You are not logged in.
Sorry, I got distracted doing unexpected chores outside until the sun went down, then I had to get the chicken and the cat "in" for the night. By that time, the meeting start was an hour and a half earlier.
GW
It would appear that the AI could not distinguish between the Crew-10 launch of a Falcon-9 and the Flight-10 launch of a Starship/Superheavy. Why? Both had a "10" in the names.
GW
There are no scramjets that will ever use a pitot/normal shock inlet. There are probably no ramjets that will use one, if the flight speed is to exceed Mach 1.5. Gas turbines, usually afterburning, will use a pitot/normal shock inlet if max flight speed is under about Mach 1.5, at the very most 2. These kinds of designs must use very volatile gasoline as fuel, although the octane number is irrelevant.
The supersonic inlets only work above Mach 1.8 at the very minimum, and perform like crap below their shock on lip speeds, if at all. Things that fly in the Mach 2 to Mach 3.5-or-more range can use shock-on-lip speed inlets of about Mach 2. These can use air cooling below about Mach 3 to 3.5 in the stratosphere, slower lower down. Above that, everything must be protected by ablatives, and you can only use dump stabilization for your flameholding. With a Mach 2 shock-on-lip inlet, your max speed potential is still only about Mach 3.5 to maybe at most 4. These will require a wide-cut fuel like JP-4 or Jet-B. More volatile than a straight kerosene.
If you are using a Mach 2.5 shock-on-lip design, your min takeover Mach is very close to Mach 2.5, and your top speed can be anything from about Mach 4 to just a bit above Mach 6, if the exit nozzle is a big enough percentage of the total frontal blockage area. You can use a straight kerosene fuel, like JP-5 or JP-8, or Jet-A or Jet-A1. There are also kerosene-like synthetics available, like RJ-5/Shelldyne-H, which resembles kerosene, but is very slightly denser than water. It does freeze too easily, though!
All of these supersonic inlets feature external compression features (spikes or ramps), and they have some min inlet throat area, either at or close downstream of, the inlet cowl lip capture feature. Flow at the min area and slightly downstream is sill supersonic, but just barely. That is followed by a terminal normal shock-down to subsonic flow, followed by divergent area increase to accomplish the diffusion to the needed degree of well-subsonic compressible flow in the duct.
Scramjet can use the same external compression spike or ramp features, and a similar cowl lip stream-tube capture feature, but there is only some modest internal contraction in duct area to something close to Mach 2.8 supersonic speed. There must be 5-10 duct diameters length of this supersonic "isolator duct", to maintain stable combustion in the scramjet combustor, however that fuel might be injected, and however that flame stabilization might be obtained. Further, there can be no bends at all in this straight "isolator duct". Doing so immediately causes shock-down to subsonic flow.
GW
I'd like to know exactly how t figures its estimates, and from what inputs. It might well be better than what I do, which is just a approximation.
GW
Using LEO speed as a basis, did you notice the 20% gravity loss for the first stage because is launch T/W was under 1.2? I was startled to see only a 0.5% drag loss, but then, with the low T/W ratios, it really wasn't moving very fast when it left the sensible atmosphere at 1st stage stage-off. Somehow, that calculator is using dimensions and T/W ratios to estimate those losses. It didn't say how those were estimated.
GW
It would appear that the next flight is a reprise of the last flight, to (1) get more data from the booster regarding the fast flip at staging, and (2) see if they can complete the Starship upper stage mission at least through entry, and hopefully to a water landing in the Indian Ocean.
Bear in mind that what I said about a 10 Hz pressure wave signal I had been hearing in Raptor tests is not a precise frequency. That makes the resonance with any certain piping lengths less sharply precise. And an increase in thrust levels would increase the amplitude of the thrust oscillation, if they have not corrected it.
Meanwhile, the testing has recently been a lot quieter. I would hazard the guess that the Raptor-3 tests are all being done in the surface stands with the sound-dampening water deluges. That higher thrust level is apparently just too loud to be on the surface horizonal stands or that tower stand Andy Beal left for them. Raptors have a lot more thrust than Merlins. So they are very much louder.
GW
Interesting info in post 2140 about the causes of Starship/Superheavy failures. The transfer pipe failure in the Superheavy sounds rather suspiciously like a mechanical resonance failure from the 10 Hz thrust oscillation signal I keep hearing in tests. That's just speculation on my part. But, I do note I hear less thrust oscillation now than I did all last year.
