A low cost, lightweight lunar lander.

 Copyright 2022 Robert Clark

 In the blog post Possibilities for a single launch architecture of the Artemis missions I discused that a single launch architecture is possible for the SLS rocket if there is a lightweight lunar lander. Such is possible using currently existing space stages. Firstly, a lunar crew module of ca. 2 ton mass is possible based on Orbital Sciences Cygnus capsule, discussed in Budget Moon flights: lightweight crew capsule. The Cygnus is actually build in Italy by Thales Alenia Space. As it is already built, the additional modications for added life support would be comparatively low cost.

 Note Thales Alenia Space is already adding life support sysmtems to a larger version of the Cygnus for the lunar Gateway. Then a lower cost version would simply use the smaller Cygnus itself, at a ~2 ton dry mass for a short term stay on the lunar surface.

 As for the propulsion system, the earlier Ariane 5 EPS storable propellant upper stage prior to the current cryogenic upper stage could be used for the purpose: it had a 1.275 ton dry mass and 9.750 ton propellant mass, for a 11 ton gross mass. Then the crew module and propulsion stage would mass 13 tons.

 An advantage over the SpaceX Starship lunar lander plan is that it is only 3 meters high, the same height as for the Apollo lunar lander descent stage, making it easy for the astronauts to climb down to the lunar surface, compared to the 25 meter height for the Starship.

 Calculation.

The delta-v to the lunar surface from low lunar orbit is 1,870 m/s:

http://en.wikipedia.org/wiki/Delta-v_budget#Earth.E2.80.93Moon_space

 The Aestus engine on the stage has a Isp of 324s. Then the delta-v it could achieve carrying a 2 ton crew module would be:

324*9.81Ln(1 + 9.75/(1.25 + 2)) = 4,400 m/s. Then it could work as single stage to go down to the lunar surface from low lunar orbit and back again.

 It is notable a stage derived from the Space Shuttle OMS pods would also have this capability.

 The specifications for the OMS pods are given here:

SHUTTLE PERFORMANCE ENHANCEMENTS USING AN OMS PAYLOAD BAY KIT 1991

https://cp3.irmp.ucl.ac.be/~ringeval/upload/spaceshuttle/PBK/shuttle_perf_with_omskit.pdf

The middle size version has a dry mass of 3,955 lbs, 1,800 kg, and propellant load of 25,064 lbs, 11,400 kg. It has an Isp of 316 s. Then with a 2 ton crew module it would have a delta-v of 4,300 m/s:

316*9.81Ln(1 + 11.4/(1.8 +2)) = 4,300 m/s, sufficient for single-stage lunar lander.

  Robert Clark

SLS for Return to the Moon by the 50th Anniversary of Apollo 11.

Copyright 2012 Robert Clark

Credit: NASA's Space Exploration Vehicle (SEV).

Very interesting report about using NASA's proposed Space Exploration Vehicle for cislunar space exploration:

Lunar Surface Access from Earth-Moon L2.
A novel lander design and study of alternative solutions.
1 October 2012 | Washington, DC

http://www.sei.aero/eng/papers/uploads/archive/SEV-L2-Lander-Presentation_1Oct2012.pdf

