sevenperforce

It would have to be in resonance with the current moon (a la the Galilean moons of Jupiter) to stay in a stable orbit. With that sort of resonance in play, it might become possible to use gravity assists to transfer back and forth between the moons and LEO without as much propellant consumption, which would open up additional trajectories and mission configurations. With a smaller second moon, we'd also have a lower barrier to landing and thus potentially an earlier initial landing. This would, however, lead to a two-part race to gain "first" status over both bodies, which might have extended the space race for longer and led to more accelerated development.

I wrote the Wikipedia sub-entry for cislunar delta-v budget, based on the 2015 NASA manuscript detailing options for staging orbits in cislunar space and other NASA resources. That latter paper is probably the most instructive here (see page 6, labeled page 232, in particular). For three-day direct-transfer trajectories between LEO and LLO, TLI will cost you 3.152 km/s and LOI will cost you 893 m/s. The LOI burn is reversed to return to Earth entry interface. The absolute lowest-LOI-cost direct transfer, at 4.5 days transit, is 813 m/s. If you want to get under 700 m/s then you need a long-duration low-energy transfer. Page 10 (labeled page 236) shows a transfer to LLO which costs only 670 m/s but takes 84 days. If you can handle a 129-day transit then you can get this as low as 651 m/s, per Table 4-4 on page 14(240). Absurdly so. No, he's talking about the amount of propellant needed for Orion to go through LLO rather than through NRHO.

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. With an added 0.6 tonnes ESM dry mass as proposed by @RCgothic, Orion needs ~6 tonnes of remaining props to develop the 900 m/s of dV needed to return from LLO to Earth entry interface post-mission. The launch mass of the entire Apollo LM (initial pre-extension configuration, Apollo 11-14) was 15.2 tonnes. Unfortunately, Orion can't brake that much weight from TLI into low lunar orbit, not even with @RCgothic's upgrade. It would need to be carrying a minimum of 13.4 tonnes of propellant plus the 6 tonnes it needs for the return. SLS Block 2 is already expected to be capable of delivering this much to TLI. If you're proposing a single-launch architecture for SLS Block 2, that's one thing; if you're proposing a different version of SLS Block 2, that's a different thing. I explained to you four months ago that the minimum mass of a Standard-Cygnus-derived crew module would be over 2.6 tonnes, not 2 tonnes -- before adding life support or astronauts -- and would be end up taking up double the maximum amount of vertical space available for co-manifested cargo. There is no uncertainty here at all. The proposed Exploration Augmentation Module (which is based on the proposed four-segment "Super" version of the Enhanced Cygnus in your post, not the much lighter Standard Cygnus) could support a crew of four for up to 60 days while berthed to Orion. It cannot do so independently, and there was no suggestion or implication by anyone that it could do so independently. Not exploding is an important requirement for a rocket, especially for one intended to carry crew. Agreed. Important requirements for a rocket include: Not exploding (AS-203, A-003) Engines not failing (AS-101, Apollo 6) Helium staying out of the combustion chamber (AS-201) Maintaining steering control during reentry (AS-201) Avoidance of re-contact between stages (A-001) Recovery parachutes remaining intact (A-001) All of those important requirements failed during the Apollo test program. Fortunately, the Apollo test program was a test program and not an operational mission program. Also fortunate, the Starship test program is a test program and not an operational mission program. As noted, I explained to you four months ago why neither the H10-3 nor the H10+ would be acceptable for this due to having the wrong engine and too much vertical height and not enough mass budget for landing legs and a low-boiloff system. You can be certain that it does not. The Cygnus itself has a mass of 3,300 kg, which could maybe be reduced to 2,630 kg if you do a complete redesign and strip away everything that makes it useful.

Or is there? Agreed. This stuff is all going to be super cooperative for the foreseeable future.

Have they downgraded the New Glenn second stage to a single BE-3U, or am I missing something?

