sevenperforce

Relight coming up soon! Expected re-entry in about 10 minutes.

Prop transfer demo started.

We got a PEZ door shot!

PEZ door appears to be open? Doesn't look like we have views of it though.

The vehicle was yawing and rolling really hard well before landing engine burn startup. I'm thinking that the grid fins didn't have enough authority to keep the vehicle properly oriented, which meant there was too much propellant slosh to get the engines consistently lit.

Over 7 km/s. Starship engine cutoff!!

Boostback complete; Raptor shutdown. The split screen view is amazing. As expected, no entry burn for the booster. Should hit the atmosphere in a few seconds here. We've got grid fin movement! Coming in super hot. Seems to be controlled so far. Looks like the grid fins don't have enough control authority. I'm guessing Superheavy broke up from aero. Back to the ship cam. Only a few missing tiles from launch.

Loving the booster and ship cam shots! Gorgeous stage separation shot!!!

Off the pad!!

Hold or no? PAST TEH HOLD! Light this candle! Go SpaceX!

Bumped launch to 8:10 CST to give boats more time to clear.

Changing a first stage from a standard bell nozzle to an altitude-compensating nozzle does not increase the performance of the first stage. Quite the contrary. Switching to a plug nozzle (or other altitude-compensating nozzle) DECREASES performance at liftoff, which results in lower actual specific impulse, lower takeoff thrust, and greater gravity drag. The slightly-increased specific impulse at altitude is not enough to make up for the loss in thrust at sea level. The only reason to use altitude-compensating nozzles on a first stage is if you have a sustainer architecture that goes all the way to orbit, where (a) you aren't worried about liftoff thrust because you have separate boosters, and (b) you have enough time for the increased performance at altitude to make up for your laggy performance at sea level. It's entirely doubted. This is a commercial environment, not an academic one. More importantly, ideas don't launch rockets. Engineering launches rockets. The engineering is the hard part.

This is mostly speculative, but it's possible that they will be putting it into the original trajectory from IFT-2, then executing a radial-in burn to ditch in the Indian Ocean instead. Alternatively they may do an acceleration burn just before re-entry. The second stage failure in IFT-2 was the result of the LOX dump that took place at the end of the burn. Because IFT-2 wasn't a true orbit, they didn't need quite as much propellant in the second stage as they would have otherwise needed. HOWEVER, they still wanted it to be as close as possible to an ordinary flight for T/W ratio reasons, so while they carried only as much CH4 as they needed, they carried significantly more LOX. Slosh of this LOX would have interfered with center-of-mass balance during re-entry, though, so they dumped the LOX out the back end just as they were approaching orbital velocity. This somehow caused an explosion (possibly due to some sort of leak of fuel-rich turbopump exhaust or fuel-rich pressurant). Presumably they didn't want to dump the LOX after reaching orbital velocity because they didn't want to risk the dump causing a propulsive effect that shifted the impact zone. For this test, it seems like the plan is to carry a full propellant load (both LOX and CH4) but never dump propellant. Instead they will cut off early, perform the tests (prop transfer and payload door operation) during ballistic coast, then relight the engines to both (a) burn off the remaining propellant and (b) alter their trajectory. This could be a radial burn at apogee that alters their impact point from Hawaii to the Indian Ocean, or it could be a prograde burn that increases their velocity just before re-entry in order to better simulate re-entry speeds.

Love this! Current progress:

Apropos of nothing, I have recently gotten a 3D printer and have gotten pretty good at messing around with it. I'm going to be printing this and assembling it with my 10-year-old for his birthday. I'll post photos once it is done!

The article describes the extent of the current "life support systems": it has everything necessary to "transport pressurized cargo and experiments" to ISS. In other words, it has the ability to maintain pressure and temperature. This makes it suitable for supporting crew while berthed to Orion, because it can independently maintain pressure and temperature. It does not have what we would consider to be a life support system for crew.

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.