 The report proposes using the lightweight SEV, at only a 3 mT empty weight, and all cryogenic propulsion as a shuttle between the L2 space station NASA has recently discussed and the lunar surface. However it could also be used as the crew capsule between LEO and the Moon's surface.
 The architecture discussed is very interesting in that the SEV would be used as the single crew module to carry the crew all the way from the L2 station to the lunar surface and back again, i.e., no separate lander crew module. There would also only be a single propulsive stage to carry the SEV from low lunar orbit to the lunar surface and back to lunar orbit, i.e., no separate lunar descent and ascent stages.
 This has similarities to the architecture for the Early Lunar Access(ELA)[1] proposal of the early 90's. This also used all cryogenic space stages to save weight, only 52 mT required to LEO. ELA also saved weight and cost by using a single crew capsule for the entire flight from LEO to the lunar surface and back again. It also used a single propulsive stage for lunar descent and ascent. But instead of linking up with a stage waiting in lunar orbit for the return, the ELA proposal was to have this single lander stage return all the way back to LEO.
 An alternative architecture discussed on page 23 in this report on using the SEV for cislunar travel does not use the method of first stopping in lunar orbit, then having a separate lunar lander stage. Instead it uses the "direct descent" method of descending directly to the lunar surface. This landing method is analogous to that used in the ELA proposal to save propellant. Interestingly the SEV report on page 23 gives the delta-V for the direct descent method as 2,610 m/s. This compares to the 760 m/s + 2,150 m/s = 2,950 m/s for the method that first stops in lunar orbit, then descends to the surface as indicated in the image above. So according to this report a savings of 300 m/s in delta-V for the trip from L2 to the Moon is possible using direct descent, a significant savings.
 I had wondered if it was possible to save delta-V and propellant in this blog post 'Delta-V for "direct descent" to the lunar surface?'[2]. The SEV report suggests it may be possible to save in the range of 300 m/s by the direct descent method.
 The only technical complaint raised against the feasibility of the ELA proposal back in the 90's was the suggestion of getting a 2-man crew capsule at only a 3 mT empty weight. So the fact the SEV is expected to have this low an empty weight is important, since it suggests the possibility with just the 70 mT first version of the SLS of a manned lunar lander mission using currently existing cryogenic stages.
 Actually the 70 mT payload of the SLS is so much better than the 52 mT needed for ELA that likely we could even use a heavier hypergolic stage for the lunar ascent stage. During the early planning of the Apollo program when the possibility an engine might not ignite was regarded as a definite possibility, it was decided to use hypergolics, which ignite on contact, for the lunar ascent stage. At this point though the cryogenic RL10 engines have had decades of use and are regarded as highly reliable.
 Still for these first versions of these new lunar landers we might still want the certainty of using hypergolics for the ascent stage. I suggest using the engine and propellant tanks of the shuttle orbiter OMS pods for the purpose. This would be quite appropriate actually since the OMS pod engines were derived from the Apollo lunar lander engines. By the Astronautix page on the OMS pods[3], they are each about 10 mT propellant mass and 1.8 mT dry mass. Then using its 316s Isp, one of them would suffice for the ca. 2,740 m/s delta-V to go from lunar surface to LEO even with a 4 mT crewed and supplied mass for the SEV with plenty of margin: 316*9.81ln(1 + 10/(1.8 + 4)) =  3,100 m/s.
 The first version of the SLS, called Block 1, is expected to launch by 2017. I would expect a test lunar lander mission, especially if using all cryogenic in-space propulsion, to be done first before a crewed mission is sent. But certainly by 2019, the 50th anniversary of Apollo 11, a crewed mission could be sent. This is in contrast to a post-2030 proposed time frame for a crewed lunar landing using the full 130 mT version of the SLS when it first becomes available.
 There is the cost issue of mounting a manned lander mission. Oddly, the high cost of the SLS might be helpful in this regard. The cryogenic Centaur-like upper stages are already available at a cost in the range of $30 million [4], so the modifications there would be comparatively low cost, compared to the already high cost of the SLS. As for the development cost of the SEV, I suggest use of NASA's commercial crew program's financing procedures. SpaceX was able to develop the Dragon as largely privately financed for reportedly $300 million. And Boeing is paying much of the cost of the development of the CST-100 capsule. It is highly dubious they would be spending a billion dollars of their own money for its development. Then likely its total development cost is in the few hundred million dollar range. Therefore it is likely the development cost of the smaller SEV under commercial crew procedures would also be in the few hundred million dollars range, again comparatively low cost compared to the SLS.
 As I discussed in the blog post "SpaceX Dragon spacecraft for low cost trips to the Moon", SpaceX will also be able to mount a manned lunar landing mission using the 53 mT Falcon Heavy by following, it turns out, the ELA architecture. This will be much cheaper than using the SLS launcher, perhaps only in the few hundred million dollars range cost. But you would have to get private financing for that, since NASA would not fund it as it would undercut NASA's own program.
 In contrast, NASA using the SLS in such an early time frame for a manned return to the Moon would provide further support for continuing the SLS funding. No longer would the SLS be referred to as "a rocket to nowhere".


  Bob Clark

Update, Sept. 28, 2013:

 Finally, NASA has acknowledged that the Block 1, first version of the SLS to launch in 2017 will have a 90+ mT payload capacity not the 70 mT always stated by NASA:

SLS Dual Use Upper Stage (DUUS).
http://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20130013953_2013013757.pdf

 This is important since it means we will have the capability to do manned lunar landing missions by the 2017 first launch of the SLS:

SLS for Return to the Moon by the 50th Anniversary of Apollo 11, page 5: A 90+ metric ton first launch of the SLS.
http://exoscientist.blogspot.com/2013/09/sls-for-return-to-moon-by-50th.html

REFERENCES.

1.)Lunar Base Studies in the 1990s. 
1993:  Early Lunar Access (ELA). 
by Marcus Lindroos 
http://www.nss.org/settlement/moon/ELA.html 
(Note a typo on this page: the payload adapter mass should 
be 2,000 kg instead of 6,000 kg.) 

2.)Delta-V for "direct descent" to the lunar surface?
SATURDAY, SEPTEMBER 15, 2012
http://exoscientist.blogspot.com/2012/09/delta-v-for-direct-descent-to-lunar.html 

3.)Encyclopedia Astronautica.
Shuttle Orbiter OMS.
http://www.astronautix.com/stages/shueroms.htm

4.)Encyclopedia Astronautica.
Centaur IIA.
http://www.astronautix.com/craft/cenuriia.htm