The attitude thrusters in the Starship upper stage were cold gas thrusters powered by the pressure maintained in the methane tank. If that pressure failed, that's why attitude control was lost on-orbit before entry. Yes, you can fix the pressure regulation system, but the surest way is to separate the attitude thrusters from the tank pressure system. For that, you need propellant-type thrusters, not cold gas thrusters.
As for the COPV failure due to "unexpected damage", all composites, but especially carbon composites, are very vulnerable to handling impact damage, and you cannot just inspect the surface and see it. Infamously, the damage is hidden within the material. The surest way to avoid this risk entirely, is just accept the extra mass and use a metal tank whose welds you can reliably X-ray. It's not that big a penalty because that is a relatively small tank.
Just food for thought, on my part.
GW
It sounds like you did the right sorts of things to define a best L/D trajectory. That would be very nearly a constant dynamic pressure trajectory, such as was studied decades ago. Yes, there's some variation of the angle of attack that gets you best L/D, as Mach number varies, but I do understand the concept you are trying to employ.
Lessee, Mach 0.3 at sea level would be a dynamic pressure of about 133 lb/sq.ft, which would be about 6.37 KPa, unless I missed a key somewhere. That's rather low. Is that about what you are trying to fly? They were looking at 1000 psf+ in the old days! So too, are you, I would think. Mach 1 at sea level is 1480 psf. I rather doubt you can come off the sled, manage to start climbing, and still quickly accelerate to a real flying speed, at any sort of optimal L/D. You must get onto the right trajectory at the right speed any way that you can, then you can control to best L/D (or to constant dynamic pressure, as a pretty good approximation).
My only caveats are (1) do you actually have the thrust to make that best L/D (or constant q) trajectory actually happen, which was always the bugaboo with it decades ago, and (2) while CFD is your only available option for estimating hypersonic aerodynamics and heat transfer, be aware it can still seriously lie to you about what happens inside hypersonic airbreathing engines. Just having a good turbulence model and heat transfer correlations is not sufficient to model all the processes going on inside all but the simplest propulsion devices. It can become a garbage-in/garbage-out problem pretty quickly.
L = CL q S for the lift force and L = W*cos(theta) - T*sin(AOA) is the normal-to-path force equilibrium, where theta is the path angle above horizontal. That means (W/S)*cos(theta) = CL*q for your trajectory, ignoring the T*sin(AOA) term, which should be rather small. In turn, if the best CL is near 0.5 in value as a guess, that says your wing loading is CL*q/cos(theta), or somewhere in the vicinity of 60-70 psf at fairly low path angles, which I do not find credible for a re-entry-qualified craft. Something at or above 100 psf is more credible.
Along the path, the force equilibrium (for no pathwise or cross-path acceleration) is T*cos(AOA) = D + W*sin(theta), and you must further exceed that equilibrium thrust T by the amount that will actually accelerate you to the next speed and altitude along your trajectory, at whatever mass you have at that time point. It takes something resembling a trajectory code to properly explore that.
Have you done anything to define the thrust requirements all along your trajectory, using that kind of free body diagram calculation embedded in some proper code? You should have, for any realism at all. Once that is defined, then you have to figure out what sorts of propulsion items you must have, that can supply those amounts of thrust, at all of those speeds and altitudes.
You will find the airbreathers start to fall well short in the thin air down nearer 30-40 km than up nearer 50 or 60 km. Their thrust is rather closely proportional to combustion chamber pressure, which is in turn a fairly fixed ratio to ambient pressure, for pretty much any type of jet engine imaginable. Once the air thins too far, you have no chamber pressure, because 3 to 6 times essentially nothing is still nothing, for ramjets and scramjets! Which means in turn your airbreather thrust is essentially nothing. And if the thrust force is near nothing, then the higher airbreather Isp is worthless to you! Simple as that.
I'm not saying your designs won't work, because I do not know enough about them to figure anything. I'm just saying you have to worry about a whole lot more than just L/D and Isp.
GW
Thanks, Void. I honestly thought it was hydrogen. I didn't go look it up.
GW
To fly at one and only one L/D ratio, you would have to fly at one and only one angle of attack, thereby maintaining a constant lift coefficient and drag coefficient, at just the optimum values. The only possible way to accomplish that is to fly at one and only one value of dynamic pressure. Theoretically, that is possible to control. Practically, it is not, especially down low where you are trying to get started.
