sevenperforce

It's a kerolox rocket that uses an aft-first entry heat shield and parafoil recovery: More power to them if they can make it work! Kerolox doesn't offer much margin here. It looks like they are contemplating a cluster of 7-9 engines inside the toroidal heat shield, possibly relying on stagnation inside the heat shield to protect the engines. It's claiming 25 kN of thrust which, for an assumed 1.3 T/W ratio, caps the liftoff mass at around 2 tonnes. This looks like a pressure-fed engine so I don't see how they get a vacuum specific impulse greater than 325 seconds or so. To achieve over 9300 m/s of Δv, it will need a dry mass no greater than 106 kg, of which 13 kg is the payload. That's pretty tight.

you had one job

Well, "loss of efficiency" is a bit of a misnomer. Some estimates for the maximum speed of a hydrogen-based scramjet go all the way up to Mach 24, but the theoretical maximum speed is probably closer to Mach 12-16, with practical considerations kicking in even lower than that. So it's not "loss of efficiency" so much as it's a total shut-off in thrust well below orbital velocity. Airbreathing engines are good for cruise, but not for acceleration. Mach 5.5 is less than a quarter of orbital velocity, so you've burned a LOT of props and still haven't gotten very far. As noted, the effective specific impulse between Mach 5 and Mach 6, for an airbreather, is often less than hydrolox because of the low L/D ratio. I thought we were comparing to Falcon 9, not Starship. 10 tonnes to LEO is still pretty solidly toward the upper end of medium-heavy lift vehicles. A dual-engine Common Centaur upper stage with a 10-tonne payload develops 4.3 km/s of Δv, well short of the 6.3 km/s you said is needed to bridge the gap between a Mach 5 hypersonic ramjet and orbital velocity. Bump up to a Centaur V and you get 6.7 km/s, which is enough...but you're now looking at dimensions of 5.4 meters wide and 33.3 meters long, plus your payload. That's at least twice the length and diameter of the bomb bay on the B-52 Stratofortress. A hypersonic ramjet aircraft eight times the size of the B-52 is a pretty huge airbreather.. Let's go back to the XB-70 Valkyrie, and let's suppose that upgrading the engines and changing the fuel type can get us up to Mach 5.5 with no OML changes whatsoever. After all, it was designed to cruise for half an hour at Mach 3 but it doesn't have to cruise at all in this instantiation. Let's also suppose that the range-reducing changes can more than double both the width and length of the bomb bay, from 3x9 meters to 4x20 meters, all still without changing the OML. Finally, let's suppose that the release of a highly sensitive hydrolox rocket stage in the atmosphere at hypersonic airspeeds is a simple, easily solved problem. A pure hydrolox solution just isn't going to work here. What we can do, however, is use a two-stage solution. Put a Centaur III on the front and a solid kick stage on the back. The Castor 30B might work. Coming in at 3.5 meters long, 2.3 meters wide, and 14 tonnes, it has 270 seconds of specific impulse, which should allow it to nudge our Centaur-III-based terminal stage from Mach 5.5's 1.9 km/s up to around 2.8 km/s. That should be enough for Centaur III to take over and get up to 7.8 km/s (neglecting gravity drag and air resistance) if we limit the payload mass to about 7 tonnes or so. In today's dollars, each XB-70 cost over $8 billion to produce (not including operational costs), and achieved (on average) just five high-supersonic flights before being retired. That's a pretty hefty pricetag for theoretically getting less than 8 tonnes to LEO. Hermeus looks like a fantastic company with some cool ideas. I have sketches of a combined-cycle turboramjet dating back years; being able to build a working one around an existing jet engine is really impressive. But these speeds still just aren't useful for orbital applications. The Halcyon's planned quad-engine, 125-passenger hypersonic business jet puts it in a payload class smaller than the P-8 Poseidon bomber, which can only carry about 10 tonnes of bombs...not really enough to get anything at all to orbit, even starting at Mach 5.5.

Well, no, an airbreathing engine can't make the entire flight to orbit at all. The Isp-vs-speed math does not particularly care whether you are an SSTO or a first stage. Presumably we would need to demonstrate the reuse of a hypersonic airbreather one time before supposing that a hypersonic airbreather could be reused thousands of times. The only turboramjets ever flown achieved an average of 234 hours of high-supersonic flight time per airframe. Let's compare to the Falcon 9 and use the typical 16 tonne Starlink type delivery. Based on your numbers the upper stage needs to deliver 6.3 km/s. This is too beefy of a job for a closed expander cycle, so let's go with a bigger engine running on staged combustion, like the currently-in-dev Chinese YF-90 boasting 2.2 MN of vacuum thrust and 453 seconds of specific impulse. The rocket equation helpfully tells us that we need a propellant fraction of 76% to make that work. The engine comes in at 4.8 tonnes so we're looking at an upper stage with a mass of at least 87 tonnes, triple the size of the Delta IV 5-meter Cryogenic Second Stage. We're looking at a length of probably 30 meters or more. The XB-70 Valkyrie, the largest flight-tested supersonic bomber of all time, had a huge bomb bay...at just under 30 feet. So your hypersonic carrier aircraft would need to be something like triple the size of the Valkyrie.