There would be virtually no weight saving on SuperHeavy because it is already as lightweight as it can be apart from relatively lightweight things like grid fins. You can't apply the mass ratio of a lightened Starship to Superheavy because they don't correspond. Not at all. So when you plug your numbers into Silverbird and it gives you a wildly high figure, and you conjecture "This surprising result must be due [to] the greatly reduced dry mass of both stages," this should be a clue. If you put nonsense numbers in for the Superheavy In your blog post you eventually propose that your super lightweight expendable Starship be converted into a horizontally-landed upper stage. This would require a fairing, wings, and a heat shield...all of the things that were removed to get it down to ~40 tonnes. See the problem?

Let's suppose the crew cabin based on the Dragon 1 pressure vessel -- loaded -- comes in at 3.5 tonnes. Dragon 1 had a dry weight of 4.2 tonnes but part of that was aeroshell and heat shield, so shaving off 700 kg while adding in crew, consumables, ECLSS, engines, and tanks seems reasonable. Volume is limited and safety is paramount, so storables are going to be necessary. Assuming a vacuum specific impulse of 316 seconds and flying all the way back to LEO, that's going to require something on the order of 20 tonnes of props. Lunar liftoff mass of 23.5 tonnes corresponds to a necessary liftoff thrust of around 46 kN or one AJ-10. Fortunately, hypergols have great bulk density, so we'd be looking at needing only 15.4 cubic meters of tankage on the ascent vehicle. An octet of capsule-shaped tanks inside the specified OML accommodates this easily. The challenge is getting that from LEO to the lunar surface. This time, four engines certainly won't do it, and six won't fit in the same OML (assuming the same RL-10C-1-1s). So let's try and squeeze five engines under there. We'll make the LOX tank slightly smaller than the 5.2 meters to allow for some sort of landing legs to fold down from the OML, but we'll keep the LH2 tank at the full diameter: By my math, this gets a volume of 116.94 cubic meters on the lower stage, allowing for 42 tonnes of hydrolox. Assume a slightly worse mass ratio than Centaur, given that we need landing legs and more structural support -- let's say 4.5 tonnes tank and structure, plus another tonne of RL10s. So dry mass on the lower stage is 5.5 tonnes plus our 24-tonne upper stage, which gets us 3.93 km/s Δv. Nowhere near enough (and besides, this stack exceeds what Falcon Heavy can put into orbit, meaning the lower stage also needs to circularize). We are volume-limited here by the AJ-10, so we can't really add more hydrolox. If we want to get that full 5.93 km/s Δv required to reach the lunar surface, then we need to seriously shrink the ascent stage. The lunar module came in at 4.7 tonnes with 2.4 tonnes of propellant, but that's not enough for us; we need to get all the way back to LEO. We need something like 14 tonnes of propellant to get from the lunar surface to LEO, ignoring added tankage mass. On the positive side, the LM was only 2.83 meters high including the engine, allowing us to add another 4.2 meters of height to the lower stage tanks, increasing our volume by 89.6 cubic meters and bumping our propellant load up from 42 tonnes to 74 tonnes. This stack would have a total mass of 99.2 tonnes, meaning that Falcon Heavy will leave it 756 m/s short of reaching orbit. It will need to burn 15.5 tonnes of hydrolox to reach orbit, leaving it with 58.5 tonnes in reserve...about 600 m/s short of what it needs to go from LEO to the lunar surface. Of course you can imagine replacing the liquid hydrogen with liquid methane....

