Exoscientist

The fueled Orion/service module weighed about 26.5 tons: Specifications Spacecraft type Crewed Launch mass CM: 22,900 lb (10,400 kg) ESM: 34,085 lb (15,461 kg) Combined mass: 58,467 lb (26,520 kg) Total with LAS: 73,735 lb (33,446 kg) Dry mass CM: 20,500 lb (9,300 kg) landing weight ESM: 13,635 lb (6,185 kg) https://en.m.wikipedia.org/wiki/Orion_(spacecraft) Then with 10.7 tons extra propellant and tankage and a 15 ton lunar lander the total mass needing to get to TLI is: 26.5 + 10.7 + 15 = 52.2 tons . Perhaps the higher 60 ton you’re getting comes from including the Launch Abort System(LAS)? But that is jettisoned well before reaching orbit . Robert Clark

Not exploding is an important requirement for a rocket, especially for one intended to carry crew. For the alternative proposal using a hydrolox stage for the propulsion of a lunar lander, this propulsive stage already exists and been flown multiple times for the upper stage of the Ariane 4 rocket. For the Cygnus capsule, it has been flown multiple times as a unmanned cargo craft and is planned to be a manned component of a space station. As I mentioned though I’m not certain it has it’s own independent life support systems. The Cygnus itself weighs 2, 000 kg. I don’t think adding life-support would add too much more to that mass. Robert Clark

Your estimated extra propellant of 6.5 tons required might not be including the required propellant to also get the Orion back to Earth. I estimated 10 tons extra propellant required. But anyway, this latest blog post was to keep the current size service module, still stage from NRHO not low lunar orbit, but use a hydrolox lander to deal with the larger delta-v needed for staging from NRHO. The proposal would still be heavier to TLI than before though. The Orion/service module weighs 26.5 tons. A 15 ton lander would bring the TLI mass to 41.5 tons. I made two suggestions to deal with this added mass to TLI: adding a Centaur V at 50-ton size as a 3rd stage or going back to the J-2X engine for the Boeing EUS upper stage that had been planned. A preliminary calculation suggest they should be able to do > 42 tons to TLI. A ca. 2 ton mass lunar crew module should be able to do over a week stay on the Moon based on the Cygnus module. The uncertainty is in whether the usage described in the article below includes independent life-support on the Cygnus or if it is taking life-support from the Orion: 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. https://www.spaceflightinsider.com/missions/commercial/orbital-proposes-future-deep-space-applications-cygnus/ As to why this approach, the Starship HLS is still uncertain for an Artemis III or Artemis IV timeline. Even if SpaceX succeeds at flights to LEO, they still need to demonstrate reusability and orbital cryogenic refueling. This approach still allows sustained lunar presence because it would not use SLS for cargo and habitat delivery. That would use far cheaper commercial launchers instead. Bob Clark

This earlier post expanded out: Possibilities for a single launch architecture of the Artemis missions, Page 4: lightweight landers from NRHO to the lunar surface. https://exoscientist.blogspot.com/2024/01/possibilities-for-single-launch.html A major flaw in the design of the SLS was that the Orion capsule was saddled with a too small service module: Rather than giving Orion a larger service module than Apollo’s due to its twice larger size, it was instead given one 1/3rd smaller. Why? It stems from the earlier, now cancelled Constellation program. Constellation was to have rocket the Ares 1 to carry astronauts to LEO in the Orion capsule, and a larger launcher the Ares V to carry the other lunar mission elements to rendezvous with the Orion in LEO for the flight to the Moon. However, the Ares I was powered by a single SRB as the first stage and that did not have enough power to get a large service module and the Orion to LEO. Then the service module was cut down in size. But when Constellation was cancelled it was decided to keep the Orion, and with that its service module also. But now the decision to use a smaller service module made no sense. The Ares I was cancelled so it made no sense to use limitations on the size of the service module coming from a rocket that wouldn’t be used. But government programs have inertia. It’s hard to change programs already decided on. It becomes like they are fixed in stone. Since the Orion capsule is kept, its service module is kept as well. It would be a relatively easy fix, both technically and financially, to increase the size of the service module propellant tanks so that the service module could carry the Orion and an Apollo-sized lunar lander to low lunar orbit with enough propellant left over to carry the Orion back to Earth again. The only problem is convincing NASA to do it. It’s because of that too small service module that NASA had to come up with other destinations at the Moon that the Orion could reach and come back from. Thus was selected the NRHO orbit(near rectilinear halo orbit). Since the NRHO is decided upon, could we get a lunar lander that could be used on the SLS to get a single launch lunar landing architecture that staged from NRHO? The answer is yes. If we made it an Apollo-sized lander but hydrolox powered instead of using storable propellant engines. The hydrolox lunar lander would need low-boiloff tech. But ULA has done a lot of research on this in regards to their ACES(Advanced Cryogenic Evolved Stage) so the tech likely could be used in the near term for a lunar lander. Bob Clark