That will happen at the higher scramjet speeds, and even then only for hydrocarbons. Scramjet theoretically could operate beneficially up to say Mach 15. That is to say, if they could be made to operate reliably. [snip] Ramjets have been seen to operate at Mach 3.5 to 4. Theoretically they should be able to reach Mach 5.5. That is why the test by Hermeus next year is so important. If they can reach Mach 5.5 even if not used as a first stage for an orbital rocket then it can be used as a hypersonic transport. Instead of 6 hour flights cross-Atlantic or cross-continental USA they could be done in 1 hour. The propulsion performance curve you've provided, while accurate as to the engine cycle itself, fails to account for the impact of drag vs acceleration on the vehicle in relation to gravity. I do suggest you take a look at that link I provided above. (The original page is dead but I've given the path to Wayback.) The logarithmic curve takes the shape it does because of the nature of airbreathing engines. Imagine first that you have a ducted fan propeller-based engine driven by an electric motor. At a standstill, your engine is essentially acting like a fan, simply blowing air out the back at some particular velocity, which produces thrust via momentum transfer. Let's say for simplicity that the fan velocity is 200 m/s. Your maximum thrust is achieved at a standstill, because once you start moving, the amount of additional velocity you can impart to the air starts to drop. At 50 m/s, you can only impart 150 m/s so your max thrust has dropped to 75%. At 100 m/s, you can only impart 100 m/s so your max thrust has dropped to 50%. At 200 m/s, your engine can be operating at max thrust and yet you won't be accelerating at all, even if you disregard friction entirely. (Obviously in real life propellers have differentially curved blades to allow them to continue operating at higher speeds with reduced efficiency, but this is a first-order approximation anyway.) This point ALSO applies to jet engines. You're sucking in an airstream with some inlet velocity, mixing it with fuel, igniting it, and then allowing it to expand out the back. The thrust is the mass flow times the DIFFERENCE between the inlet velocity and the exhaust velocity. Although the amount of energy you're adding to the airstream remains essentially constant, the net thrust drops as you move faster because that difference shrinks as the inlet velocity gets closer and closer to the exhaust velocity. But that's only half of the problem. Airbreathers have extremely poor thrust-to-weight ratios compared to rocket engines, so you'll need to use aerodynamic lift to continue climbing. Lift-to-drag drops as Mach number increases: a Boeing 747 has a L/D ratio of around 15:1 at its cruising speed of Mach 0.85 while the Concorde and SR-71 had L/D ratios of just over 7 at Mach 2+. The X-15's L/D ratio was 4:1. The very best designs for a hypersonic aircraft might be able to achieve 5:1 at high Mach numbers. But if your T/W ratio is 0.3 gees (which is still really impressive for a hypersonic engine), then two-thirds of your thrust is being converted into lift-induced drag just to resist gravity. If two thirds of your thrust is being wasted, then that means your effective specific impulse is just a third of what you expected. If your expected ramjet specific impulse is 1300 seconds at Mach 5.2, you'll be rather disappointed when you get an effective specific impulse of just 433 seconds. At that point you're better off just using a pure hydrolox engine for your spaceplane. Where are you getting the idea that a hypersonic ramjet engine would be able to boast thousands of reuses with one-hour turnarounds?

On the other hand, a Raptor still did explode on both booster landing attempts. Engines that run on liquid do tend to explode when you feed them gas.

Not when your vehicle has to be ten times larger because the engines have a TWR that is a dozen (or more) times smaller. By way of elaboration: The engines have a TWR that will start fairly low, increase modestly with velocity as ram effect compression initiates, then drop to extremely low -- along with specific impulse -- at hypersonic velocities. This is all due to the airbreather's burden. At hypersonic velocities, an airbreathing spaceplane is fighting like crazy to even maintain speed, because it has to somehow gulp up air, accelerate that air to its own speed, then burn and push that air out the back end faster than it entered, while a rocket continues to merrily accelerate without a care in the world. The effective specific impulse of an accelerating spaceplane drops so low at hypersonic velocities that it will generally require several TIMES more fuel than a pure rocket, despite avoiding the need to carry oxidizer.

Not when your vehicle has to be ten times larger because the engines have a TWR that is a dozen (or more) times smaller.

Who's to say that hypersonic jet transport is any more feasible -- as a business case -- than hypersonic rocket transport?