No one wants a sortie lander. To achieve the goals of Artemis, we need substantial downmass -- something closer to the 46 tonne (launch) mass of the Altair lander. Add a reasonably-sized crew capsule, and doing this in a single launch means a vehicle capable of throwing upwards of 70 tonnes to TLI. 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... If you go back and look at the actual sequence of replies, you'll see that I was replying to your comment where you said that a moon rocket should be a single-launch affair: "We did this 50 years ago. There is no reason why we can't do that now." It's disingenuous to start by criticizing a distributed-launch architecture on the grounds that we need a single-launch architecture, then respond to criticisms of a single-launch architecture by saying we actually need distributed launch. Zubrin's 2018 op-ed (which was very short on detail) proposed three Falcon Heavy launches to LEO at 60 tonnes each, in the form of a cargo lander that could take itself from LEO to the lunar surface. Assuming 453 seconds Isp, that's 29.2 tonnes to TLI per launch, which comes to 87.6 tonnes to TLI total, substantially more than the 70 tonnes I proposed for a single-launch architecture that you asked for. Don't get me wrong -- I love the idea of distributed launch. In fact, that was my point; I was explaining why a single-launch architecture was prohibitively difficult SPECIFICALLY because you said it was necessary. Distributed launch is definitely the way to go. That said, Zubrin's numbers are...rather aspirational. You need 5.93 km/s Δv to go in either direction between LEO and the lunar surface. Starting at 60 tonnes in LEO with engines pushing 453 s Isp, you would need to burn 44.2 tonnes of hydrolox to land on the moon with bone-dry tanks. You'll need to pull this off in a single TLI burn to avoid Oberth losses and multiple passes through the Van Allen belts -- no less than 400 seconds preferably. So you need to push at least 110.5 kg of propellant through your engines per second, giving you a minimum thrust of 491 kN. The RL10C-1-1 gives you a thrust of 106 kN, so four won't be enough. But even with four of them in an EUS-style cluster, that's 752 kg of engine alone, leaving a mass budget of only 3 tonnes to hold over 44 tonnes of hydrolox. By comparison, Centaur III's tanks weigh in at just over 2 tonnes and carry less than 21 tonnes of hydrolox. Such a vehicle wouldn't even fit on Falcon Heavy. Hydrolox has a bulk density of around 0.36 kg/L, meaning that ~45 tonnes will require 125 cubic meters of tank volume. Four RL10C-1-1s will fit inside the fairing -- barely -- but they take up 2.4 meters of vertical space (although a carefully-designed thrust plate can reduce this to around 2.13 meters). The Falcon Heavy extended fairing has a usable internal diameter of 4.57 meters, but we need room for landing legs. The landing legs on the Apollo Lunar Module added just over 2 meters to its diameter while stowed on top of the S-IVB, but let's assume implausibly that a modern design can cut this in half, requiring only 50 centimeters of stowed clearance per landing leg. Taking tank skin thickness to be negligible (not a good idea for a lander, mind you), that gives us a maximum internal tank radius of 1.785 meters, for a cross-sectional area of 10 m2. Assuming ellipsoidally-capped tanks with a rule-of-thumb half-radius height, the total stage height would need to be 15.2 meters, leaving a cosy 1.34 meters of vertical volume on top and impinging on the fairing separation system: I'm not sure how Zubrin expects anyone to construct a moon lander slightly taller than the New Shepard booster, complete with landing legs and thrust structures and egress ladders, all while somehow getting a 44% higher mass ratio than Centaur III.

I couldn't have said it better myself.

I saw the original post, started typing a reply, scrolled up a bit, and then saw you had said this, which was 95% of what I was going to say. Milankovitch cycles are particularly interesting (for me) because they leave directly observable traces on Earth's surface...some of which are even visible with the naked eye. The full Milankovitch cycle is the fusion of periodic changes in axial tilt, orbital eccentricity, perihelion, and precession rate. Of these, eccentricity has the largest forcing function on global temperature, and it operates on a roughly-100,000-year cycle created primarily by Jupiter. I believe that the amplitude of the changes are forced primarily by Saturn, though I'm not sure. The function looks like this (going back and extrapolated forward by just under a million years): Whenever eccentricity reaches a maximum, the increase in solar insolation at perihelion reaches a maximum, and so global temperature starts to rise. This increase in temperature releases carbon dioxide stored by the oceans, which further drives up global temperature in a rapid spike. As glaciers melt and ocean levels rise, the dissolved gas carrying capacity of the ocean increases, allowing the oceans to slowly scrub the atmosphere of excess CO2, allowing global temperature to gradually trickle back down until the next peak in eccentricity. These temperature cycles are recorded on Earth in a number of ways, both organic and inorganic. Foraminifera, or forams, are single-celled organisms with calcium carbonate shells, and since calcium carbonate (CaCO3) contains oxygen atoms, foram shells trapped in benthic seafloor sediment create a record of the isotopic concentration of oxygen in the atmosphere at the time. Because 18O evaporates more readily at higher temperatures, higher ocean temperatures lead to a higher 18O/16O ratio in the atmosphere. Thus, benthic forams record global temperature. Similarly, air bubbles trapped in ice cores also preserve samples of the global atmosphere. Sure enough, both benthic forams and ancient ice cores reflect this sudden temperature spike and gradual decline on the exact same period as Earth's variation in eccentricity: And it's not just these two records. Because global temperature impacts sedimentation rates, and sedimentation rates impact the density of sedimentary rock, we can literally see Milankovitch cycles recorded on the sides of cliffs in certain areas, where the rapidly-deposited sediment has weathered away faster than the slowly-deposited sediment: I'm a fan of this because it's one of the bodies of evidence that helped me to leave the science-denial cult I grew up in.

