You are not logged in.
In point of fact, funding for "SR-72" is USAF, not NASA.
This is most likely a gravy-train program for Lockheed-Martin. I doubt this will produce a real operational aircraft, because the combined-cycle engine they describe is still more science fiction than science fact. The money will get spent trying to make an incompatible set of propulsion technologies work together, not to actually fly anything.
To fly Mach 5 to 6, you don't need any of that combined-cycle stuff. Nor do you need scramjet (which is still very highly experimental). You don't even need the heavy-cored gas turbine. All you need are rockets that get you to Mach 2-2.5, and a ramjet that takes you from there to Mach 5 or 6. That job can be done in an entirely-fixed propulsion geometry, no variable inlet, no variable nozzle. Use the rockets to land not quite deadstick.
The heat protection required for flight beyond Mach 4 is unlike what they do in reentry, which is fundamentally heat-sinking through a 2-3 minute transient event, whether done ablatively or refractively. For steady flight, there will be a conductive load into the interior that is the difference between convective input and hot-skin re-radiative passive cooling. The key to make this work is skin surface thermal emissivity > 0.80, while limiting the temperature such that the material still has the strength to bear structural loads.
The steels and superalloys have far-higher usable temperatures (1200 F +) than titanium (750 F), by the way. Inlet structures are harder to protect than skins or even leading edges, because some of these surfaces have little or no view of the environment by which to radiate. You'll be speed-limited by the higher convective heat transfer at lower altitude. Mach 6 is for well above 80,000 feet.
The most practical way to arrange the rockets and ramjet is parallel-burn, with a mix of rocket types. At speeds like that, there is no such thing as cooling air, so reusable structures for the combustor and nozzle are essentially impossible, unless they are complicated liquid-cooled items, expensive, and heavy. That's just not a recipe to go fly.
Instead, build the combustor/nozzle unit as an easily-replaceable item, just like a big JATO bottle. Do it an an integral rocket-ramjet with a solid booster sized to get to ramjet takeover speed at relatively low altitude. Line the combustor and nozzle with ablatives, just like missiles. Keep an inventory of these IRR things on hand at the bases, and they just unbolt, slide out the tail. New ones just slide in and bolt-up. You'll need a stab-in slide joint to connect the inlet with the combustor. Leave the fuel injection stuff in the inlet, and put all the ignition pyro with the propellant in the IRR combustor. That takes care of ordnance safety, as well as practicality.
In parallel alongside this replaceable unit, add two small pressure-fed liquid rockets whose thrust combined only slightly exceeds the weight of the vehicle. These burn the same jet fuel as the ramjet, and use either IRFNA or high-test H2O2 as the oxidizer. Hypergolic ignition is what makes this practical. These burn in parallel with the ramjet to make a near-vertical climb to mission altitude possible, without burning off all the ramjet fuel. They also serve as go-around or divert propulsion at landing.
No real development of anything there, just adapting existing stuff together in a new way. That's a proper recipe to go fly.
GW
Last edited by GW Johnson (2017-07-15 09:48:21)
GW Johnson
McGregor, Texas
"There is nothing as expensive as a dead crew, especially one dead from a bad management decision"
Offline
Follow-up regarding the TBCC engine concept illustrated, 7-18-17:
This TBCC engine concept is still science fiction and unlikely ever to come true, because (1) the common inlet geometry is impossible, (2) if the turbine core is fully isolatable to protect it, you might as well do parallel-burn of entirely separate devices, and not compromise either propulsive item’s performance, (3) even plain scramjet is still far too experimental to apply, (4) the dual-mode scramjet is far too-compromised to serve, and quite distinct from the dual-combustor scramjet, which could serve once scramjet is ready to apply, and (5) the common nozzle concept is as geometrically impossible as the common inlet concept.
Common inlet problem: the geometric features are identical for scramjet or supersonic gas turbine up to the inlet throat. From there they are entirely incompatible. The scramjet demands a long constant average Mach number isolator duct, which is very nearly constant area. Without it, the scramjet will not function, and may even explode.
