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It was in another thread somewhere on these forums that I described the fundamental change between the NASA of 1958, and the NASA of today. But I’ll repeat it here, because that’s what the most recent conversations in this thread seem to be about.
NASA was originally formed to put man in orbit, and carry out science and aeronautics, too. But the “prime show” or “front-burner project” was manned spaceflight. 3 years later, the mission got upgraded to the far more demanding man-on-the-moon, which really energized the little agency.
In those days, NASA was rather small, very heavy on engineering talent, had a definite front-burner mission, money was no object, and no one told them how to do their jobs. They got to figure that all out themselves. And, miss-steps notwithstanding, it worked quite well. It was 8 years from assigning the moon-as-mission to the agency, until two men first walked there.
It might have taken perhaps 3 extra years to do this, had budgets been a problem, but that basic approach of assigning the mission and then “stand back and let them do it” works really well either way.
All that changed in the middle of the moon landings in 1972 (there should have been missions through Apollo 22, not 17), when Apollo got cancelled early and all manned flight outside Earth orbit forbidden by presidential order. NASA has never had a front-burner mission, an agency reason-to-be, ever since. They have only had major projects mandated upon them mostly by Congress, with some from the various presidents. Projects like Space Shuttle, like ISS, like X-30, like X-33, etc.
Science and aeronautics are still small-time background, but by dint of the successes of the probes (derived from being left mostly alone by Congress), the planetary probe program kind-of falls in-between in that spectrum. Some of these projects, like the Mars landers and Hubble, turn out to be quite popular with the public, too popular to kill, even though Congress often tries.
Two of these mandated projects flew men in space, the rest didn’t. There were some tests as exceptions that never led anywhere. None of these were actually managed in an overall sense by the agency. Instead, the project, its detailed objectives, how it would be done, and where things would be built (by that I mean in whose districts) were all mandated by Congress. That’s exactly what Constellation was, and what its resurrected form Orion/SLS is. There is no one in Congress at all competent to do any of this work, which is precisely why their mandated management plans are so egregiously ineffective and nonsensical.
Meanwhile the agency has grown to enormous size, trying to be “everything to everybody” in lieu of a front-burner mission / reason-to-be. Once an organization gets too large, it gets very inefficient, worrying more about preserving departments, people, and budgets than actually doing anything real anymore. NASA is no exception. There’s more managers and support functions at NASA these days than there are real engineers. That’s not a good recipe to get anything major, actually done. Not in industry, not in government.
When you add bureaucratic inertia to mandated-but-nonsensical-projects to be done, you have what we see now: no man has flown beyond Earth orbit, or explored anything off-world, in person, since 1972. And with the projects they have to do sopping-up most of the available money, we’re having a hard time not spending trillions just to go back to moon, the same moon that we visited over 4 decades ago! None of the stuff they are doing now (with the serious money) can take a crew to Mars alive, much less land there.
I’m talking about Orion/SLS. The PR about that is nothing but lies. And everybody who knows much at all about these things knows it. Most within the agency are too afraid to tell the emperor that he has no clothes, though. And as an agency, NASA is afraid to tell Congress that it too has no clothes.
There are small groups within NASA that are working on the right kinds of things for men to go beyond Earth orbit again. But these are not funded with any serious money. Some of these groups are better managed than others. Some of these groups have better talent than others. So, their ideas and plans vary considerably in practicality and feasibility. That should not be unexpected, given the situation. But until these things get the money and attention to perfect them, they will take no one anywhere. That, you can count on.
That’s fundamentally why what is funded seriously often makes no sense. A lot of you on these forums seem to have noticed that. Nothing about that situation will change, until the operating model for NASA-as-an-agency goes back to that 1958 version. They need a front burner mission as a reason-to-be, and they need to be left alone to accomplish it. Period.
Unfortunately, Congress craves ever-more control, not less. Ergo, no change is forseeable.
GW
The video they released shows it right over the deck coming down fairly slowly, but cocked over at a rather large angle, something like 40 degrees off vertical. The accompanying news story said one of the four grid fins had quit working, and that the gimballing rocket motors (motor?) couldn't compensate for it. The story said the fin failed for lack of hydraulic fluid. (It's my understanding that only one of the 9 engines is firing at touchdown, but that all are gimballed for steering when operating.)
