Copyright 2024 Robert Clark
Congress is becoming increasingly concerned that with the continuing delays of the Artemis missions that China may beat the U.S. back to the Moon:
US must beat China back to the moon, Congress tells NASA.
By Mike Wall
'It's no secret that China has a goal to surpass the United States by 2045 as global leaders in space. We can't allow this to happen.'
https://www.space.com/us-win-moon-race-china-congress-artemis-hearing
I had previously proposed correcting an error in the design of Orion's service module that instead of making it larger than Apollo's service module because of Orion's twice larger size, instead made it 1/3rd smaller:
Possibilities for a single launch architecture of the Artemis missions, Page 2: using the Boeing Exploration Upper Stage.
The proposal was to give an additional approx. 10 tons propellant to the service module. This would allow the Orion capsule/service module stack plus an Apollo-size lander to be carried all the way to low lunar orbit, not just to NRHO(near-rectilinear halo orbit).
This though because of the higher payload may require use of the higher thrust J-2X engine on the Boeing EUS(upper stage) rather than the 4 RL-10 engines now planned on the SLS Block 1B. It's higher thrust would result in a greater payload to LEO and TLI, perhaps to ca. 120 tons to LEO rather than the 105 tons planned to LEO.
This approach requires additional propellant tanks be added to the service module and a change in the EUS upper stage engine to the J-2X. As I discussed in that blog post, it may also require an additional Centaur V sized third stage be added atop the Boeing EUS. This is dependent on what is the TLI(trans lunar injection) payload for the Boeing EUS using the J-2X engine. It may be it can perform the needed TLI payload without an additional Centaur V 3rd stage.
In any case, I'll propose here an alternative approach to a single launch Artemis architecture without increasing the service module propellant load. This again will use a light-weight Apollo-sized lander with all the components of Orion capsule/Service Module/lunar lander all carried on that one single SLS launch. Because of the lower propellant load on the service module though I'll also send it to NRHO instead of to low lunar orbit.
Note the NRHO was chosen by NASA as the orbital location because it has a lower delta-v requirement to get there than going to low lunar orbit. Here’s the the delta-v requirements:
The second group of delta-v’s shows the delta-v to NRHO as 0.45 km/s and the delta-v to and from the lunar surface from NRHO as 2.75 km/s, or 5.5 km/s round trip.
I’ve seen various numbers for the Orion and service module dry mass and propellant mass. I’ll use 16.5 total dry mass for the Orion+service module together, and 9 tons of service module propellant mass, but only 8.6 tons of this as usable propellant because of residuals.
Then we'll use 6 tons of Service module propellant to get the Orion/Service Module/lunar lander to NRHO after being placed on TLI trajectory by the EUS, for the 16.5 ton Orion/Service Module dry mass, and 15 tons gross mass Apollo-sized lander with 2.6 tons left over for the return trip.
We'll need every bit of performance to accomplish the mission within these constraints. So we'll assume we can get a 324 s Isp out of the storable propellant engines on the service module. This is higher than specified for the Orion service modules engines but is doable because of the storable propellant Aestus engine on the Ariane 5 EPS storable propellant upper stage which gets this vacuum Isp. We'll assume we can get this increased Isp by using a larger expansion ratio nozzle or even by swapping out the engine on the service module to use the Aestus engine. Then we get:
324*9.81Ln(1 + 6/(16.5 + 15 + 2.6 + 0.4)) = 510 m/s, or 0.51 km/s, sufficient for placing the stack in the NRHO orbit, where the 0.4 in the equation is for the unburnt residuals.
Then with the 2.6 tons usable propellant left over for the return trip, after the lander is jettisoned, we get:
324*9.81Ln(1 + 2.6/(16.5 + 0.4)) = 450 m/s, 0.45 km/s, sufficient for the Orion return.
To increase performance even more we may want to switch even to the RS-72 engine. This is a turbopump-fed storable propellant engine with a vacuum Isp of 340s. It achieves this by using a higher chamber pressure of 60 bar and higher nozzle expansion ratio of 300 to 1 than the Aestus engine. A turbopump engine also has lower residuals, typically less than 1%. A disadvantage is that pressure-fed engines are simpler with fewer moving parts, and so higher reliability, important for an engine to place the spacecraft in orbit and for leaving orbit.
