of carrying a second stretched Centaur as a payload on one of the flights, the Atlas V 552 uses the stretched Centaur as its upper stage. In order to tank up the LH2, it carries an LH2 drop tank between the lander and the command module. It gets transfered right after reaching orbit, and gets dumped shortly before TLI.
Here are the major components of the system:
- Command Module: This module is based on the Apollo outer mold line, but only carries two people, and enough life support consumables for the mission. I budgetted 11,000lb dry and 3250lb of propellant for the capsule (not including RCS propellants). I assumed hypergols for the stage, with a crappy 314s Isp. The Apollo CM wet mass was 12.8klb, and the SM weighed 54klb wet, 13.5klb dry. However, most of the SM mass was due to the CSM performing the LOI burn for the Apollo Stack. About half of the dry mass of the CSM was the huge main engine, and a good chunk of the remaining mass was electrical equipment and the huge tanks for the 40klb of propellant. With modern materials, electronics, a smaller crew, solar panels instead of fuel cells, and the much lower propulsive requirements for the Command Module in this architecture, I think 11klb is actually pretty conservative for such a system. For another comparison the latest CEV numbers I’ve heard (which are pretty far out of date) were ~18klb for a four person capsule.
- Stretched Centaur Lunar Transfer/Crasher Stage: As mentioned above, this is a dual engine centaur using two RL10A-4-2 engines, but with a 50% barrel stretch to the tanks. The tanks are actually less than 40% of the dry mass of a centaur stage, but you also need more helium for pressurization of the larger stage…assuming that the 50% greater propellant load requires a 50% higher dry mass should be a conservative estimate. The idea of a stretched Centaur shouldn’t be too crazy when you realize how many iterations General Dynamics, Martin Marietta, and Lockheed Martin have done on the Centaur just in the past 20 years (including 5m diameter Centaurs for use on Titan IV among other things). The 1750lb for the extended mission kit is also based on numbers from previous papers LM/ULA has published about converting their stages over for longer-duration missions. Total dry mass I assumed was 9850lb. Note that the Atlas V 552 performance numbers also include 5400lb worth of Centaur burnout weight, so you only have to provide ~4450lb worth of “payload” for the Stretched Centaur. Also note, that if you tank the stretched Centaur up all the way for launch, it should probably increase the payload capacity of the Atlas V 552 a little compared to a normal Centaur, but for purposes of this analysis we’re assuming only the nominal payload of a normal Atlas V 552, to be conservative.
- Single Stage Lunar Lander/Ascender: This stage takes the crew the rest of the way to the lunar surface after the Centaur has provided the first part of the descent burn, and then provides the ascent burn, and the burn to take the crew to the L2 staging point to rendezvous with the Command Module. I budgetted 1100m/s for its portion of the descent burn, 100m/s to allow for a 90s hover to find the best landing spot, 2650m/s for the lunar surface to L2 burn, and about 50m/s more for contingencies. This is probably the most aggressive part of the mission. For this vehicle, I’m assuming a piston-pump-fed LOX/CH4 stage, based off of the piston pump and LOX/Methane engine work XCOR has done (possibly combined with stuff that we at Masten have done that they haven’t like gimbals, throttling, etc). The piston pump requires very low net peak suction head, which allows for very low pressure tanks, that can be made of the LOX/Cryo-compatible Nonburnite composites that XCOR has been devleoping. XCOR developed the piston pump and Nonburnite composites explicity for making propellant tanks out of shapes that aren’t typical for propellant tanks (in their cases to make the CG numbers work, they wanted to do LOX-filled “wet wings”). Using this technology, instead of heavy pressure fed tanks and heavy helium tanks, you have lightweight composite tanks that can actually form part of the load-bearing structure of the vehicle. As I understand it, based on my recollection of their public statements, the piston pumps they’re looking at using scale to about enough flow for a 2500lbf engine in a single pump. By combining them with the 7500lbf engine XCOR developed (with a nozzle extension of course), you have significantly more thrust than you need for landing. More importantly, you can possibly make the three pumps operate in a redundant