GW
An air separation plant has nothing to do with producing methane fuel! It has everything to do with producing liquid oxygen, which for the Raptor engines, outweighs the methane by a factor near 6. It also produces more liquid nitrogen than they could ever use, which they could sell.
The methane for the Raptor engines comes from natural gas, which is a mix out of the ground of methane, carbon dioxide, water vapor, ethane, propane, butane, and some others, sometimes including even hydrogen and nitrogen. Methane is separated from the rest in a cryogenic chilling plant, as the purest of several products from such plants, such as LNG.
Your household or camping propane or butane comes from these same natural gas cryo separation plants. The ethane usually goes to the chemical industry. Power plants can use unseparated natural gas right out of the ground , although the "higher quality" fuel that causes fewer maintenance problems has been de-watered by the mildest of the chilling in the separation plants.
There are a bunch of these facilities on the Texas coast near the Houston ship channel, amongst all the oil and chemical plants. They are not owned in any way by anyone but the fossil fuel industry. They are not very far by road, rail, or ship, from Brownsville, Texas. Less than 400 miles.
The BE 4 engines in the new ULA Vulcan rocket burn hydrogen as the fuel, not methane. The oxidizer is the same liquid oxygen that the Raptors, and almost all other modern engines, use. Hydrogen is also made in chemical plants, by reaction, from natural gas.
Lets just say that with technical flaws like that, I have my doubts about the strategies and motivations claimed in that video.
GW
Photonbytes:
I've slowly been looking at you postings and references, trying to understand what Spaceplane Corporation is actually trying to do.
A question: does your ascent presume near-optimal L/D ratio all the way up?
GW
Your time comparisons are apples and oranges.
Reentry only takes 30 minutes if you time it from the reentry burn instead of the entry interface altitude of 140 km. From there, using a low-angle trajectory, it's about 3 to 4 minutes from entry at orbital speed to a fairly sudden drop from near-orbital speed to well-below orbital speed, the max-gee "point", which is several seconds long actually. The several-second-long max heating "point" precedes this by several seconds. Then within several seconds more, you are "out of hypersonics" at about Mach 3 speed, if you are blunt. That's 4-5 minutes from interface to out of hypersonics. It works this way because most of your descent is above half orbital speed, for a short time exposure to the most extreme entry heating. The peak gees is usually somewhere around 4 gees.
The ascent heating exposure is a lot longer than that, because most of your ascent time is for when you are below half orbital speed, and very little exposure time above it. While not as extreme, simple hypersonic heating is still quite severe. And you will NOT suddenly gain a lot of speed at 4-ish gees, unlike descent, because unless you are a rocket, there is no way your propulsion will support accelerations like that. The altitude is too high (around 40 km, near 140,000 feet) and the air too thin (almost a vacuum) for any airbreather to accelerate anything.
Using a rocket there is a waste of rocket propellant that should have been used instead on the non-lifting ballistic ascent trajectory that leaves the sensible atmosphere at only high-supersonic speeds, not even hypersonic.
GW
To quote the computer from the command prompt window:
"The system cannot find the path specified"
This is the 4th time tried. All results identical.
GW
It would appear that this spaceplane concept is intended to do a lifting flight to orbital-class speeds. That is essentially trying to fly reentry in reverse, more or less, something discredited since the X-15 days in the early 1960's. The drag will be horrendous at hypersonic speeds, and the aeroheating far worse. Even NASA knew better than to try that, when it kluged-up its spaceplane notion as the space shuttle: vertical rocket launch to get out of the atmosphere as quickly as possible, leaving it before you get high supersonic, much less very low hypersonic. Their trajectory was non-lifting ballistic, to get minimum drag. And it worked!
As for airbreathing vs rocket, there is far more to the picture than just specific impulse, and in NONE of these postings can I find a recognition of that. Ramjet and scramjet, both airbreathers, will have a thrust at any given speed that is at least approximately proportional to the ambient air pressure. This shows up as combustion chamber pressures only about 3 to 6 times ambient. Above about 100,000 to 125,000 feet, 3 to 6 times nothing is still nothing for chamber pressure (it likely will not burn at all, such low pressures), which in turn means essentially nothing for thrust. It WILL NOT accelerate nor will it climb, because vehicle mass does NOT decrease with low air pressure the way thrust and drag will. That effect is called "service ceiling", and rockets do not suffer from it. But ALL airbreathers do. Ramjet, scramjet, turbojet, piston, and airbreathing combined cycle, you name it.
