sevenperforce

OMG YES

OOOOH the camera itself is on a flap. Re-entry heating!! HOLY excrements that is beautiful.

Flaps are moving. Are we stable?? I'm seeing a persistent roll. Possibly not good.

OMG, what gorgeous views.

The onboard computer has decided not to perform the in-flight restart burn test. Re-entry in 8 minutes!

Operating the flaperons in the hypersonic regime is completely untested. There was some early talk of putting a subscale demonstrator on Falcon 9 just to make sure the flaps operate properly...who knows how that will go. Looks like we have a good SES-2? Wait, no, that was just a cold-gas flip.

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.