fashion, so the loss of one pump can be tolerated at any point in the mission, and the loss of a second pump can be tolerated through most of the mission. If done right, the pumps could be “armored” as XCOR calls it, but placed in such a way that they have removable manways between them and the main compartment that would allow for shirtsleeve troubleshooting/repair (the pump compartments would need to be done in a manner that if something went horribly wrong, that any debris/blast would be directed away from the crew cabin…but I can imagine a few ways that could be done). All told, I’m assuming a 4350lb dry weight, a 9000lb propellant weight, 500lb worth of hardware to be left on the moon, and a 360s Isp. The LM ascent stage was 4200lb, but held only 65% of the propellant mass, and only about half the propellant volume of this lander, and didn’t have to do landings, and didn’t have to support the crew for as long (about 3 days vs. the target 9 days to give you a week on the surface and 2 days in transity to L2). But as mentioned above, it used pressure fed tanks, with the mass of a helium blowdown system, had to provide significant RCS capabilities since the stage did not have a gimballed main engine, was using crappy 60s era electronics and electrical systems, and had tanks that were entirely non structural, and also didn’t have access to modern materials like lithium-aluminum or modern composites. However, the 13,850lb total mass for the lander actually compares pretty well with the 13,510lb currently assumed for the pressure-fed, hypergol-fueld Altair Ascent stage (from this document), which carries 4 crew for the same mission duration.
- Pre-Depot LOX Tank: This ~2.2klb Tank holds ~57.1klb of LOX for the Stretched Centaur. It includes a docking port (possibly using LIDS technology?), a sunshield, and a Centuar-derived LOX tank. It gets launched as the sole payload for the Delta-IVH, using up all but about 200lb of its capacity. But since it is so dense, it might be able to get away with using a shorter (and lighter weight) fairing than is typical for Delta-IVH if that wouldn’t require lots of expensive aero analysis. This tank, if launched with the LOX pre-chilled can hang out for over a month waiting for the Atlas V 552 launch.
- LH2 Drop Tank: This ~62.5 m^3 tank weighs about 2000lb (with another 2000lb budgetted for connecting structures between the various parts of the launch stack). It would be housed between the Lander and the Command Module on the Atlas V 552 launch. It would possibly use 5m tankage derived from the Delta-IV US. After reaching orbit, the LH2 from this tank would be transfered (using propulsive settling) into the Stretched Centaur. After the Command Module docks with the Pre-Depot LOX tank, and has transferred all the propellants from that (and discarded the pre-depot LOX tank), the CM and empty LH2 drop tank would separate from the stack, the drop tank would be discarded, and the CM would reattach to the lander much like was done on the Apollo Missions.
Now, this mission model isn’t perfect. It uses most of the capabilities of the two launchers without a huge amount of margin (except in the fact that the Atlas V 552 with stretched Centaur probably has some margin built in that isn’t being explicitly called out). And I’m not a fan of launching the crew on an EELV with 5 solid strapons. It would be a lot easier if you assumed the development of something like the Common Upper Stage that ULA has been talking about recently. With that, you would have tons more margin (since a CUS would add nearly 7mT of capacity to the DIVH, and probably at least 5mT to the Atlas V 552–possibly enough to go with less or no strapons on the crew launcher). But it demonstrates that a 2-launch EELV mission using almost no modifications to existing launch vehicles (beyond the Centaur mods) is within feasibility.
The system also has several good things going for it. First off, it can deliver lunar crew to the surface without a depot. It doesn’t need Autonomous Rendezvous and Docking (since the rendezvous and docking can be piloted), or tankers to be developed. It doesn’t need HLVs or 10m fairings (everything can fit within a stock Atlas V fairing). It doesn’t need really long term LH2 storage in orbit. It only requires two launches for the mission, and doesn’t put anywhere near as much launch timing constraints as the ESAS architecture does. It can provide for cargo missions (~19klb delivered mass to the surface assuming that 2klb of the lander stage is in the form of a removable crew cabin, which just happens to be enough to land a Bigelow Module).