That sort of sounds like a hydraulic leak, without enough reserve to make up for it. Common enough failure mode. Edit: the hydraulic press in my shop recently blew a rod seal. It loses lots of hydraulic fluid unless I park it in just the right position.
Grid fins work subsonically like an extremized form of biplane and triplane wings. The channels forcibly redirect the airflow to a new angle, which is lift. Supersonically it still does that, but at large drag due to the big normal shock standing out in front of the grid fin unit. These are operational on the Russian AAM that NATO designates AA-12. Supersonically-draggy fins are "perfect" for what Spacex is attempting.
GW
Edit: Looks to me like they very nearly pulled this landing off. Kudos to them! It's about the only imaginable way to recover a first stage with a modern stage inert weight fraction in the vicinity of 5%. In terms of landing loads, that's a very flimsy structure.
Tom said: "By the way, what politician is not a publicity hound? Publicity is how politicians get elected." I quite agree. But I would add this: The trouble we have now is that campaigning is completely out of control at 90+% of term in office. That's no time left to actually do the job we pay them for. A job which they don't want to do, because of the necessity to extremize everything for purposes of campaigning.
GW
The asteroid re-direct mission is a mission for Orion/SLS to go and do with humans, nothing but a reason-to-exist. It is a mission that falls within what Orion/SLS can actually do, which is inherently restricted to cis-lunar space and no more than 2-3 weeks cooped up inside a capsule. Scientifically, it makes little, but not non-zero sense: a way to get lots of samples from one body, but not a way to sample representatively an excruciatingly-diverse family of bodies. That second choice is what you really want to do. It supports both asteroid mining and asteroid deflection. Orion/SLS cannot ever do that as a one-rocket/one-mission shot, which is the mission model Congress seems hung up on.
Anything that could successfully take men to an NEO in-situ could take men to Mars. Mission times and velocity requirements are quite similar, and way, way beyond what Orion/SLS can do in single shots. (Used as a launcher for a transport to be assembled in LEO, Orion/SLS could make sense if it could be launched for commercial prices, but it cannot and never will be.) The difference between those missions is a lander for Mars (why go all that way and not land?). Sampling multiple NEO's in-situ, which is what should be done first, is more properly an unmanned mission, because the flight times are too long for human crews (one example only: Ceres-Vesta probe). That kind of things gets smaller samples, but identifies where men should go to explore and get lots of samples.
GW
Third cent: propellant ullage. Why use a big motor like OMS to do a small motor's job? All you need is a very few hundred pounds of thrust acting on something the size of a Saturn S-II second stage. They did it back then with small solid propellant cartridge motors shaped like a thick pancake, about 5 inches in diameter, and about an inch and a half thick. There were 3 for redundancy. I crudely estimate 300 lb thrust each. Today you would do the same job with a few very small hydrazine thrusters, if you wanted something reusable. If it's one-shot, use the solid, it's cheaper and stores "carefree".
GW
Two cent's worth/first cent: I think the upshot of the postings about using propellant tanks of any kind as human habitation space, shows that flying it as a tank and modifying it in space has very serious penalties associated with complexity, weight, and safety. Modifying it on the ground to habitat configuration and then launching it dry to space as dead-head payload will work, and is what we did on Skylab. I'd hazard the guess that a purpose-built habitat riding as payload will be the better option, since there is no longer a surplus of Saturn S-IVB stages laying around unused. Pulling stuff out of museums is not a solution at all. The Bigelow inflatable approach has an awful lot of promise.
Second cent: when judging the wisdom (or lack) of what NASA did or does, you have to remember both (1) that and (2) why the NASA of post-1972 is not the NASA of 1958. Early on, a small organization heavy on engineers was tasked to do things fast, sparing no expense, but was not told how to do them. That actually worked quite well. NASA was the manned space program, with science and aeronautics sort of in the background. Plus there were dozens and dozens of contractors to choose from in letting contracts.
After 1972, the President and then Congress started telling them how and what to do in great detail, which micromanagement inherently strains budgets that have become the deciding factor in most decisions. The same micromanagement allowed the start of "consolidation" of all those contractors toward a monopoly. Budget pressure led to the side-mounted cluster of a shuttle that eventually killed two crews, and cost around $1B per launch to put at most 15 tons in LEO for ISS.