Yep, precisely. In a funny sort of twist, part of this whole thing is to make the engines as Kerbal as possible. An engine that just works, no other systems needed, a set dry mass and nothing more, etc etc. -- repressurizes the tank for every bit of propellant it uses, and so forth. I mean, at that point, why not just have a "small helium tank" instead of a smaller LOX tank? Also, how do you propose they "trap" the water and CO2 in the small tank? I'm mildly curious to learn why they didn't use the existing heat exchangers in the engine chamber and bell -- the ones used to preheat the propellant while regeneratively cooling the structure -- and tap off that for autogen repress. I suppose it would tend to decrease overall thrust, while running more volume through the preburners doesn't have that problem. While I cannot claim to be an engineer, I do have a degree in physics along with my graduate degree in law, and I concur. Musk may make many stupid decisions, and some of them might even be engineering-related, but that doesn't mean SpaceX has problems surrounding a lack of proper engineers. We'll see. They spent a lot of time on deving the hot gas-gas thrusters for Superheavy and then just scrapped them completely when they determined that venting ullage did the same trick with lower dry mass. This is all a pretty fine balance. Samesies

Two boosters back to Terra Firma and one in the drink; GOES-U continuing to orbit. Go Falcon Heavy! Unrelated: is it just me or is the forum site doing REALLY bad?

The only hypersonic aircraft that has ever been reused, the X-15, used rocket-based propulsion and required a drop from a carrier aircraft. No attempt has ever been made to reuse scramjets. The fastest actually-reusable aircraft capable of taking off under its own power, the SR-71, had a maximum speed of Mach 3.5 (1.2 km/s) and flew an average of 540 times each across the 32-craft fleet, with at least a week of maintenance required between each flight. It is much easier to build a reusable rocket stage than it is to build a reusable hypersonic airbreathing vehicle. Hypersonic airbreathing makes full-flow combustion look simple. Also notable that the payload fraction for a hypersonic aircraft is extremely low. The SR-71 had a max takeoff weight of 78 tonnes and carried just 1.6 tonnes of actual mission payload (let's round up to 2 tonnes to account for the pilots and cockpit weight), a payload fraction of 2.6%. Compare this to Falcon 9 -- the first stage has a mass of 443 tonnes compared to the second stage mass plus payload of 131 tonnes, a payload fraction of 22.8%. To loft a second stage and payload comparable to the Falcon 9, your notional hypersonic launch vehicle would need to have a rolling takeoff mass of over 5,000 tonnes, nearly nine times the size of the largest plane to ever leave the ground, the Antonov An-225 .

If a particular use of language communicates the intended point to 99.9% of the expected audience, then that use of language is working just fine.

So it doesn't have anything to do with engine reliability after all, then? All engineering ideas are bad ideas until they are standard practice. Your feelings about the term "Chief Engineer" are about as strange as your feelings about the term "full duration".

From what I'm hearing, the description is of fully-integrated coolant flow paths that eliminates the need for secondary leak containment.

That's...really incredible. Absolutely the kind of "make it better simpler" engineering that SpaceX is becoming known for (and Tesla seems to keep screwing up). With gun magazines it doesn't matter either. The springs involved are strong enough that gravity is not meaningful. Precisely why Tim annoys me a little.

This is the hottest summer of my life. It is the coolest summer of the rest of my life.

Indeed. A "partial duration" test would be a weird and confusing way of describing a test fire. Language is descriptive, not prescriptive. But that doesn't mean anything goes. We use language because it is useful for communicating; if a particular use of language fails to communicate, then that use should be deprecated. Reminded of this.... If a usage of language is sufficient to communicate the desired information within readily available context, then that is a good use of language. "The rocket engine completed a full-duration test" could mean that the engine completed a test fire equal to the duration of a launch, but it could also mean that the engine completed the full planned duration of the test. That's where you have to look to context. If this claim is accompanied by a video of a 15-second static fire, then context has provided you with all of the necessary supporting information. I had a dream last night that I had the chance of getting a seat on a New Shepard launch with Bezos on board, but one of my wife's high school friends ended up taking my place. Bezos had pulled a Musk and decided to launch New Shepard out of the top of a barn for "rule of cool" rather than from the existing launch mount, and the engine clipped the struts of the barn, causing the launch escape system to trigger while still in the building. I simultaneously mourned the loss of my wife's friend and rapidly shorted Amazon stock, making tons of money in the process. This anecdote has nothing to do with the terms used by SpaceX or Stoke, of course. I suppose that we could say there is a difference between a "full duration test fire" and a "full duration test burn". A "test fire" can be any length, while a "test burn" suggests the length of an ascent burn. "Full duration test" is therefore ambiguous and so reference to a full duration test should be accompanied by additional context clues. For modern engines that use regenerative cooling, the thermal conditions are really quite extreme. Within a few seconds of startup, the system will reach steady-state operation: all of the heat going into the engine and all of the heat coming out of the engine is in equilibrium. As a rule, rocket engine failures don't happen slowly or gradually; if something is going to go then it goes fast. So achieving complete startup and getting 10 seconds or so of a burn is pretty good evidence that the engine won't eat itself.