Aeroheating during such an ascent is your other enemy. The total temperature in the stream adjacent to your vehicle is pretty close to the driving recovery temperature for heat transfer. At only Mach 3.5 in the stratosphere (below about 70,000 feet), this is 886 F. At only Mach 5 same altitudes, it is 1880 F. At Mach 10 it is 7730 F, which explains why the vehicle is fully surrounded by an ionized plasma sheath through which no radar/radio can penetrate and no visible light can see. Anything unable to cool adequately will quickly (in seconds) soak out to similar temperatures. And unlike reentry, your time spent at such conditions is orders of magnitude longer than 3-4 minutes of re-entry, precisely because most of your speeds are far, far below orbital class. And that last is why you do NOT want to try to fly reentry-in-reverse!
The max recommended soak-out service temperature of a carbon composite with organic binder would be about 200-250 F. Max for aluminum is about 300-350 F. Max for both titanium and low-carbon steel is about 700-800 F. Max for almost all the austenitic stainless steels is about 1200 F although 316 and 321 will go a little hotter around 1600 F, and 309 and 310 will go hotter still to almost 1900 F. Of those, if you must have good cryogenic properties, only 304L in plate and sheet is weldable. That's why most earthly cryogenic tanks are far preferred to be made of it, not aluminum! The exotics like Rene 41 will go to about 2200 F. But they have no cryogenic strength, and they are hard to machine and hard to join.
Don't launch this thing horizontally, and expect the wings to do you any good during ascent. Above about 100,000 feet, they simply cannot, even if they are huge! The air is just too thin! As I said, this has been known, theoretically and experimentally, since the X-15 days. Thinking you are going to do this single stage is just nonsense. It has been nonsense for over 6 decades now! Which never stopped people proposing, and the government funding, proposals to explore it anyway.
GW
I show a similar result, in terms of the "tank efficiency" R-factor. R = Wp/(Wp + inert), basically that fraction of the loaded tank that is actually propellant mass. The tank must contain the volume of the propellant, and larger is inherently a heavier tank, just because it has a larger set of surface areas. But if the density is low, that tank does not contain very much actual mass. For tank section L/D's in the neighborhood of 5:1, with stainless steel balloon construction, I'm showing overall (fuel and oxidizer tanks combined) R's near only 0.94 with LOX-LH2, when its nearer 0.98 or 0.97 with both LOX-RP1 and LOX-LCH4. The oxidizer-to-fuel ratio "r" figures into that, as well as densities of oxidizer and fuel. I was throwing in small allowances for a few frames and stringers in my tank inert masses, plus leaving the upper dome unfilled of liquid, in order to have a vapor pressurization space.
The effect of higher Isp with hydrogen is quite dramatic, probably the largest influence, but the tank efficiency effect is also significant, as it directly increases your stage inert mass fraction. Very quickly, you run out of payload fraction as you try to impose a large dV upon your stage with lower-Isp propellants. Hydrogen does not buy you as much as its Isp might suggest, but that higher Isp effect does offer a bit more payload fraction, despite the higher tank inert.
The problem with all this is the VERY strongly-constrained nature of the stage design problem: the vehicle as launched must be a low drag "clean" shape of the right L/D, or you must start doubling your estimate of the drag loss to cover.
The thrust/weight at liftoff has to be high enough for the vehicle to reach significant subsonic speed while still at very low altitude, in order not to burn up the majority of its propellant before it ever hits the speed of sound. Otherwise, you will be doubling or even tripling your expected gravity loss you have to cover.
Empirically, that liftoff acceleration must be about half a gee net, above weight, or T/W = 1.5 at liftoff. But, the actual engines as sized to create that thrust level still have to fit behind the stage, or else you end up having to double your drag loss that you must cover.
Increasing those losses drives up the dV you must demand of the stage quite dramatically, which drives up mass ratio and propellant mass fraction. At any given inert fraction, you run out of payload fraction VERY quickly! And lower Isp makes that worse even quicker!
You see the mathematical problem here: this is a VERY strongly-constrained optimization, not a free optimization for which you can just run one simple rocket equation calculation.
GW
Offtherock:
Welcome to the forums!!