The gas turbine requires a well-subsonic Mach number at its compressor inlet face, period. No way around that. This requires the inlet duct between inlet throat and compressor face to be strongly divergent, with a terminal shock wave standing in it. No way around that.
So you ether build two vastly different ducts fed “somehow” by the same supersonic air from the inlet throat, or you do an incredibly complex, heavy, expensive, and likely-unreliable duct variable geometry design.
The common nozzle incompatibility is similar: the gas turbine exhaust is modestly-high subsonic, and requires convergence to a sonic throat for best thrust, followed by a slight supersonic divergence. The scramjet requires no throat at all, its combustor exit stream is already well supersonic. Its nozzle is entirely divergent supersonic expansion. You cannot have that in one device without heavy, expensive, complex, and likely unreliable variable geometry.
Isolatable turbine core problem: if you isolate your turbine core to protect it from overheated air, then you must bypass all the air around it. There are two very serious problems doing this: (1) this effectively doubles the cross-section of your engine installation, and (2) it is a complex, expensive, heavy, and likely unreliable variable geometry problem. So, if your frontal cross section doubles anyway, then why not just do parallel burn with entirely separate engines of the two types, and avoid all these complications and performance compromises?
Plain scramjet is not ready to apply: NASA’s X-43A scramjet flew twice of 3 attempts, for two 3 second burns (one near Mach 7, the other near Mach 10) with impractical hydrogen gas fuel, in which neither test demonstrated any vehicle acceleration capability as an airbreather.
USAF’s X-51A flew twice of 4 attempts, for two 3-minute burns (both at only Mach 5) of hydrocarbon fuel, in which neither test demonstrated any vehicle acceleration capability as an airbreather.
A few other countries have achieved similar very-restricted results. This technology is simply not ready for general application.
In contrast, plain ramjet has flown for decades to Mach 4 or 5 with very practical hydrocarbon fuels, demonstrating very strong vehicle acceleration capabilities, at least at lower and middle altitudes. One even reached Mach 6.
The dual-mode scramjet problem: this is a concept that first operates as a ramjet (subsonic combustion), then as a scramjet (supersonic combustion). Ramjet requires a subsonic air duct approaching its combustor, the same as the subsonic air duct approaching the compressor face in a gas turbine.
Ramjet also requires for its nozzle the same convergence to a sonic throat with mild supersonic divergence as does the gas turbine. Neither the inlet nor the nozzle geometry is compatible with scramjet, same as gas turbine.
This dual-mode scramjet idea is thus utter nonsense without enormously complex, expensive, heavy, and likely-unreliable variable geometry for both its inlet and its nozzle.
That is quite distinct from the dual-combustor scramjet concept, wherein about 25% of the inlet air goes through a fixed geometry subsonic duct to a subsonic combustor with a convergent nozzle embedded within the scramjet combustor. The other 75% of the air is fed to the scramjet via a supersonic isolator duct. The scramjet has a divergent-only nozzle. The embedded ramjet combustor is the pilot light and vapor fuel generator for the scramjet. This is entirely fixed geometry.
The SR-72 engine concept would be much more believable if it were a gas turbine combined with a plain subsonic-combustion ramjet. At least the inlet and nozzle geometries would be compatible.
It would still suffer from the doubled cross section problem of bypassing 100% of the captured airflow past the turbine core. But with a belly-mounted installation, that bypass could be accommodated at the cost of volume within the fuselage.
But, that is NOT what they proposed! They proposed something virtually impossible to do, which effectively means that all the money gets spent on trying to make the engine work, and none on actually flying anything. And THAT is why I called this nothing but a “gravy-train” program for Lockheed-Martin.
GW
GW Johnson
McGregor, Texas
"There is nothing as expensive as a dead crew, especially one dead from a bad management decision"
Offline
GW,
What about the supersonic compressor that RamGen is working on?
Considering who is providing the funding for that project, it's hard to imagine that DARPA, the US military, and Lockheed-Martin are all unaware of the technology. At the very least, the inlet could be much smaller and there's no requirement for subsonic flow.