Today, NASA is a gigantic organization, filled with inertia, and totally micromanaged by Congress in terms of projects and money. It lets contracts mostly to the ULA monopoly (which Spacex is attempting to break). Only the big projects that Congress mandates get big budgets, there are other groups within NASA doing good things, just on a shoestring. The planetary science guys are sort of in the middle of that spectrum, but being neglected, are more successful because of less congressional micromanagement.
Overall, the organization has been trying to be "everything to everybody", for the last couple of decades at least. That never works out well. Without change, the dinosaur will very slowly die. That change has mainly to do with micromanagement coming from Congress instead of a real human mission to be done.
GW
Cruz is nothing but a self-promoter and publicity hound, but with enough legal education to make him dangerous.
He tells audiences only what they want to hear, so they'll vote for him. Speaking in Texas, he promotes space stuff. In DC, it's cutting budgets. See why he seems to say different things?
He also will do anything, no matter how bizarre, for publicity. He's really good at that. But not at governance.
GW
See paragraph 2 of post 53 above.
It went out of control when the grid fin steering system ran out of hydraulic fluid. Whether part is on deck and part on the bottom of the Atlantic does not change that outcome.
If they can maintain control to touchdown, then they just might be able to pull this stunt off. It'll be a lot easier on an island with no up-and-down "heave" motions.
GW
Not all of us in Texas think well of Ted Cruz. I find him to be a total embarrassment to the state of Texas. Too bad so many of my neighbors fell for his nonsense that he got elected.
GW
There's a reason the vast majority of the missiles out there (from any source anywhere in the world) are solid propellant rockets. It's because they are both the cheapest option that there is, and the most carefree that there is, when looked at from a life cycle standpoint. The second closest is ramjet, with turbine a very, very distant third.
Yep; solids are cheap. And they are operationally carefree. Better than anything else we know how to do. The world's weapon arsenals are what they are, for a very bloody damned good reason.
The expectations and preconceived notions for SRM's that most people seem to have, are at variance with what actually obtained on the shuttle program. Those motors actually did rather well, in spite of all the expensive impediments and out-and-out mismanagements, that NASA managers contrived to heap upon them.
A fair fraction of these case segments really were reloaded and reflown. (The loss rate just goes to prove that parachute landing in the ocean is unsurvivable, even at 10% inert fraction.) They turned out to be about as reusable as the shuttle orbiter itself, or more so, by about any measure that is quantifiable.
That's why both Atlas-5 and Delta-4 (ULA products) both use SRM's when the core thrust is inadequate off the pad. Low down in the atmosphere, Isp isn't all that important, it's nothing but brute force thrust that matters. There's nothing on this Earth that beats solids for brute force thrust, with the sole exception of nuclear explosions.
That being said, what you have to consider for reusable stages is how hypersonic you are at atmospheric entry for your recovery. Most cores are 5-10% inert mass fraction stages. These are very structurally fragile, even if aeroheating were not an issue, and it is.
Spacex Falcon 1st stages generally burn out near 10,000 fps (3 km/s), as all TSTO 1st stages do. In the absence of any other influences, that's the entry speed upon hitting air again (energy is conserved, after all), and it is not very survivable (shuttle Columbia broke up at about Mach 12, or about 12,000 fps or 4000 m/s).
You have a choice: either put a heat shield (heavy!) on this thing, or carry some extra propellant (also heavy!) to slow it down to a more survivable entry interface speed. Speeds like that are nearer Mach 3 (3000 fps, 1 km/s). Spacex has done the latter, and seems near to succeeding, far sooner than any of us ever expected. I am surprised and pleased by that outcome. Spacex has "done good"!
For second stage of a TSTO, entry speed is very near LEO orbital speed: near Mach 25 (25,000 fps, or 8+ km/s) at entry interface. If you think Mach 10 protection is hard to achieve, try this! It's worse (by far) than speed squared dependence on speed. So that's an all-up heat shield. No chemically-powered machine could ever afford the propellant to slow from there to near Mach 3 by retro-thrust.
