Possibilities for a single launch architecture of the Artemis missions.

 Copyright 2022 Robert Clark

 In the blog post ESA Needs to Save NASA's Moon Plans I noted that the original plan SpaceX submitted to NASA for a lunar lander required 16 launches due to multiple refueling flights, with the refueling flights to orbit requiring a time of 6 months to accomplish. I argued in the blog that if instead NASA used an Ariane 5/6 as the upper stage of the SLS rocket replacing the current Interim Cryogenic Propulsion Stage(ICPS) then it could be done in just a single launch of the SLS, with no launches of the Starship required at all.

 After their proposal was submitted by SpaceX and accepted by NASA, Elon Musk, stung by the criticism it would take so many launches, suggested it probably could be done in only 4 refuelings since a stripped down Starship for a lunar lander mission would weigh much less.

 SpaceX needs to be open about what the mass would be for such a stripped down Starship since that would directly affect how much NASA, and the U.S. taxpayers, would have to pay to SpaceX for refueling launches. See discussion here, 

The nature of the true dry mass of the Starship. 

 My suggestion to use the Ariane 5/6 as an SLS upper stage was critiqued on political acceptability grounds for a such a large contract to be taken from a U.S. company and given to a European company. 

 Here I'll propose a solution using existing, pretty much, American upper stages for the SLS. It's the ULA Centaur V upper stage coming into service next year. I considered using the Delta IV common core stage but at a 40 meter height it might be too tall for this use.

 

Architecture.

 The Centaur V has a 54 ton propellant load. Following the approx. 10 to 1 gross mass to dry mass ratio of the original Centaur, I'll take the dry mass to be ~5 tons. Then I'll examine  two options: 1.)2 Centaur V's combined into a single stage, and 2.)2 separate Centaur V's.

 The current Block 1 version of the SLS gets about 27 tons to trans-lunar injection(TLI). This is the speed needed to get a spacecraft once in orbit to reach the Moon. The 27 tons is just enough to get the Orion capsule and its service module to TLI. 

 However, the current approach is not to put the Orion in low lunar orbit around the Moon. Instead, it will be placed in a higher altitude orbit of Earth-lunar space called a near-rectilinear halo orbit(NRHO). The reason is the current version of the SLS did not have enough power to put the Orion in low lunar orbit and for it to be able to escape again.

 Our plan then is to first increase the payload capacity of the SLS so that enough additional propellant can be given the Orion service module so the Orion can actually reach and leave low lunar orbit. 

 The Orion with its fully fueled service module has a mass of 26.5 tons. The propellant load of the service module is ~10 tons, with 16.5 total tons dry mass of the Orion and service module. We'll add an additional 10 tons propellant to the service module to bring the total mass to 36.5 tons, including 20 tons of propellant.

 The AJ-10 engine used has a vacuum ISP of 319s. We'll assume a lunar lander of size ~15 tons, comparable in size to the Apollo missions lunar lander. In a following blog post we'll describe it in more detail. So, 16.5 + 15 = 31.5 tons dry mass needs to be put in low lunar orbit.

 For the delta-v calculation, after the SLS places the Orion/Service Module/lunar lander stack in trans-lunar injection(TLI) towards the Moon, we need .9 km/s to put the stack into low lunar orbit. This requires 13 tons of propellant, leaving 7 tons remaining:

319*9.81Ln(1 + 13/(31.5 +7)) = .910 km/s. The lunar lander will then be launched to land on the Moon while the Orion and service module remain in lunar orbit.

 After the lander mission is completed, the lander returns the astronauts to the Orion in lunar orbit, and the lander is then jettisoned. The Orion's service module is then fired to bring the Orion back to Earth. After lander jettison, the dry mass of the Orion and service module will be 16.5 tons. Then the 7 tons of remaining propellant is sufficient to perform the trans-Earth injection(TEI) burn of 900 m/s to escape lunar orbit and place the spacecraft back onto the free return trajectory back to Earth:

319*9.81Ln(1 + 7/16.5) = 1,100 m/s.

Calculations for Earth escape stage to TLI.

 That's the plan if we can upgrade the SLS to carry sufficient payload to give the Orion service module that extra 10 tons of propellant. The total mass that needs to be put into TLI is 36.5 + 15 = 51.5 tons. Here's a calculation for the first approach of two Centaur V's combined into a single stage. I'll use the payload performance calculator of Dr. John Schilling, on Silverbirdastronautics.com. The specifications for the 5-segment SRB's are taken by scaling up the numbers from the 4-segment SRB's used on the Space Shuttle system.

 I'll give this stage 4 RL10 engines instead of the Centaur V's 2 because of the larger size, in effect just transferring two of the RL10's from the second Centaur's to the first. The input page looks like this:

                                                               

 The payload estimator then gives the payload to LEO of ~127 tons:

 

  And the for the payload to TLI we'll use a C3 of -1.00km2/s2.

 This gives a payload to TLI of about ~52 tons:

  It is notable though the Schilling payload estimator has rather large error bars. These numbers need to be confirmed by more accurate payload estimators.

 The payload can be increased by using instead of the RL10's, a single Blue Origin BE-3U, the vacuum optimized version of the BE-3 engine used on the New Shepard. This engine has a vacuum optimized thrust of 710 kilonewtons. Placing this in for the upper stage thrust gives a payload to LEO of 136 tons, and to TLI of 54.7 tons. Again this needs to be confirmed by more accurate payload calculators.

 The intent here is to find a low cost approach to an upper stage that would allow a single launch architecture for the Artemis lunar lander missions. A combination of adding additional engines and also combining two tanks would ratchet up the costs.

  The second approach would use two separate Centaur V's. However, because of the large mass that needs to be carried by the either Centaur as payload we'll give both Centaurs 4 RL10's. The input screen looks like this on the Schilling calculator:

  And the LEO payload is ~129 tons:

 And the TLI payload is ~54.7 tons:

  Again, these payload estimates would have to be confirmed by more accurate payload estimators.

 This second approach would not incur the extra costs of combining two Centaur V's into a single stage, but it would require 4 additional RL10's. As before though we could get increased payload by replacing the RL10's by the BE-3U, and likely lower cost.

 We still need to come up with that lunar lander of comparable gross mass as the Apollo lander, ~15 tons. In a following blog post I'll show our European partners can come up with such a lander at low cost and at a relatively short time frame.

  Robert Clark