You can always Ctrl+A to select all text on the page and read it quickly without changing your forum theme. On a related note, thanks for pointing out that there is a dark theme. Much easier on the eyes. Our polymath fellow is proposing wings for Superheavy, which is a nonstarter in just so many ways.

This is very helpful if you do not have a well-instrumented engine with modern realtime telemetry and must physically inspect parts to see what went wrong. This is much less helpful when you have so much advanced instrumentation and realtime telemetry that you typically don't perform physical inspections of your parts even after a static test because you already know what went wrong before the vehicle can even be detanked. It does not. You cannot take the mass ratio of an aspirational upper stage and imagine that it will simply scale to the same mass ratio for a lower stage. Lower stages have to be stronger than upper stages. That's basic. If you're removing mass left and right for an expendable version, it's not going to be able to support itself horizontally.

To expand on this point, let's say that @Exoscientist's speculation was 100% true -- the majority of these tests are aiming for exactly 120 seconds, and some significant number of those tests fall short of 120 seconds because the engines spontaneously fail somewhere between 110 and 115 seconds (or something like that). Raptor is clearly unreliable, right? Nope, this still doesn't provide any meaningful evidence that Raptor is too unreliable for flights, let alone flight tests. For example, let us imagine that the Raptor manufacturing process has some chance (let's say 3%) of introducing a fatal defect in the turbopump exhaust injection manifold, and that defect is undetectable except through static testing. Let us further suppose that engines with the defect have a 99.999% chance of failing before 120 seconds, and that engines without the defect cannot develop the defect by static fire testing. Because SpaceX is hardware-rich, they could simply elect to static test each of their engines a few times and throw out the bad ones. This would be a completely reasonable way to eliminate the defect, if that's the way they chose to do it. Absolutely nothing can be inferred from these observations, other than the fact that SpaceX has an active, healthy, aggressive test-firing program.

I think you're correct. If I remember accurately, Rocketlab is fairly flush with engines but not so flush with stages, the opposite of ULA (which can churn out stages rather quickly but is depending on BO for the engines). Might explain why Rocketlab is less worried about saltwater incursion in its engines, because it can just replace them if they don't perform well in test fires. The Rutherford engines are probably more accepting of harsh conditions because they lack combustion-based turbopumps.

One of the primary reasons to pursue SMART is to increase launch cadence, not decrease launch costs. Blue Origin is not exactly flush with engines and their factory isn't moving very fast, so if ULA (or its successor) wants to get a reasonably high launch cadence for launching the Kuiper satellite constellation, they might need to recover the engines simply to avoid schedule delays. If they can recover and refurbish the engines faster than Blue Origin can build and ship them, then SMART makes a lot of sense for that reason alone, even if the cost savings are marginal.

It's got a photo of me and my kids, too. All Hail Sherman Vulcan Centaur 0 (the one with no SRBs) can't quite get off the ground on its own. The BE-4s are underpowered for launching it single-stick, so they have to launch it partially detanked. With two SRBs (like in this launch), they can fully load the first-stage tanks but it's still rather slow getting up. Sustainer architectures are weird.

I am pretty damn sure that all-up, full-mission-duration static fire tests are NOT more challenging for engines and stages than all-up, full-mission-duration launch tests.