It would make more sense to recover this 2nd stage in LEO after the circularization burn, if there were a use for it up there in LEO. There isn't a use yet, that's why no one has done it. Sorry, fact of life. But, it could be done, if there were a use. That kind of on-orbit re-use would be the cheapest and most technologically-feasible thing we could do.
There just is not yet a use for such hardware, and the orbital debris problem weighs against it until there is a specific use. Sorry, that's just life.
I think the shuttle proved that large spaceplanes are not a very good idea, for the kinds of technologies we have at our disposal. That says nothing about small ones, I might add. But large payloads are probably most efficiently (and cost-effectively) placed in LEO with big vertical launch rockets, for the forseeable future. If you can lower the price by recovering stage 1, fine. If not, well, that answer still stands.
If you need solid strap-ons to get your launch rocket off the pad with big payloads, well, fine. Do it. There's nothing wrong with that. Design them correctly (NASA still does not), and you can even man-rate them.
GW
That's a pulsed combustion device, not explosion propulsion. The mechanism is different.
The one in the article is an academic lab play-toy. But, these things have been experimented-with for missile and vehicle propulsion. The upshot of that was you have to build hell-for-stout (which is heavy!) to take the pounding, and it isn't likely you could use it for manned vehicles because of that pounding.
There are some rather severe restrictions on what kinds of fuels (or propellants) you can use that can be made to detonate. Think low molecular weights and gas phase only.
GW
2nd stage recovery on a TSTO is quite different from 1st stage recovery. The difference is orders-of-magnitude worse aeroheating. Falcon-9 (or heavy) 1st stages (cores in the case of Heavy) already retro thrust rather significantly to reduce what would have been a Mach 10-ish entry to something nearer Mach 3 or 4.
The second stage (either vehicle) would hit air in the vicinity of Mach 25 without retro thrust, or else would need a "real" heat shield (and a lot more airload gee resistance structurally) to do most of the job by aero-deceleration. The retro-thrust requirement is so huge (Mach 25 down to 3 or 4?) your only practical choice is a real heat shield and gobs of structural "beef", which is very heavy. I don't see that activity going on anywhere.
So, I don't think I'd worry too much about 2nd stage recoveries just yet. Spacex doesn't seem to be looking at it. Not for Falcon-9 or Falcon-Heavy. It wouldn't be any different, no matter how big the rocket, as long as it was TSTO.
GW
Spacex already knows what happened. The news reporters are technically rather ignorant, generally speaking. That's why the truthsome tidbits get left out of their stories.
I found one buried in today's newspaper: they simply ran out of hydraulic fluid for the steering grid fins before it landed. Presumably, it lost attitude control, which also costs them their retro thrust vector. The news story said they already upped the hydraulic fluid quantity by 50% for the next shot.
The barge has steering thrusters. Marine engineers use these quite frequently in modern passenger ships. These not only help control position, they also help a great deal with reducing pitch and roll. But they can do nothing for straight up-and-down "heave" of the wave action .
Bob Clark is quite right about Falcon-9R needing to hover. That capability coupled with a small radar to perceive the relative motion of the "heave" would enable self-compensation for a more reliable ship landing. That kind of dynamics is elementary to a Navy carrier pilot, but not to an Air Force pilot.
If however, Spacex really plans to land on an island instead, then what they are doing will work on dry land which does not move. Unless they add hover capability and the small radar, it seems unlikely to me that barge landings will be reliably successful in anything but the very calmest seas.
GW
I see in today's news that Falcon-9/Dragon launched successfully to ISS. The attempt to land and recover the 1st stage was only partially successful. They actually did hit the recovery barge, but they crashed hard upon it, near as I can tell. Still, they hit the target at all, that's significant in and of itself.
GW
OK, look. The dynamics of fast interplanetary travel demand propulsion that has at least a low-1000's of sec of Isp, and lots of vehicle acceleration capability: 0.1 to 5 gees. We have some technologies that have shortfalls, and we have those benchmarks. Not both! And we have some concepts we haven’t tried yet.
There are a bunch of electric propulsion schemes, some more mature than others. The best ones have the low-1000's of sec Isp, but all fall way short on vehicle acceleration: 0.001 gee or less. The most critical need is for a lighter-weight, more powerful electric power supply. That's what needs to be worked on. If the SLS budget were going toward that, we might have a solution in sight. But we do not.