Congress becoming concerned with the frequent delays of the Artemis program, that China may beat us back to the Moon: US must beat China back to the moon, Congress tells NASA. By Mike Wall published 3 days ago '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 Bob Clark

Team Reveals Remaining Asteroid Sample Johnson Space Center Office of Communications JAN 19, 2024 A top-down view of the OSIRIS-REx Touch-and-Go-Sample-Acquisition-Mechanism (TAGSAM) head with the lid removed, revealing the remainder of the asteroid sample inside. Photo: NASA/Erika Blumenfeld & Joseph Aebersold https://www.nasa.gov/image-article/nasas-osiris-rex-curation-team-reveals-remaining-asteroid-sample/ Bob Clark

You could always use his NotAFlamethrowerTM Robert Clark

Odd they didn’t just partially fill the tanks. Bob Clark

Toby Li @tobyliiiiiiiiii Jan 6 Elon Musk to provide 2024 Starship Update on January 11. Confirmed on an X livestream, @elonmusk announced he will conduct a SpaceX company talk including a Starship update next Thursday. I'm hoping to hear details regarding Starship IFT-2's post-flight analysis, plans for IFT-3 & beyond, and updates on Starship HLS milestones. https://twitter.com/tobyliiiiiiiiii/status/1743529920322846812 This is supposed to be an internal speech to SpaceX employees, but in the past was released to the public. Robert Clark

Thanks for that. I can tell now when I am copying something from another forum with different formatting when it becomes unreadable on this forum when reading in dark mode. By the way, I just realized that Elon tweet I copied from twitter where he says the expendable Starship might have a dry mass of only 40 tons, has that same problem. But you can read the tweet by clicking the link to the twitter post. Bob Clark

How does this look? ——————————————————————————————————————————- I have been accused of being anti-SpaceX because of my criticism of the Starship. Actually, after a calculation I'm convinced the Starship can be operational, like, tomorrow, with relatively small design changes: Towards advancing the SpaceX Starship to operational flight: SpaceX should lower the Raptor chamber pressure and thrust level. https://exoscientist.blogspot.com/2024/01/towards-advancing-spacex-starship-to.html … ————————————————————————————————————— I applied the option the forum offers of inserting text in plain-text format. Robert Clark