GW
I've recently been hearing a new type of testing out of SpaceX's McGregor site. These are repeated start-stop tests. It starts, burns several seconds, then shuts down for several seconds, then restarts. This process repeats several times. I see no steam or plume clouds, so it's not on the tower stand or the two big surface deluge stands, it has to be on one of the horizontal open-air stands. The level of loudness and a hint of combustion roughness suggests these are Raptor tests. Odds favor these are Raptor 3 development tests, but that's just a guess on my part. The shutdowns have a sort of "bang" sound to them, not very loud. Every now and then I hear a loud one, though.
GW
TH:
The forums refused to let me log out today. This happens fairly often. I click on the logout button and nothing happens, repeatedly.
GW
The title of this thread was "oil, peak oil, etc". We would have already seen peak oil, except for the successful advent of fracking, which among other effects, made the Permian Basin a rich supply again. Oil recovery there was drastically ramping down, until fracking made so much more of the resource available.
There are two prices to be paid for that. (1) initially, we needed immense amounts of fresh water for the base of fracking fluid, when there already isn't enough fresh water for the population. Frack fluid comes back up the well about 10 times as salty as seawater, contaminated with heavy metals, and slightly radioactive. (2) used frack fluid disposal by deep well injection has been confirmed to cause earthquakes, with magnitude increasing as you inject higher volumes faster in smaller regions.
They are only recently beginning to re-use frack fluid. That helps with cost (1), but there is still a long way to go. And the recent floods notwithstanding, there is still a risk of not enough fresh water supply in Texas.
All you have to do to lower the induced earthquake risk is not inject so much so rapidly, and also spread it out over much wider regions. That costs a bit more, which corporate CEO's object to, but with the verification of too much disposal injection causing earthquakes, the liability from earthquake damage would cost them more.
So, oil and natural gas are going to be with us for a while yet. Along with nuclear, wind, and solar. And for a while yet, coal.
Don't complain about the subsidies for wind and solar, without also complaining about the subsidies fossil fuels also get! That often gets lost when people make political arguments! The depletion allowance is almost 150 years old now. And most of the necessary environmental clean up costs do not appear in the prices of oil, gas, and coal; they appear instead in your taxes. So you still pay it, regardless of how and where it appears.
I think Obama was right, the correct strategy is "all of the above" for the shorter term, transitioning into majority nuclear and renewables, in the longer term. It takes decades to significantly change large industries, and fortunately that started a couple of decades ago. Which is why about 25% of Texas electricity is wind, with a tad of solar. And rightly so, based on the prices! The variability of the renewables is what limits their percentage: you need transient surge capacity, best offered by natural gas fueled gas turbine plants.
I would like to see natural gas replace coal as power plant fuel a bit faster than has been happening, but not dramatically so. Gas burns much cleaner than coal, and does less environmental damage in its extraction than coal. And there is no ash to dispose of after burning gas. And MW for MW, there is less CO2 emitted per MW produced by gas, than by coal. And less acid from the sulfur that is more-or-less inherent in coal, but not so much in gas.
The biggest problem faced by nuclear is disposal of used fuel, made worse (by roughly a factor of 10) because we do not recycle it. Even so, we built the disposal facility on the old nuclear test range in Nevada, but we never used it! Instead, spent fuel rests in cooling ponds at every nuclear plant, some not inside containment. How incredibly stupid is that? Politics demanded that stupid outcome, not anything technical.
As the grid scale storage solutions actually come on-line, the percent renewables can increase, but not until then! And we will still need surge plants using natural gas for decades to come, because nuclear as we know it now, does NOT surge well at all!
It would behoove us to fix the leaks in natural gas production and transportation venues. Gas is a bad greenhouse agent. Again, the costs of actually doing that are what CEO's object to. And they will continue objecting, until made liable for the damage those leaks do. The regulations for that are still largely lacking, and being resisted politically. We’ve seen this movie many times before!
Energy sources-wise, it's not an either-or thing. It's just doing what you know how to do, but doing it as responsibly as you can (which often requires regulations that can be enforced). All the while knowing that both knowledge and technology improve over time, especially if you do not cut the research funding!
And, as the motor fleet electrifies, grid capacity is going to have to increase significantly. There is no way around that, there is only how fast we really do it. As I said, changing industries takes time! Best to get started soonest! The climate disasters are already upon us, and will worsen for decades to come.
GW
Photonbytes:
I do not yet understand what this "airbreathing rocket" is, or your overall propulsion concept. But I gather from your postings that its performance is based on CFD predictions.