Chemical and solid core nuclear thermal meet the thrust benchmark, but fall short on Isp. Many folks have objections to nuclear for any of a number of reasons, but we may just "have to get over it", and be careful how and where we employ it. Both technologies are essentially ready-to-use, especially the chemical. Here in the US, nobody is making LOX-kerosene engines but the newcomers like Spacex and XCOR. That needs to change. We also need to work on in-space cryogen transfers and long-term storage. The practicality for real interplanetary travel is not quite there yet, but could be.
Gas core nuclear thermal is absolutely immature, but holds the promise of meeting both the thrust and Isp benchmarks simultaneously. So, we ought to be working on it, too. That technology could enable travel if the search for a practical electric propulsion power supply fails. It's stupid to put all your eggs in one basket, as we all already know. If that word "nuclear" offends, well tough. Life ain't fair. Neither is physics. The electric power supply we seek for electric might well be nuclear. So, get over it.
BTW, my candidate power supply concept would be a nuclear thermal rocket device feeding an MHD generator. None of that heavy steam loop nonsense. Use it as a rocket for getaways and captures, use it for MHD electric on the transits between to raise midpoint speeds. Get thrust from the electrics and the released plume. Same rocket hardware does both, which saves vehicle weight. That might really start looking good if gas core nuclear was the rocket device. Then we could start looking at excess electricity to use for processing propellants while electric-thrusting on the transits. It pays to dream big. Tells you what you really need to be working on, too.
The only other technology holding the promise of meeting both benchmarks simultaneously (that I know of) is the nuclear explosion drive. We have known since 1959 that this scheme would work. The side effect to fear is not so much fallout from launches to LEO, but EMP effects, and by far! The devices make lousy blast weapons, and blast weapons make lousy explosion-propulsion devices. Trouble is, it has been ignored and forgotten since 1965. Plus, it is only efficient in very large (10,000+ ton) vessels. I suggest we ought to be working on this, too, but slower and longer term, with an eye toward giant colonization transports by next century. Could always be accelerated if we find we need it sooner.
OK, that being said, now return to electric propulsion. The problem is power supply, as I said just above. The thrusters themselves are also a bit short on basic device thrust/weight, even when you don't count the power supply, so that's another area we ought to be working on, as well.
But, a good power supply could make the simple arc jet feasible, which could use just about anything as propellant, at least in principle. To use oxygen-containing materials as electric propellants, will require solving the reactive-oxygen problem, so we should just belly up to the bar and get on with solving that issue. That solution makes in-situ electric propellants a lot more feasible-looking.
Some (few?) parts of Mars have big buried glaciers, we think, so there is massive ice we could mine, just not at every site where we'd like to base. Maybe not very many at all. Who yet knows? The atmosphere there is mostly CO2, which at those temperatures is not all that hard to freeze for easy storage. How about CO2 as an electric propellant? Could we make that work? Venus also has CO2, but it has hellish conditions and big gravity well, which make it not so attractive at this time in history.
Titan has a largely-nitrogen atmosphere and a very weak gravity well, although Saturn's is considerable. Could we use Titanian nitrogen as an electric propellant to support activities in the outer solar system? The surface "rocks" seem to be water ice, so you have that as well. Plus, methane and ethane in liquid form in the lakes. Could any of these work as electric propellants?
There's the ices on the surfaces of 3 of Jupiter's big moons. Water, CO2, perhaps ammonia, and some other things. Could these be used? The moons have weak gravity wells, but Jupiter's is enormous, as is its radiation hazard. Not sure how soon we might be capable of safely recovering those resources, but somebody ought to be looking at it.
Guys, the power supply issue and the what-can-we-use-as-propellant issue are the two biggest issues facing electric propulsion. Those are what we need to be working on, very big time. I just don't see that going on in anything being funded through NASA or the other space agencies, nor through DARPA or anybody else. Which is why I really don't believe any of those agencies are serious yet about sending men beyond LEO. Not even to the moon, or the SLS/Orion test schedule would look more serious about actually accomplishing something.
Solve those two issues with electric propulsion, and we can ride the ions where we want to go. Right now, the acceleration is way too low to use it for manned flight. The travel times are just too long. Vehicle acceleration at departure and capture needs to be 0.1 gee or better. Can tolerate 0.01 to 0.001 gee during transits, but higher would be better. That’s what we need.