For the nth time -- if the duration of the test was intended to be two seconds, then this engine completed the full duration of the test fire. If I attempt to complete a certain amount of work and I complete 100% of the work I attempted to complete, then I have completed that work even if another part of the underlying task remains outstanding for the next day.

I was on the forums for ages before I even bought the game.

I completely forgot about that! I love this forum and everybody here. Honestly it was a meaningful outlet for me during some rough times in my life. My time messing around on the KSP forums has seen me father several children, go through a nasty divorce, go through law school, get custody of my kids, become a licensed attorney, and get married to the love of my life. It's been quite a ride and the group of people here is really quite decent. If anything ever happens to these forums, I am u/lawblawg on Reddit and @justlydeserved on TikTok.

What evidence do you have to suggest that any engines exploded on relight during either Superheavy booster landing burn?

No, there was not and has never been any plan for a re-entry burn for the ship. SpaceX uses supersonic/hypersonic retropropulsion for the Falcon 9 booster during its re-entry, but there has never been any possibility of some sort of ultrasonic retropropulsion for entries from orbital velocity. The ship is not in the correct orientation for doing that and it would not make any sense. "which suggests" is doing a lot of heavy lifting there. It's a non sequitur. I might as well say "Apollo 12 getting struck by lightning is not normal, which suggests it should have deployed parachutes during ascent." There is not. It is not. There were no other ship landing burns on this flight or on the previous flights. Even if RCS and the flap control surfaces were sufficient to reorient the ship, this would not explain (a) why we saw a landing burn, (b) why we had landing burn callouts, or (c) how the ship came to a hover over the water. What's more likely -- that the laws of physics were suspended to allow a plummeting object to suddenly freeze in place in midair, or that the graphic didn't update properly? How is SpaceX "continuing to promote" this idea? Simply by having a contract? If having a contract to deliver crew means that you are promoting the idea that any engines you would use to carry that crew are already at full reliability, then that is a problem for a number of NASA contractors who have contracts but don't even have engines yet. Agreed that current reliability is where we would expect it at this stage. Additionally, as long as we don't have catastrophic or sympathetic failure modes, engine-out capability should be treated a little differently.

Spending a long time at higher altitude is a good thing for Starship. There are (basically) two different types of heat shields: ablative heat shields and radiative heat shields. Ablative heat shields function by allowing their surface to boil away to shed heat, while radiative heat shields employ a low-conductivity, high-emissivity material which can handle high heat by simply absorbing it and radiating it away into the environment. Ablative heat shields have the advantage of being able to handle higher peak temperatures because their surface can simply boil away faster to adjust, while radiative heat shields are able to handle greater total heating because they don’t get used up. If you have an ablative heat shield, peak heating is really not an issue: you can get as hot as you like and it’s not a problem because your heat shield will just boil off more quickly. Total heating, however, is a bigger issue: you need to get through the peak as quickly as possible or your entire well heat shield will boil away before you finish. As a result, you want a steep trajectory to slow down quickly. With a radiative heat shield, you need to limit peak heating so you need a more gradual re-entry with as much lift as possible. The more body lift you can achieve, the longer it will take you to slow down and the lower the peak heating will be. Starship, like the Shuttle, uses a radiative heat shield, so it wants to enter with an attitude that prioritizes body lift and keeps converting speed into altitude for as long as possible to reduce peak heating. The Shuttle had an additional issue, though: it needed to be able to achieve subsonic glide for landings, so it had WAY too much body lift in re-entry and had to execute complex S-curves to avoid (essentially) skipping off the atmosphere and losing aerodynamic control. Starship has a much more flexible control envelope since it doesn’t have to carry its entire launch structure to orbit and doesn’t have to worry about subsonic glide, so it can glide for as long and high as it wants.

[snip] I found it really interesting that the burn-through happened LONG after expected peak heating/temp. This suggests that aerodynamic forces, rather than heat alone, were a dominant force in the burn-through. The strakes are a temperature hot point. I wonder if the heat transfer past the tiles into the skin was enough to damage the skin structurally, which led to later weakness and peeling away once it got down in the denser, less-searing-but-still-very-spicy atmosphere. It could also be a glue issue. Those tiles around the strakes are glued on, so if the peak heating weakened the glue on a material integrity level, the glue still could have held until it got down into the thicker atmosphere, at which point aero forces ripped off the tiles and allowed the hot air (not hot enough to threaten the tiles but hot enough to melt steel) to enter.