Now for the ca. 15 ton gross mass lander, because of the higher delta- v needed from NRHO we’ll use hydrolox rather than storable propellant stage. The Ariane 4 H10 hydrolox upper stage had a 11.8 ton propellant mass and 1.2 ton dry mass. We’ll use a 2 ton dry mass of the crew module:
ORBITAL PROPOSES FUTURE DEEP SPACE APPLICATIONS FOR CYGNUS.
SPACEFLIGHT INSIDER
MAY 1ST, 2014
Orbital’s proposal, outlined in this PDF, involves docking a Cygnus spacecraft with Orion to serve as a habitation and logistics module on longer flights. For these missions, the re-purposed Cygnus would be called the Exploration Augmentation Module (EAM). With its current life support systems used to transport pressurized cargo and experiments to the ISS, Cygnus is stated as being already suitable for the long term support of a crew. While berthed to Orion, Cygnus could support a crew of four for up to 60 days. Cygnus also has the capability of storing food, water, oxygen, and waste and features its own power and propulsion systems. The EAM would utilize the enhanced configuration Cygnus, which will begin flying larger cargoes to the ISS beginning with CRS-4 in 2015. An even larger version is also being proposed, featuring a 4-segment pressurized cargo module.
Note though the phrasing here is ambiguous. The Cygnus capsule as used as a cargo transport to the ISS contains air, as it would have to for the astronauts at the ISS opening it to retrieve the cargo, but not life support systems. I'm inclined to believe for the usage cited in this article it would be taking life support from the Orion capsule. Then the calculations need to be made for how much mass it would take for life support, thermal management, consumables for an independent crew module.
Now for the delta-v calculation for our hydrolox lander, we'll assume we can match the max 465 s Isp of the RL-10 engine by giving the Ariane 4 upper stage engine a nozzle extension as used on the RL-10, then we get:
465*9.81Ln(1 + 11.8/(1.2 + 2)) = 7,000 m/s, 7 km/s. This is quite a bit higher than the 5.5 km/s needed for the round trip from NRHO to the lunar surface and back again. But it uses hydrolox propellant so needs extra mass for low-boiloff tech.
Low boiloff-tech and long duration hydrolox stages are an important enabling technology. ULA engineers and ULA CEO Tory Bruno have written about this extensively in regards to for example the proposed ACES derivative of the Centaur upper stage. Because of the prior research on low-boiloff tech, an operational version to be fielded in a short time frame to be used on the Artemis missions likely can be done.
This shows a single launch mission is doable if going to NRHO, but it is not my preferred plan. A complete orbit around the Moon at NRHO altitude takes about a week, and for the Orion capsule being at NRHO and not low lunar orbit, the lander's crew would have to remain on the Moon about a week before they could return to the Orion in the NRHO orbit. The landers crew module would have to be larger with heavier life support and consumables in this scenario.
If instead the Orion was at low lunar orbit it takes two hours to complete an orbit and the lunar lander could launch every two hours to rendezvous with the Orion.
Since the Orion's service module being given an insufficient propellant load is such an obvious design mistake, the preferred route to take would be to correct that error, thereby allowing the missions to take place from low lunar orbit instead of from NRHO.
Robert Clark
1 comment:
Vehicle characteristics must match mission requirements. Ideally the mission drives the vehicle design, although for Apollo, the reverse was true. But that is another story.
What you have described would indeed allow NASA's SLS/Orion to reprise Apollo missions, but that is no longer the mission at hand! The lunar poles were not within the reach of the Apollo/Saturn design, but that is where NASA (and everybody else) wants to go now, because of possibly-mineable ice.
It is the shortfall of SLS/Orion to even do Apollo missions, coupled with the more demanding polar mission, that drives the boondoggle approach of Gateway in that halo orbit, that in turn makes landers so hard to design by driving up the dV so high.
Further, SLS/Orion Block 1 is so deficient in capability that the halo orbit they are forced to use is unstable, with its apoapsis beyond the moon's Hill sphere. Its periapsis is also very far from the moon's surface,
which is what drives up the lander dV so high.
So not only are they required to send lots of tankers to the halo station to refill supposedly-reusable landers, they also have to refill the station propulsion, which must burn frequently, just so as not to lose the station. Tankers to the moon is a very expensive proposition indeed.
That blind corner is where NASA is, letting the vehicle drive the mission design, but to more challenging requirements than Apollo had to meet.
Corporate welfare politics and incompetent Congressional micromanagement is exactly why the vehicle is deficient in capability, and (!!!) won't let them fix this and do it right. -- GW
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