GW
The separate flow paths make the turbine feasible as long as it can be isolated when speeds exceed about Mach 3.3. But, I'm very skeptical of the "dual-mode ramjet-scramjet" that share inlets and nozzles. The internal inlet geometries are very incompatible, and scramjet requires a slightly divergent combustor to a strongly divergent nozzle, with no throat convergence. Ramjet on the other hand requires 65% throat convergence on area before you pass into the expansion bell, or it performs quite miserably, if at all. Since scramjet min takeover is about Mach 4, and turbine's max is Mach 3.3, you must have a ramjet.
And then there's the service ceiling effect with any airbreather. Thrust is made with airflow. If the air massflow is low due to low density, so is the thrust. As is the drag. But weight is independent of air density. Your thrust margin over drag is what you may use to accelerate or to climb, or a little of both. Multiplied by flight velocity, this is the "specific excess power" that relates directly to increases in potential or kinetic energy, or both: Ps = (T-D)*V/W. In the thin air above 100,000 feet, T-D is a number approaching zero, W is not. It won't climb, and it will have a max speed that is more driven by airframe L/D than by engine characteristics.
Why not just do a turbine-plus ramjet, and use a much cleaner (higher L/D) airframe to reach Mach 6 at about 100,000 feet? Then go rocket from there. Or, do simple rocket plus ramjet, and delete the heavy turbomachinery core at high altitude where the service ceiling effect is so bad?
GW
Add to the mix for SRM's: learn (finally) how to do a one O-ring segment joint design, and more importantly, why. Learn also how to do combustion stability analyses in SRM's. NASA's 5-segment version of the shuttle 4-segment design is not only worthless, it's actually quite dangerous. As is their 3 O-ring joint design.
GW
I know nothing about "ceramic foams". I don't think acetylene is a monopropellant, though. It never burns for me without air or oxygen. Explosive? YES!
I do know the 50-pound steel acetylene bottle on the welding cart in my shop holds around 5 lb of acetylene dissolved in around 5 lb of acetone. When full and for most of its useful "life" welding and cutting, bottle pressure is near 250 psig. That's awfully low compared to most rocket motor chamber pressures (500-1000 psig). Isp's are quite crappy when figured at 250 psig, no matter what you assume about the nozzle.
Once withdrawn from solution in the bottle, you NEVER use acetylene at more than about 10 psig. PERIOD. Your torch will blow up and kill you if you try. That makes the rocket chamber pressure mismatch even worse. There is no usable Isp at 10 psig chamber pressures in a rocket engine.
Check the inert weight fraction for 5 lb acetylene in 50 lb worth of bottle (90%). Now scale THAT up to 1000's of lb of acetylene on some rocket design. Not so attractive, is it?
The lesson here is things that look "theoretically" good from a chemistry standpoint are often quite worthless in practical terms. I think you can safely forget acetylene as a rocket fuel.
GW
The only geometrically similar items between an (afterburning) turbojet and a ramjet are the afterburner duct and the supersonic inlet components. Yet even these are suboptimal shapes if you try to switch roles. Neither of these has anything geometrically in common with a rocket engine.
That's why no combined-cycle engines have actually flown in all these decades. Getting around those incompatibilities is just too heavy, and penalizes the component engine performances too greatly. You are quite literally better off in terms of weight, performance, and complexity to just put the two engine types separately in the airframe.
Myself, for the drone (or piloted) hypersonic plane that folks call "SR-72", I'd use the ramjet in the fuselage of D-21 drone shape, and mount two small rocket engines in the aft ends of the wing fillets. The ramjet can take this plane as fast as Mach 5 or 6 in combat at high altitudes, and cruises for long range at Mach 3. Take off on simple rocket using large propellant mass in drop tanks, accelerate to Mach 2-ish for ramjet takeover, and climb at about Mach 2.5 to 3 into the thin air where ramjet is range-efficient. Then go do your thing. Glide landing with on-board rocket propellant residuals for go-around capability.