Dude, this is hogwash. You have absolutely no idea what the planned burn time was for any of these tests, what was being tested, whether these were acceptance tests or tests to failure or outlier tests...nothing. You're looking for patterns that don't exist. You might as well throw in your lot with the day-trading dopes arguing about which candles predict a new stock market trend.

Sort of, but not entirely. The Saturn I used an RL10-based second stage, while the Saturn IB used the same basic upper stage with a J-2 engine, which became the S-IVB-200. It flew a few times before Apollo 4. However, the third stage that flew on the Saturn V in the Apollo 4 test was a different configuration, the S-IVB-500, with a flared interstage, a different helium pressurization system, a new auxiliary propulsion system, and a different separation system.

I'll even go one step further and point out that the development of the Saturn V actually deviated significantly from prior US launch vehicle development by doing an all-up test on the first actual launch. Prior to the Apollo program, virtually all rockets were tested one stage at a time. The first stage would be ground-tested, then test-launched with a dummy upper stage and payload. The second stage would then be ground-tested, and only after all of this would it be stacked onto the first stage with a dummy payload for an integrated flight test. For three-stage rockets, this would proceed even slower (first stage with two dummy stages, then first two stages with one dummy stage, then an all-up test with a dummy payload, then a true integrated test launch). The Apollo program deviated dramatically from this approach by doing an integrated flight test of all three stages AND a functional CSM on Apollo 4. They focused on validating all of the systems independently (and in parallel) so that they would be able to put everything together on the first go. And of course the Apollo program was wildly successful. So if there is a "lesson learned" from Apollo, perhaps it is the lesson that deviating from past practices can be a really good idea if you have a consistent vision and the resources to make it work.

Also, keep in mind: a lot of start ups gravitate toward kerosene for their first stage (or even solids). Because of the high density of kerosene, you tend to have a much thinner rocket. Not only are they going for full reusability, but they want a methane first stage, so it makes sense to use a wider stage due to methane’s low density, which in turn gives them more volume to work with for their hydrogen upper stage.

No problem. I would suggest making a spreadsheet, carefully laying out the dry mass and propellant capacity of each stage, and so forth. How you compensate for lunar gravity is up to you; just keep it consistent. The TWR is the total thrust of all firing engines divided by the weight of the vehicle; the weight of the vehicle is its mass times the gravitational acceleration. Personally, I just imagine it's Earth gravity for everything and then if I am dealing with lunar touchdown or takeoff I apply the appropriate transformation afterward. Yes, that's correct. The lunar module ascent had two phases: a ten-second burn straight up to clear terrain, and then a pitchover and burn to orbit. The precise moment and orientation of liftoff was chosen to minimize phasing and plane changes. If the ascent propulsion engine had failed during the last thirty seconds of the orbital insertion, the four little aft-facing RCS thrusters had sufficient umph to complete orbital insertion. That was one of the reasons for the valve that would allow the main tank propellants to flow directly into the RCS thrusters. Yes; following successful orbital insertion, the rendezvous maneuvers were performed entirely by the RCS system. Although the RCS system had three-plane translational capability, all of these maneuvers were being done pretty much manually, often using slide rules to compute the proper heading, time, and duration of each burn. As a result, the aft-facing RCS thrusters were the only ones used for these maneuvers; the other thrusters only provided attitude control. For Apollo 11, the initial launch reached an 87.6x17.6 km elliptical orbit at main propulsion system burnout. RCS was used about an hour later, at apoapsis, to circularize. They had planned for a plane change burn as well but didn't need it. By this point, they were nearly on a collision course with the CSM so they did a Constant Delta Height burn (still with RCS) which ensured that the two vehicles would be constantly 28 km apart and they would have time to plan the remaining rendezvous burns. As an example -- the RCS circulation burn for Apollo 11 required 15.7 m/s of dV or around 13.7 kg of propellant. The burn took just over two minutes, starting at MET 125:19:34.70 and ending at around MET 125:21:36. Had the burn been performed with the actual ascent propulsion engine, it would have taken less than three seconds. I don't know the specific procedures but it definitely all would have happened pretty quickly, so it was likely automated. The RCS propellant tanks used teflon bladders to hold the propellant inside a pressurized tank. As helium pressurant was vented into the tank, the teflon bladders were compressed and pushed the RCS propellant out, obviating the need for any separate RCS ullage burn. If all of the RCS tanks had COMPLETELY failed and the main tank was being used exclusively, then ullage burns would become a problem. There was a chance that the surface tension of the propellant in the feed lines would be enough for the initial ullage puff but it would have been tricky. But that was an unlikely contingency (lots of other stuff would have to fail which would probably be LOCV anyway). What was tricky about it? It used multiple helium tanks with burst discs (for simplicity) so it could only be started twice. No, definitely not. The CSM never did anything until separation from the third stage. The S-IVB third stage of the Saturn V was equipped with a pair of Auxiliary Propulsion System modules using hypergolic propellants. Each module carried around 120 kg of propellant and boasted a trio of RCS engines and a single ullage engine. They were used to provide ullage for third-stage restarts, roll control (and backup pitch/yaw control) during the third stage burns, and general attitude control during the transposition and docking maneuver.