I frequently copy and paste from other forums which may introduce new fonts this forum is not familiar with. When putting them together I frequently use the “Note” app on the Apple iPad. That might introduce other fonts. Here’s an experiment: I’ll try using the “Pages” app on the iPad instead which allows you to select the font. So here’s the previous post you couldn’t read in “dark mode” but copied into “Pages” and put in Arial font then recopied onto this forum: _____________________________________________________________________________________________________ I have been accused of being anti-SpaceX because of my criticism of the Starship. Actually, after a calculation I'm convinced the Starship can be operational, like, tomorrow, with relatively small design changes: Towards advancing the SpaceX Starship to operational flight: SpaceX should lower the Raptor chamber pressure and thrust level. https://exoscientist.blogspot.com/2024/01/towards-advancing-spacex-starship-to.html The Raptor engine has shown continued failures on all of test stands, Starship low altitude landing test flights, and the two orbital test flights. But the Raptors on the booster on the last test flight were able to complete the ascent part of the flight without failures. They failed only after the attempted to relight. Multiple-lines of evidence suggest that on that last test flight SpaceX throttled down the Raptors on the booster to less than 75% while those on the Starship were run at ~90%. I've suggested this is why the booster engines were able to fire reliably during the ascent and those on the upper stage were not. If this is the case, then it suggests a method to get Raptor reliability: run them at ~75% throttle on both stages. But if keeping the same stage dry masses this would result in the payload of the reusable version being reduced to approximately in the range of 100 tons from 150 tons. Instead, I advise first start with reducing the dry masses by optimally lightweighting the expendable versions of both stages. Surprisingly this gives a greater expendable payload than the expendable payload of the current version. Secondly, I suggest using winged, horizontal approach to reusability gives a much reduced payload loss due to reusability. Thirdly, basic orbital mechanics shows high delta-v missions such as to the Moon or Mars are done more efficiently by using more stages. Then a third stage is suggested for the Superheavy/Starship, a mini-Starship as it is called by Robert Zubrin. This allows single launch and fully reusable missions to the Moon or Mars. No refueling flights required. _____________________________________________________________________________________________________ Let me know if this is legible. By the way, about your question I wanted to do the optimal lightweighting for both the Starship and SuperHeavy as expendables. The savings in weight would be more extreme in the case of the Starship though. My entire argument is based on this one Elon tweet: Elon Musk @elonmusk Probably no fairing either & just 3 Raptor Vacuum engines. Mass ratio of ~30 (1200 tons full, 40 tons empty) with Isp of 380. Then drop a few dozen modified Starlink satellites from empty engine bays with ~1600 Isp, MR 2. Spread out, see what’s there. Not impossible. 9:14 PM · Mar 29, 2019 https://x.com/elonmusk/status/1111798912141017089 The difference of 80 tons between this Elon estimated dry mass of 40 tons for the expendable Starship and the current reusable dry mass of 120 tons is *huge*, especially for an upper stage where the difference subtracts directly from the payload mass. Robert Clark

I’m responding to this comment from the Starship discussion here since my response is more general in nature on lunar missions. I don’t agree you need rockets giving 70 tons to TLI to get sustainable architecture allowing manned lunar surface stations, a la how the ISS is in low Earth orbit. As I mentioned before the SLS is too expensive, upwards of $4 billion per flight, to be used for cargo only missions. Better to use far cheaper commercial flights for that purpose. Robert Zubrin gave a plan for producing a Moon base using three launches of Falcon Heavy plus a launch of the Falcon 9 to carry the crew to LEO: Op-ed | Moon Direct: How to build a moonbase in four years Robert Zubrin March 30, 2018 https://spacenews.com/op-ed-moon-direct-how-to-build-a-moonbase-in-four-years/ Key for his plan is using a “Lunar Excursion Vehicle” (LEV) of ca. 12 ton gross mass that is hydrolox powered as a lunar lander. This would require near zero-boiloff tech, but Zubrin thinks this is doable with current tech. For the two cargo missions launched by Falcon Heavy, Zubrin would use a larger hydrolox stage for the lander at ca. 40 ton size, a bit smaller than Centaur V, that could deliver ca. 12 tons of cargo to the lunar surface. Quite notable about his plan is as far as the manned flights it would require a single Falcon Heavy rocket as the launcher to get the hydrolox in-space stages to LEO. The FH has a payload capacity of 63 tons to LEO. So an only 63 ton launcher could get the required in-space stages to LEO, which could then do a manned round trip flight to the lunar surface. This small size for the launcher is coming from the fact the in-space stages are so much lighter being powered by hydrolox. The Falcon Heavy however is not manrated though. So it would need a separate man-rated launcher to get the astronauts to LEO. This man-rated launcher could be the Falcon 9 but it doesn’t have to be. Robert Clark