And most importantly, if depots do come into existence, it can immediately take advantage of them. With just an LEO depot, you can both cut down on the number of EELV launches to just one (and use lower-cost systems like Falcon 9’s, Zenits, Ariane-Vs, Soyuzes, future commercial RLVs, etc to launch the remaining propellant). Also by getting rid of the huge LH2 drop tank, you simplify the stack, remove about 15klb worth of hardware from the Atlas stack , dropping it to the point where it can possibly be launched by a 502 launch instead of a 552 launch (since the stretched Centaur provides almost as much propellant as a Phase 1 Atlas, which was supposed to boost the LEO capacity of the single-stick Atlas to almost 30klb). Or you could use that saved mass to beef up the lander and/or command module for more capable missions.
If you have both a LEO and an L1 or L2 depot, the Centaur can top itself up again that depot, and provide a much larger chunk of the descent burn to the lander stack. With enough propellant left over to return to LLO then to L1/L2 after separating from the lander, allowing the Stretched Centaur to be reused multiple times. With such a system you could actually soft-land bigger payloads than the Altair cargo lander…and you’d have the capability of making the lander and transfer stage fully reusable. The transfer stage, since it wouldn’t see atmospheric flight, reentry, lunar dust, or even particularly bad thermal environments should actually be reusable for several flights–the RL10 is after all rated for 200 relights. The lander may be tougher, but by the time you have an L1/L2 depot, you’ve probably had enough time (and enough surface infrastructure built up) that you can work that out to.
Ok, so maybe it’s not so bad of an idea after all.
Jonathan Goff
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I don’t get the basic mission shape, where does the crew go to the lander?
You can’t launch the crew with the lander you can’t abort in it since it’s not built for atmospheric flight.
I agree with gravityloss. I accept the fact that you have run the numbers and that it may be possible to do this, but there’s no clear synopsis of the actual mission mode. What is the sequence of events? Where do all the pieces go? And for how long? I think I may have pieced some of it together, but it would certainly help if you could give a play-by-play of how the mission is supposed to go.
Some elements are unclear to me, too. It might be simpler to do a propellant depot, and a spacecraft / rocket configuration closer to what actually exists.
I really like the idea Jon, except the margin’s probably too small as you said for it to really be feasible –all the downrating for launch profile and LAS that a crewed Atlas V would undergo is what I’m thinking of, not to mention the extra weight of ‘keep alive’ and rendezvous systems the prelaunched LOX tank would require.
But programmatically as an evolutionary idea it’s a perfect “pathway” between what exists now and what some deride as the ‘fantasy’ of a depot. By having the LH2 drop tank already attached to the Centaur, you eliminate a lot of the risk of what most see as the risky parts of a depot, as well as the boiloff concerns for your average depot.
Meanwhile you still have the rendezvous and LOX transfer in orbit to demonstrate the basics of orbital depot technology. From there you’ve proved basically all the mechanics of a depot, it’s just having the will to launch the necessary commercial depot supply train.
Speaking of LOX transfer, where were you thinking that happens physically — I mean physically where would the Delta-delivered LOX depot connect to a stretch Centaur to load LOX once it’s buried in the middle of the launch stack? Were you thinking in the vicinity of the dropped LH2 tank?
Hey guys,
I’ll try to work up some cartoons tomorrow to explain what I was thinking (I try not to do rocket stuff on Sundays). Sorry if it wasn’t clear.
~Jon
Jon,
Great post.
It may help you to assume that the SpaceX Dragon capsule is used in your 2-EELV lunar architecture, and that a Falcon-9 Heavy is a launch option as an EELV. It may also help to assume that ULA will build a Delta IV Heavy upper stage with 4 RL-10B engines (and 100,000 lbs thrust) like Aerospace Corp assumed during the Augustine Panel a few weeks ago. This would probably increase performance by 10-tons or so, and ULA could probably easily finance this themselves if ULA were a real company that re-invested to improve its products and if ULA wasn’t a joint-venutre/government-contractor.
