Waiting for a Pump-Fed Methalox Lunar Descent Engine

[jongoff]

And, if that isn't crazy enough, you could "exchange" landers, have one coming and one going, sharing an uncrasher stage's propulsion all at once on the same trip. No room at all for contingencies, however, for cargo/resupply this might be possible albeit quite a tour de force.

I don't know if I ever ran the numbers on that scenario, but it was an idea I mentioned at the end of this blogpost back in 2013:

http://selenianboondocks.com/2013/09/centaur-uncrasher-stages-for-simplified-lunar-landings/

One of my few regrets about running a startup is not having the time to run numbers on crazy ideas like this. Maybe some day we'll be profitable enough that I can afford hiring someone or sponsoring some grad students just to run simulations for me. :-)

~Jon

[TrevorMonty]

Liked your depot fiso talk, pity you run out of time for more Q&A.

When I did rough calculations on uncrasher stage (OTV might be better description) I was thinking of human landers. Your idea of using for small cargo landers has lots of possibilities. In some ways it is better than a XEUS lander as payloads only need to handle forces in one direction.
If lander attaches to OTV and it's LV via it's base all major forces are in downward direction. With XEUS launch and orbital burns are horizontal compared to landed position.

Having propulsion the lander can rendezvous with OTV after being dropped off by LV.

The OTV dry mass can be less than ACES US as it doesn't need to handle launch stresses with payload. Plus development costs are lot less than XEUS. Lander should be straight forward and low cost, especially for likes of Moon Express and Masten. With refuelling at LEO and EML1 a 30t OTV could land close to 20t.

[sevenperforce]

The ITS is planned to use 10-tonne metha(l)ox-gas thrusters for RCS, fed from the autogenous pressurization coolant loop of the Raptor.

The dev Raptor is smaller than the final design, at around 1000 kN SL thrust; I'd project around 1150 kN vacuum-optimized thrust.

If you built a methalox upper stage for Falcon Heavy based around two of the dev Raptors and gave it a few of those gas thrusters (dual thrust-axis lander design), then refueled it in cislunar space, it could perform most of the landing burn using its main engines and just use the gas thrusters for the last little bit of the burn. It would have no trouble carrying enough fuel along with it to return to lunar orbit, or even to an Earth entry trajectory.

[spacenut]

If worried about batteries, a landing could be done at the begining of the 2 week solar cycle, and lift off near the end, with limited battery needed.  It would allow for about a 1-1/2 week stay before getting back into lunar orbit and docking with mother ship or moon station. 

[Propylox]

The thesis of this discussion is that all the many designs which have been proposed recently for crewed lunar landers are hobbled by the lack of an appropriate engine for the braking phase of lunar descent.

Do you accept the premise that the "tyranny of the rocket equation" won't be broken soon? If so, your design for the descent burn just has have better efficiency (Isp) than hypergolic propulsion provides.

... So personally I envision a hypergolic ascent stage which provides abort to LLO as a descent contingency capability. This implies the descent stage isn't reused, but perhaps it could be repurposed on the lunar surface.

The braking phase of lunar descent isn't an issue, it's the landing and liftoff that is. Reliability and/or redundancy are paramount at this stage. Pressure-feed hypergolic aces this requirement, but they compromise isp (which can exceed 350sec) for thrust and vice-versa.
But a lower-thrust, pressure-feed hypergolic engine for orbital maneuvering (inclination, perigee, etc) as an auxiliary/emergency engine to the main engine of higher thrust/isp achieves all safety, thrust and isp needs. These engines are ready with minimal development.

Secondly, I reject any proposal that abandons its descent stage and engines. The propellant required to send a new descent stage from Earth to LLO far exceeds the propellant required to return that stage from the Lunar surface to LLO.

Both ideas combined; My "design" (which you asked for) is a fully reusable propulsion module composed of two engines, replaceable main tanks (and baffles for the hypergolics) to be attached to any habitat or downmass. This module would be docked, outfitted, refurbished and refueled in LLO before mounting to whatever its mission requires it carry. (note: landing legs are part of and specific to the habitat or downmass, not part of the propulsion module)

[Steven Pietrobon]

Secondly, I reject any proposal that abandons its descent stage and engines. The propellant required to send a new descent stage from Earth to LLO far exceeds the propellant required to return that stage from the Lunar surface to LLO.

You're forgetting the extra LM propellant and the extra stage mass required to hold the extra propellant that you need to send from Earth. That extra two or three tonnes of stage mass that you "save" from not sending from the Earth requires an extra 10 to 15 tonnes of propellant that you need to send from Earth into low Lunar orbit (LLO). That is the extra amount of propellant that is required to reuse the whole LM (plus you now got the problem of your ascent engines possibly being damaged during landing). At say $25,000 a kg into LLO, that works out to $200M to $300M extra that you need to pay. That's probably about the same price for a separate descent stage. Just reusing the ascent stage makes more sense and should make operations easier, since you just plugin in a new descent stage with its payloads to your ascent stage and have the descent stage refuel the ascent stage.

[Propylox]

Q: Why the fascination with Methalox for lunar operation considering it's poor thermal properties requiring boil-off mitigation? There are other options, some with equivalent isp and better density.

One technology that is worth keeping an eye on is battery powered, electric pump fed engines such as the Rutherford ...

Do any of you really think a pump (or multitude of pumps) not part of a powerhead could reliably deliver enough propellant at pressure to achieve high isp and +80kN thrust for a lander? Wouldn't all the pump and/or battery hardware outweigh, literally and metaphorically, a pressure or pre-burner engine accepting boil-off?
For example: About four Rutherfords (electrically driven) or four XCOR engines (their thermodynamic piston pumps appear capable of similar thrust) could power a lander, but why not keep it light, simple and reliable with a 4-chamber RD-58 similar to Yuzhmash RD-8s? (not proposing Kerosene, just comparing engines)

You're forgetting the extra LM propellant and ...

No, I did not and Yes, I can add.
Different lander and mission designs will give different results, but mine works out as every kg abandoned on the surface, needing replacement, only saves about 0.9kg of initial propellant. That 10% difference is what I called "significant" as it works back the chain to Earth's surface. (for reference; About 46% of my initial mass in LLO is landed and of that about 46% is return fuel when using a fully reusable lander)