I have been accused of being anti-SpaceX because of my criticism of the Starship. Actually, after a calculation I'm convinced the Starship can be operational, like, tomorrow, with relatively small design changes: Towards advancing the SpaceX Starship to operational flight: SpaceX should lower the Raptor chamber pressure and thrust level. https://exoscientist.blogspot.com/2024/01/towards-advancing-spacex-starship-to.html The Raptor engine has shown continued failures on all of test stands, Starship low altitude landing test flights, and the two orbital test flights. But the Raptors on the booster on the last test flight were able to complete the ascent part of the flight without failures. They failed only after the attempted to relight. Multiple-lines of evidence suggest that on that last test flight SpaceX throttled down the Raptors on the booster to less than 75% while those on the Starship were run at ~90%. I've suggested this is why the booster engines were able to fire reliably during the ascent and those on the upper stage were not. If this is the case, then it suggests a method to get Raptor reliability: run them at ~75% throttle on both stages. But if keeping the same stage dry masses this would result in the payload of the reusable version being reduced to approximately in the range of 100 tons from 150 tons. Instead, I advise first start with reducing the dry masses by optimally lightweighting the expendable versions of both stages. Surprisingly this gives a greater expendable payload than the expendable payload of the current version. Secondly, I suggest using winged, horizontal approach to reusability gives a much reduced payload loss due to reusability. Thirdly, basic orbital mechanics shows high delta-v missions such as to the Moon or Mars are done more efficiently by using more stages. Then a third stage is suggested for the Superheavy/Starship, a mini-Starship as it is called by Robert Zubrin. This allows single launch and fully reusable missions to the Moon or Mars. No refueling flights required. Robert Clark

On the static stand you can gradually build up to full thrust and full flight duration. If you see out of nominal state for one or more engines, you can shut the test down, examine the engines not operating properly and compare to the ones that are. Then try again gradually building up to full flight conditions. In contrast if one or more engines fail in flight, even if it didn’t explode but had to be shutdown, usually you have to destroy it by FTS for fear it will stray too far out of the safety zone. Robert Clark

Consider, that is the standard industry practice. It is not just SpaceX and Nasa doing in that way in the past. Every company in the industry does it that way. Every company could just test single engines on test stands if they wanted to. That would be much cheaper. Instead they follow the accepted practice of doing full flight analog testing prior to test flights. That the Soviet N-1 experienced not just engine shutdowns but engine explosions on every test flight and the SuperHeavy/Starship also experienced the same thing, including the engine explosions, suggests this is not a better approach to getting reliable engines. Robert Clark

You don’t have to take the video authors word on this. You can judge for yourself if there is a most common length over which most of the tests are done: Robert Clark

I have discussed this. It is standard industry procedure to do full thrust, full up(all engines), full flight duration(actual mission length) static fire tests before doing flight tests. Even SpaceX did this when qualifying the Merlins on the Falcon 9 for flight: SpaceX Testing - Falcon 9 Engine Test 147,136 views Nov 25, 2008 The full mission duration test firing of the Falcon 9 first stage lasting nearly 3 minutes. The nine Merlin engines produced 855,000 lbs. of thrust and consumed over half a million pounds of liquid oxygen and rocket grade kerosene during the test. And this video shows this being done for the Saturn V first stage: Saturn V S IC Static Firing (archival film) 34,163 views Aug 15, 2018 This film provides an overview of testing of the Saturn V first stage, the S-IC stage at NASA's Marshall Space Flight Center during the Apollo Program. The five F-1 engines which made up the stage generated a combined seven and a half million pounds of thrust. Such tests are more challenging for the engines and the stages. But that is the point. It is because such tests were done that the Falcon 9 and the Saturn V were spectacular successes. Bob Clark