The Russian Lunar Mission profile (discussed at the Russian Spaceweb web site) is a similiar 2-EELV approach using dual Lunar Orbit Rendezvous for a 4-man mission. The Russians basically use an Atlas-V Heavy launch vehicle with a 4 RL-10 engine upper-stage called the “Rus-M” for their 2 EELV lunar architecture.
If SpaceX decided to develop and test an LH2 version of their Merlin engine (which their propulsion team had experience doing when working in their previous jobs at TRW and Boeing/Rocketdyne), their Falcon-9 Heavy would have the performance (i.e. over 40-tons to LEO and over 16-tons to Lunar Transfer Orbit) to do a 2-EELV Lunar manned mission with a SpaceX Dragon capsule carrying 4 people. The only thing missing for SpaceX would be a lunar lander built by you at Masten with a ~ 10,000-lb thrust class engine from XCOR or with 4 of your 2,500-lb thrust engines intended for your XA-1.0 vehicle.
It seems pretty clear to me – the crew transfers to the lander prior to the now uncrewed CM doing the swingby and parking itself at L2. The crew descend to the surface, and then after their stay ascend direct to a rendezvous with the CM at L2, then do a powered swingby on the way home. I’m not sure about abort profiles compared with a CM doing an LOI, though.
Addressing the issue of when the crew transfers to the lunar lander. Is there much fuel penalty if the crew rides up in the command module (for safety’s sake), then does a rendezvous/docking and transfers to the lunar module while in LEO before TLI? If this action is practical, the CM could go on to L2 and the lander to the moon under the scenario that Jon describes and maybe this is what he had in mind.
I like the fact that under Jon’s scheme we have a single stage lunar lander that (theoretically) could be made to be reusable. Unlike the old Apollo or current Constellation plans with separate ascent and descent stages that allow only one use of the lander.
A bit OT, but does anyone know if ULA had ever considered a “Fat Atlas,” an “Atlas VI” if you will, that uses two RD-180’s?
Roderick,
An Atlas with dual RD-180s? You mean like their Phase-Two Atlas? Yeah, they’ve been trying to sell that one for a while. Their concept was to use the same tooling diameter as the DIV body (ie going up to ~5.4ish meters diameter) for both the first stage and for the Centaur. Single stick launch vehicle that could put something like 25-30mT into LEO. Me, I think it’s overkill. Any of the Phase 1/2/3 upgrades would make this whole concept a lot easier, but I like the fact that you can still probably do some useful things even without the government explicitly shelling out money for ULA upgrades (if ULA is getting enough demand that they think they can get a better value by investing in those upgrades themselves…that’s a different story).
~Jon
Jon: It would make me happy if you were to apply an analysis like this to the prospect of visiting a near earth asteroid, say, perhaps, one of the Apohole class, or really, anything with a minimal delta-v off of
Dr Lance Benner’s list.
Perhaps a hybrid – manned refuel – instrument mission might make sense for a manned asteroid mission… You save a lot of money and mass on the lander, anyway…
NASA has a history of thinking big and wanting everything right from the beginning:
– Operational space shuttle providing airline like flights without any precursor vehicles
– Space Station Freedom providing a huge orbiting laboratory
– Ares V supporting America’s heavy lift needs for the next 30+ years
Shuttle never met expectations; ISS is finally nearing completion 25 years later at over $100B; who knows what our launch requirements will be 20-30 years from now?
Jon, what you are proposing provides a fast, innovative, cost effect path to renewed lunar missions. A sustainable program requires success within politicians short career. And as you stated, there are numerous enhancements that can be made to increase capability or reduce costs.
And arguably the use of depots puts America on a more “Direct†path to Mars than developing an SDLV heavy lift that will continue to tie NASA’s expenditures to a huge albatross called rocket infrastructure rather than allowing NASA to focus on ambitious exploration.
Dave,
Once you have the L1 or L2 depot, you can pretty much go anywhere with an architecture like this (combined with something like a Bigelow Sundancer module for the living quarters for longer duration flights). I’m not as familiar with NEO trajectories to be honest. It might be possible to use this for an NEO mission pre-depot, but I’m not really sure.
~Jon