I am aware that CFD can do marvelous things these days, but also that it still can only predict from the models built into it. Combustion is not yet fully modeled by anybody, except by "assuming the answer", that being a burn based on mixture in the cell. So the CFD predictions can still be way wrong! The gold standard for CFD propulsion models is still actual test with physical hardware.
I'm no expert in CFD codes, but I do know about that circular logic fallacy when it comes to combustion models. Too many people blindly trust the computer these days, when they really need to be open-minded skeptics. It works pretty good for external aerodynamics, even at entry conditions. Not so well regarding combustion, especially at extreme conditions.
As I said in the other post, I'm an old retired guy, but still able to do a bit of consulting. I started out in the slide rule days. I did rockets, ramjet, air turborocket, pulse detonation, and some other things, plus vehicle aerodynamics and flight dynamics, heat transfer (even hypersonic), some stress-strain, and a whole lot of other things, too.
GW
The old Titan-II was a two-stage vehicle. It was used both as an ICBM and as the booster for Gemini. It used NTO-hydrazine blend in both stages, as I recall.
The old Atlas was almost but not quite an expendable SSTO. Atlas-D was the version used most as both a ICBM and as the orbital booster for Mercury, and as the booster for the Agena docking vehicle used during Gemini. It was one set of kerosene and LOX tankage, made as an inflated 300-series SS balloon, with initially 3 engines burning at liftoff. Later in the trajectory, after it bent over a lot, and after a lot of the propellant had been used, it dropped off two engines in the engine bay skirt, leaving only the exposed center engine burning. That weight reduction allowed it to reach orbit with smallish payloads like the 4000 lb Mercury capsule.
GW
It sounds like you are talking about a combined-cycle ramjet-rocket design of some sort. Don't confuse ramjet with scramjet. Only the external compression and capture features of the inlet are the same. Everything about scramjet downstream of capture is geometrically incompatible with ramjet. Your outfit might want to talk to an old retired guy like me. I did a lot of rocket and ramjet work, some air turborocket work, even a tad of pulse detonation work. I also designed high speed vehicles, and did hypersonic heat transfer. I might be of some help as a consultant.
GW
The rationale for even trying scramjet is that keeping the air supersonic lowers its thermodynamic static temperature, delaying ionization. That is the fuzzy barrier that ramjet faces at about Mach 6: combustion energy release starts partitioning more and more into ionization instead of static temperature rise. You do not get ionization energy back out as speed in a nozzle. I'm unsure as to why; I just know that it does not work, experimentally. Probably a severe residence time mismatch, but that's just an educated guess.
What that really says is that scramjet also faces the same kind of ionization-induced speed barrier, just at a far higher speed. Another educated guess says that's why the "speed limit" quoted for hydrocarbon-fueled scramjets is somewhere near Mach 10, when hydrogen-fueled scramjets can fly faster (nearer Mach 20). The species from hydrocarbon combustion are different and some of them ionize easier.
When you add some sort of pre-cooler to the system to cool the air, you delay the onset of ionization problem to higher speeds. But you do it with a loss of some of the energy in that captured air, unless you can contrive to put it directly into the fuel to be burned. There are limits to that, and also process inefficiencies. Those act to limit what you can do and how fast you can really fly. I suspect those are what limits the Sabre engine to about Mach 5 in any practical sense, and maybe Mach 6 in a best-case theoretical sense.
Yeah, you might add a pre-cooler to a ramjet, but if you did, you would negate its main advantage of very high thrust/weight. That pre-cooler stuff is heavy. It only works precisely because it has to use hydrogen fuel. That's also heavy.
--- UPDATE 7-29-2025: The tanks are heavy because they have to be so large, liquid hydrogen being of such low density. But the pre-cooler DEPENDS FUNDAMENTALLY on the fuel being liquid hydrogen! There is no fast pre-cool without the intense cold of liquid hydrogen. That's not me saying that, it's the Sabre engine people!
NASA's X-43 scramjet tests were 3 second burns of gaseous hydrogen, 2 successes for 3 attempts, one at almost Mach 7, and the other at almost Mach 10. These were rocket-boosted to test speed, and done above 100,000 feet where the thrust and drag were quite small compared to weight. Why? To lower airframe aeroheating. These did NOT accelerate as airbreathers.
USAF's X-51 scramjet tests were 3 minute burns of JP-7 thermally-stable kerosene, 2 successes for 4 attempts, both at Mach 5. These were rocket-boosted to test speed, and done above 100,000 feet where the thrust and drag were quite small compared to weight. Why? To lower airframe aeroheating. These did NOT accelerate as airbreathers.