Why use "flight tests" here? Why not just say "as long as Raptors fail, there will be questions about its reliability" period? How is a flight test magically different from a static test? No one wants a sortie lander. To achieve the goals of Artemis, we need substantial downmass -- something closer to the 46 tonne (launch) mass of the Altair lander. Add a reasonably-sized crew capsule, and doing this in a single launch means a vehicle capable of throwing upwards of 70 tonnes to TLI. We also know you dislike solids generally, so let's look at an all-liquid architecture. Three stages are going to be required, obviously. Rule of thumb splits delta-v among stages, and you need 9.4 km/s to reach orbit and 3.2 km/s to reach TLI, so that's a total of 12.6 km/s, or 4.2 km/s on each stage. Notionally, let's imagine a kerolox first stage, a methalox second stage, and a hydrolox third stage. The EUS packs 126 tonnes of hydrolox, so let's start there. With a 75-tonne payload, it develops just 3.5 km/s. If we swap out its four RL10C-3s for a pair of J-2Xs, it will get even less. But the EUS is the largest hydrolox upper stage in development, so let's stick with that. Now our first two stages need to somehow deliver 4.5 km/s each. Now for our notional methalox second stage. We know you dislike Raptor, so let's imagine a BE-4U with improbably comparable specific impulse to RVac. Let's imagine a 45-tonne stripped-down Starship as our main tank, pushed by a quincux of BE-4Us. The BE-4 produces 2.4 MN at an estimated sea level specific impulse of 315 seconds, so if we bump that up to 380 seconds then five of them will produce a whopping 14.5 MN together. With the required nozzle extension they're going to weigh in around 4 tonnes each. The third stage and payload together come in at 232 tonnes so stage dry mass is 297 tonnes, which gets us a delightful 5.5 km/s from the 1000 tonnes of methalox onboard (reduced from 1200 tonnes because BE-4 can't take densified methalox). We've got a T/W ratio of 1.1:1, which is low but acceptable. ETwo stages deep, and we're ahead of schedule! Our first stage will only need to deliver 3.6 km/s (or roughly 4.1 km/s if you want to treat sea level specific impulse as pressure drag and just use vacuum specific impulse). But here we have a problem. The most powerful kerolox engine in the world, the RD-171, gets 337.2 seconds of vacuum specific impulse. You need a wet:dry mass ratio of 3.5:1 to achieve 4.1 km/s, putting us at a liftoff mass of at least 4,500 tonnes, almost double the launch mass of the Saturn V. You'd need at least nine RD-171s or at least 17 RD-180s to get off the ground. Or you can use a cluster of 77 Merlin 1Ds. Good luck with that. How were these "lessons" from Apollo? The Saturn V used a flame trench and powered stage separation, but these weren't lessons. NASA didn't attempt to go without powered stage separation and then fail and correct it; they just chose to use solid separation motors from the start. You seem to be confusing "lessons learned" (e.g., "we tried it one way and realized it didn't work and so we found a better way") with standard operating procedure ("this is the way we decided to do it and it worked"). Besides, Superheavy has always had a flame "trench". It's actually six very big trenches that go in every direction. The "T-Bird" S-IC-T was a Saturn V first stage designed specifically for static fire tests; it never flew and never could have flown. It completed its test firings prior to the construction of the actual flight articles. AFAIK, neither the three S-IC stages launched in the Apollo 4-6 flight tests nor subsequent S-IC stages used for actual crewed missions ever received full-thrust, full-up, full-mission-duration static burns. It continues to puzzle me why you insist that building a new facility for full-duration static test fires of Superheavy would somehow be categorically and qualitatively better than conducting full-duration test fires simultaneous with the test launches themselves. We have no evidence that SpaceX fails to share thrust level and telemetry data with its government partners. Agreed. As outlined above, building an actual single-launch architecture for meaningful moon landings would require a rocket almost double the liftoff mass of the Saturn V (and likely more).