The video argues the Raptor has high reliability based on the tests on static stands at McGregor. The author says the reliability is high because “most” tests were able to reach a planned length of 115 to 120 seconds. The problem is for a rocket engine to be used to power a crewed vehicle you want very high reliability. For instance the SLS has estimated reliability for its components of 99.9% and for the Merlins based on the number of successful flights we can estimate it as better than 99.9%. That is, less than 1 in a thousand would be expected to fail. But going by counting the number of tests for the Raptor that fail to reach that 115 to 120 second mark, it may be 1 in 5 to 1 in 6 fail to reach it. Note as the author of the video observes some tests are planned to be shorter. For some for instance they were intended to be about 47 seconds long. But there are a block of tests I marked off in the attached image that appear to be aiming for that 115 to 120 second mark, and several of them don’t make it. I estimate 5 or 6 out of the 30 I marked off failed to reach that planned burn length. Another questionable issue of these static tests is the planned lengths. The largest portion them were of a planned length of about 120 seconds, 2 minutes. But judging by the two test flights the actual burn time for the booster is in the range of 2 minutes 39 seconds to 2 minutes 49 seconds range. Only very few of the test stand burns went this long or longer. The video gives a link where you can watch the test stand burns NSF.live/McGregor. Another useful aspect here is you may be able to judge the power level of the burns. There is a graphic that shows the sound level of the burns. From that you may be able to judge whether or not the engines were firing at or close to full thrust. In the image below, the burns in white are those shorter burns of about 47 second lengths the author of the video made note of. They may be tests of the boost back or landing burns. The ones I’m commenting on are under the yellow bar, which I estimate to be at about the120 burn time. There 5 or 6 out of 30 don’t reach the planned burned time.  Robert Clark

This video attempts to argue the Raptor is reliable by looking at static fire tests. A key question in the video is how many of the test fires fall short of the expected length, suggesting a test failure or for whatever reason the engine had to be shutdown. The video host suggests it is small number. I suggest it is more than it should be for an engine at this stage in its development. Someone may want to count the percentage themselves where it is displayed graphically at about the 5:55 point in the video. Also I don’t like the lengths of the engine tests. The video host says most are about 2 minutes, 120 seconds, suggesting that is where SpaceX thinks that is what the flight burning time should be. But judging from the test flights for the booster that should be in the 2 minutes 40+ seconds range, 160+ seconds, which means even for static fires of individual engines SpaceX is not doing realistic static fires: Planned mission timeline[34] …. 00:00:55 Max q (moment of peak mechanical stress on the rocket — Later and lower Max q than planned[42] 00:02:49 Main engine cutoff (MECO) — Not attempted[43] 00:02:52 Stage separation … https://en.m.wikipedia.org/wiki/SpaceX_Starship_integrated_flight_test_1#Flight_profile. Additionally, SpaceX needs to be open about how many of these static fires are done at full power. Robert Clark

I have mentioned this before. I major irritation of mine is that SpaceX dismisses the lessons of Apollo. It dismissed the importance of a flame trench and dismissed the importance of having powered stage separation. But the most egregious of these is dismissing the importance of having full thrust, full up(all engines), full mission duration(actual minutes long flight length) static burns: Robert Clark

Yes. As long as Raptors fail in flight tests, there will be questions about its reliability. Bob Clark I don’t hold in stock with any companies, including any aerospace companies. I don’t like the Blue Origin lunar lander plan or the third one either, whose name I can’t now remember. They all need multiple flights for a lander mission. My opinion a moon rocket should be A moon rocket(singular). We did this 50 years ago. There is no reason why we can’t do that now. Bob Clark

The Saturn V development program resulted in spectacular successes. The Soviet N-1 program resulted in spectacular failures. Robert Clark

The fact the SH/SS keeps exploding in flight is evidence the SpaceX little 5 second burns are insufficient to qualify the Raptor for flight. Everyone acknowledges the engines on the N-1 were insufficiently tested. The elephant in the room everyone is ignoring is the same is true of the Raptor. https://twitter.com/RGregoryClark/status/1700872620604891324/photo/ Bob Clark

What I mean to say is it is standard industry practice to do full thrust, full up(all engines), full flight duration(full length of an actual mission) static firings to flight qualify a rocket stage: SpaceX using the term “full duration” for their 5 second burns misleadingly gives the impression these 5 second burns are sufficient to qualify their stages for flight. The fact their stages keep exploding in flight is clear evidence they are not. Bob Clark