UPDATE 7-29-2025: JP-7 was a unique early version of thermally-stable kerosene, used in the SR-71 variants to resist wet-wing exposures to very hot skins in the 400-500 F range. The X-51 tests used up the last of those stocks, there is no JP-7 being made anymore. Even so, thermally stable kerosenes are being made, but today they are additive packages in JP-8.
In contrast, the one hypersonic flight of ASALM was an accidental throttle runaway on the very first flight test. This was a liquid ramjet burning RJ-5 (a.k.a. Shelldyne-H), a synthetic resembling very strongly kerosene, but slightly denser than water. It was intended to cruise at Mach 4, at 80,000 feet, and then average Mach 5 in a terminal dive onto its target, not having time to heat up and melt itself. The runaway happened in ramjet after boost to takeover, launched at only 20,000 feet at Eglin AFB over the Gulf, and accelerated away dramatically in ramjet to just about Mach 6. It was still accelerating when it ran out of fuel, reporting by telemetry that its skins were beginning to melt.
UPDATE 7-29-2025: We lost the bird for 3-4 days, then found it stuck like a giant steel dart, in a farmer's field, about 10 miles outside of Eglin AFB. ASALM-PTV was 20 inches in diameter, and about 15 feet long. The weaponized form was longer, with a longer fuel tank section.
This was a wingless chin inlet design with a wave rider flattish bottom (a technology that goes back to the F-100 Super Sabre in 1953, by the way). It was aerodynamically very clean, so its top speed was likely somewhere between Mach 6 and 7, we'll never really know. It was made of martensitic stainless steel construction, limited to about Mach 4 steady state by aeroheating. Made instead of something like Rene 41, it might have survived steady-state at about Mach 5.5-to-maybe-6. The inlet chin cowl and buried subsonic air duct passage would have been most at risk from the aeroheating.
GW
"Air breathing rocket" is a very loose term usually applied to a combined-cycle engine concept combining rocket with turbojet. The other often-proposed combined cycle engine concepts are for hypersonic atmospheric flight, combining turbojet with ramjet or scramjet. Those are different. Both types have been tried many times, and are still being tried. None are fully operational yet, even after half a century of trying. Something flying experimentally is just that: experimental. That would NOT be ready for general application.
Do not confuse "air breathing rocket" the combined cycle rocket-turbojet with either the "ducted rocket" or the air turborocket. Those are different, and the term "ducted rocket" is really vague and misleading. The air turborocket has a fuel-rich gas generator combustor where part of the fuel and some oxidizer are burned to create not-fully-hot gas to drive a turbine, which drives an air compressor that draws air into the engine like a turbojet. The compressed air, the remaining fuel, and the turbine drive gas are then burned together in what amounts to an afterburner duct leading to the propulsion nozzle. Since the afterburner pressure is lower than what is usual in a turbojet combustor can, specific impulse is far lower than turbojet not afterburning, more like a simple ramjet or a turbojet in full afterburn. But it generates static thrust, unlike the ramjet. It is NOT a hypersonic propulsion device: that air compressor is limited to Mach 3 to 3.5 at most, just like turbojets.
The term "ducted rocket" is most often misapplied to either an air augmentation duct about a rocket, or to a fuel-rich solid propellant gas generator-fed ramjet. The air augmentation ring on a rocket acts as an air ejector duct, raising thrust and specific impulse, but only at very low flight speeds. This concept does not work in any practical sense at supersonic flight speeds.
The gas generator in a gas generator-fed ramjet does the same thing that the liquid fuel tank does in a liquid-fueled ramjet: it contains the fuel. With the fuel rich solid propellant, you gain the "wooden round" handling and logistics of a solid rocket, but at the specific impulse of a ramjet. It has high volumetric loading of that energy, but specific impulse is reduced some by the lower heating value of the fuel.
The "solid fueled ramjet" has oxidizerless solid fuel packaged within the combustor. It has inherently low volumetric loading of that energy because the fuel grain must have a large open bore to pass the air, and it needs yet more combustion volume downstream of the fuel charge to enable efficient combustion. But it offers the "wooden round" advantages of the solid plus similarly high heating value and specific impulse as the liquid ramjet. You add oxidizer to the fuel to increase the regression rate, and you negate many of the "wooden round" and heating value advantages.
GW