To answer these questions reliably, your best approach is to look at the actual thrust of each engine, the number of engines on each stage, and the weight of the stack at each staging event. Then just do the math. All of that is going to be much more accurate than guessing at whether publicly-posted numbers were posted by people who did the math correctly. The wet and dry masses of each stage is all public from NASA documents. I happen to know a good bit about this so you're in luck! The same fuel type was used, yes, but it was not sourced from the same tanks. The RCS thrusters on the ascent stage provided 100% of the reaction control for both the descent and ascent; there were no RCS thrusters on the descent stage. The tanks on the descent and ascent stages were completely separate without any interconnections. The descent propulsion engine was a throttleable, restartable, gimballed hypergolic engine fed exclusively from tanks housed within the descent stage. Because this engine could be gimballed, the RCS on the ascent stage was used only for roll control while the descent stage was firing and ullage while the descent stage was starting up. While I know that the RCS thruster controls were designed to permit translational burns for docking, I am not sure whether translational firing was active during hover and landing. I should also note that although the lunar module was capable of acting as the active translational actor during docking, it never did so in practice; it would just hold orientation and allow the CSM to come to it. The ascent stage had three sets of propellant tanks: one main set of propellant tanks which fed the ascent propulsion system and a pair of redundant propellant tank sets for the RCS system. Each of the redundant RCS propellant tank systems contained half the propellant needed for descent RCS plus all of the propellant needed for ascent RCS, so that if one of the tanks experienced a problem they could still perform all necessary ascent burns. There was an additional (closed) valve linking the main propulsion propellant tanks to the RCS system so that the RCS system could be powered directly from the main prop tanks if both sets of RCS system tanks failed, although this would reduce the amount of propellant available for the ascent propulsion system. The ascent propulsion system was constant-thrust and fixed, so the RCS system had to provide pitch and yaw as well as roll during ascent. Although the ascent propulsion system was technically restartable, it was not ordinarily restarted during ascents because the helium pressurization for restarts was a little tricky: it was a single burn from the lunar surface to lunar orbit. It was restarted for disposal burns, and required an RCS ullage burn for those restarts. There was no ullage burn off the lunar surface because lunar gravity provided sufficient propellant settling. @StrandedonEarth is correct: the SIV-B did a single burn from staging to parking orbit, circled Earth a few times, then restarted for the trans-lunar injection burn. The parking orbit was both to simplify phasing and to allow time for systems check-out before committing to the rest of the mission. Oberth, drag, and boiloff losses were accounted for but minimal.

The comparison between the RL10A-4 and the J-2X is illustrative. The RL10A-4 has an expansion ratio 84 to 1 and gets an Isp of 451 s. The J-2X has an expansion ratio of 80 to 1 and gets an Isp of 448 s: http://www.astronautix.com/j/j-2x.html . So the J-2X Isp is over 99% of the RL10A-4 value. I'm not sure where astronautix got the 80:1 figure, but the actual manufacturer said the expansion ratio was 92:1. So I'm inclined to believe them. Which goes to my point. You can't squeeze expander cycle performance out of a gas generator engine. It's a different beast.