Possibilities for a single launch architecture of the Artemis missions, Page 2: using the Boeing Exploration Upper Stage.

 Copyright 2023 Robert Clark

  A comparison between the Apollo and Orion capsules:

 Rarely has a design mistake been so clearly illuminated by a single picture. Note the Orion capsule is nearly double the size of the Apollo capsule in mass. But rather than making Orion’s Service Module twice as big as the Apollo Service Module, as it should be to get similar performance, instead it is smaller by 1/3rd.

 Orion’s service module is based on ESA’s ATV cargo tug to the ISS, which had a 4.5 meter diameter and a 10 ton propellant load.

BUT THERE WAS NO REASON TO KEEP IT AT THAT SAME DIAMETER FOR THE ORION USE, NOR TO KEEP THE SAME SIZE PROPELLANT LOAD.

 If instead the diameter was made to match the capsule’s diameter, as was the case with Apollo, there would be an additional 20 cubic meters of volume inside the Service Module, well more than enough to hold an additional 10 tons of the storable propellant used.

And that is all that is needed to solve THE major problem of the SLS/Orion approach: the fact it can’t send the Orion and a lunar lander to low lunar orbit, and bring the Orion back to Earth again.

 It is because of that the idea of the lunar Gateway was proposed, where the SLS would only have to take the Orion to a further out orbit.

 But if instead the Service Module was given that additional 10 tons of propellant then it could send both the Orion and a ca. 15 ton lunar lander to low lunar orbit, and have enough propellant left over to bring the Orion back to Earth, a la the Apollo architecture.

 Rarely, has a mistake been so clearly exposed, especially when its solution is so clearly made apparent as well.

 In the blog posts, "ESA Needs to Save NASA's Moon Plans", and "Possibilities for a single launch architecture of the Artemis missions", I wrote about getting a single launch format for the Artemis lunar lander missions by using the Ariane 5 as an upper stage or by using two Centaur V stages as the upper stage for the SLS, respectively.

 This stemmed from dislike of the plan NASA was endorsing of using multiple flights and refuelings of the SpaceX Starship as the lander. I also objected to the high cost projected for the planned Boeing Exploration Upper Stage(EUS), being nearly half the cost of the entire SLS per flight, nearly $1 billion.

 However, NASA has negotiated a better price structure for the EUS. And it appears NASA is wedded to the Boeing EUS. Then I'll discuss a single launch architecture using the Boeing EUS upper stage.

 The payload to LEO of this version of the SLS with the Boeing EUS, which is version Block 1B, will be 105 tons to LEO. The current fueled mass of the Orion+Service Module is 26.5 tons. An additional 10 tons of propellant will bring it to 36.5 tons. 

 In the blog post, "A low cost, lightweight lunar lander", I discussed a lunar lander at a 13-ton total fueled mass based on the Cygnus capsule given life support as the crew module, and the Ariane 5 EPS storable propellant stage as the propulsive stage for the lander.

Calculations for the delta-v to the Moon and back.

 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 fueled lunar lander of size ~13 tons, as described in the blog post, "A low cost, lightweight lunar lander", comparable in size to the Apollo missions lunar lander. So, a 16.5 + 13 = 29.5 ton mass for the vehicles that need to be put in low lunar orbit. But remember also we need to have some propellant left over in the service module to bring the Orion back home to Earth.

 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 0.9 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/(29.5 +7)) = 0.950 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. 

 So the total mass that needs to be sent to trans lunar injection(TLI) on a path to encounter the Moon is 36.5 + 13 = 49.5 tons. Now use the rule-of-thumb that a Centaur-like hydrolox stage can send to TLI at a 3,000 m/s required delta-v a payload mass equal to its propellant load.

 So use for a third stage atop the Boeing EUS the Centaur V at an 50 ton propellant load and 5 ton dry mass. This then results in a total mass to LEO of 104.5 tons consisting of the 55 tons of the Centaur V plus the 49.5 tons of the Orion capsule/Service Module/lunar lander, within the lift capacity of the SLS Block 1B to LEO.

  Robert Clark