Offline
Well, the difference between "being worked on" and "read to apply" is light years. When a supersonic compressor is "ready to apply", the picture will change. Somewhat. Depends upon what supersonic Mach number at the compressor face can be tolerated, versus what is needed for scramjet operation.
The supersonic inlet "throat" Mach number for a scramjet will fall in the 2-4 range, or perhaps even higher. That gets followed by a mandatory constant-Mach isolator duct into the scramjet combustor. There are no turns allowed, as these invariably cause strong oblique shocks and deceleration.
For an ordinary ramjet or gas turbine installation, the inlet throat Mach number is in the 1-1.2 range, followed by a divergent diffuser duct. The forward part of that divergent duct has a shock train terminating in the equivalent of a normal shock, followed by subsonic deceleration as the duct diverges. Ramjets and gas turbines typically have end-of-diffuser Mach near 0.3-0.4.
If there were such a thing as a "ready-to-apply" supersonic compressor, I'd hazard the guess its compressor face approach Mach is closer to 1.2 than it is to 2+. That means you can use an inlet with a Mach 1.2-ish throat and a straight constant-Mach duct, but it's still the wrong throat Mach for scramjet. Extensive inlet variable geometry is therefore simply required. And the problem of shocks when you divert air from turbine to scramjet and back is still entirely unresolved.
I don't think I'd hold my breath waiting for this stuff. There are some very severe, and very fundamental, incompatibilities in such ideas.
As for Boeing and Lockheed-Martin knowing about such things, yes, they are counting on it for the never-ending series of gravy-train cost-plus contracts that fund "working on it", instead of actually flying anything. They make more money that way than they do building real flying machines. I've been hearing noises out of Lockheed for over 30 years now about this or that concept for an "SR-72" to replace the SR-71. None is flying yet. The -71 is long retired now.
And I put my evaluation of how unready the scramjet technology still is in the post above.
GW
Last edited by GW Johnson (2017-07-19 09:43:24)
GW Johnson
McGregor, Texas
"There is nothing as expensive as a dead crew, especially one dead from a bad management decision"
Offline
GW,
I don't know what the upper limit is for the compressor, but I believe it's been tested to Mach 2.4. DARPA and AFRL funded work on RamGen's advanced vortex combustion for flight applications. It's intended to provide combustion stability in a turbulent flow. They supposedly have a supersonic expander, too, whatever that is.
The decision to retire the SR-71 fleet came from Congress and the US Air Force. Lockheed-Martin was all-too-happy to keep the Blackbirds flying for what was a pittance. There's been no real foresight in defense asset procurement and retirement decisions in the past several decades or so.
Offline
Well, if the supersonic compressor can tolerate inlet face Mach numbers higher than 2, then the inlet throat mismatch that I brought up in the post above more-or-less goes away. The fundamental nozzle mismatch is still there, though. As well as how sensitive this compressor is to off-design conditions.
If memory serves, I saw something a couple of years ago that said this supersonic compressor technology had extremely-narrow operating limits, and that seemed to be rather fundamental to the concept, not something easily remedied. Perhaps that assessment has changed, but I have not heard that it has.
The more important question is just how ready is the supersonic compressor technology, to be generally applied? Or is this stuff just barely demonstrated to be feasible? There's light years between those two readiness states, and the NASA TRL classification thing has little to do with evaluating that. When folks like GE, Pratt, and Rolls-Royce start offering supersonic compressors for military engines, then you know it's actually ready to apply. Not before.
Until then, when you propose such a technology as a critical enabling item for a supposed aircraft development, you are really proposing a technology development program, not a build-and-go-fly program. It really is that simple. And I am surprised and disappointed about how many of those things get sold (in public) as go-fly programs. Deceiving the public about how its tax dollars are being spent is never a good thing.
GW
Last edited by GW Johnson (2017-07-21 10:08:19)
GW Johnson
McGregor, Texas
"There is nothing as expensive as a dead crew, especially one dead from a bad management decision"
Offline