The D-21 shape has better L/D ratios BY FAR than any of these airframe-integrated scramjet things. Once you reach about Mach 3.5 at 80+ kfeet, you're pretty well beyond the reach of air defenses. What's the point of incurring all that trouble with an unready-to-apply scramjet technology and poor L/D cutting your range, just to fly Mach 8, when Mach 5-6 is good enough, and with a mature and reliable propulsion technology?
Do you really intend to fly, or do you want a technology development gravy train?
These things (press releases for big contractor SR-72 programs) are gravy-train "sops" for the two gigantic favorite oligopoly contractors. Government's favorite, that is. None of these wet dreams will ever really fly.
Somebody else, likely overseas first, is going to do exactly what I outlined above, and stun the world with it just like Sputnik 1.
GW
edit same day: further thoughts --- the other thing that comes to mind regarding "hypersonic" is aeroheating. If you do something other than a brief ablative or heat-sinked transient, you must solve the steady-state hypersonic thermal protection problem. This gets really severe above Mach 4. By the time you reach Mach 6, even your captured inlet air is flame-hot.
Go to Mach 8 or 10, and it just gets ridiculous to the point of serious stupidity: you are simply much better off leaving the atmosphere entirely if you really want to go that fast. Reentry is only 3 minutes long from LEO, after all.
From about Mach 4 or 5 on up, you must protect skins, leading edges, and nose tips, PLUS all your combustor, nozzle, and inlet surfaces inside. There is no such thing as "cooling air" above Mach 3. And the new ultra-high-temperature-ceramic nose tip and leading edge materials are simply not suitable for those other jobs.
Sometimes I wish people would forget about comparing things on Isp. Isp varies all over the map, vacuum or otherwise, depending upon what chamber pressure and expansion ratio you think you can use. And not everybody agrees about those things.
I wish people would compare characteristic velocity c* instead. That's the true function of thermochemistry and chamber pressure. Everything else, including Isp, depends upon that, and a whole lot more (expansion ratio, specific heat ratio (implicit in c*), nozzle kinetic energy efficiency, and chamber to ambient pressure ratio, not to mention separated-bell flows if your pressure ratio is too low. But, that requires folks to learn elementary interior ballistics. It's not hard. But it's not one simple calculation, either. But, it is repeatable and all can agree upon it.
Most of the high vacuum Isp's I see quoted depend upon large (but unspecified) expansion ratios, and do not include the nozzle kinetic energy efficiency. They also assume very high chamber pressures, that most manufacturers do not yet build to. For example, shuttle vac Isp is for 3000 psia chamber pressure. That's the only engine ever made that operates that high. Most everybody else is at or under 1000 psia, which makes a huge difference.
GW
I agree, landing the big, heavy Earth reentry capsule on the moon as part of a lander design is a stupid idea, based on what we learned from Apollo. That land-the-whole-ship idea was where NASA was with the Apollo design in 1963. It took two Saturn-5 launches per moon mission, plus cryogenic refueling from one 3rd stage to the other in Earth orbit, to make such a cluster landing possible back then.
The breakthrough idea came from outside NASA, and was resisted very strongly due to "not invented here" attitudes already in place at NASA that early. That idea was "lunar orbit rendezvous", which translates to "take only what you need to the surface, leave the rest in orbit about the moon". That got them down to one Saturn-5 per mission. By 1965 they had broken down and accepted lunar orbit rendezvous as the only way forward.
Strangely enough, using that plus LEO rendezvous and docking assembly could have gotten a moon mission down to 3 or 4 Saturn-1 launches, but they didn't need to do that , they had the Saturn-5 available.
BTW, the Saturns were not originally designed to be moon rockets. Von Braun was working for the Army before there was a NASA. Saturn-1 was a very large ICBM, and Saturn-5 started as concept for a "troopship rocket" to put 100+ men on the ground in Russia from the US in an ICBM's flight time. Von Braun knew these rocket could be, or could lead to, space launch rockets, just like a decade earlier with the V-2 for the Wehrmacht. Building ballistic missiles is how he got to play with space travel concepts, such as the "Nova" rocket we never built.
GW
The article was re-edited by "staff writers" from a 5-year old document, just to have something to publish. Those "staff writers" actually know very little about their subject, or they would have realized there is no shuttle anymore.
Credentials and experience actually do matter, after all. Out on the internet in general, I find that 99+% of what is posted is really ignorant BS. I'm happy to say that percentage is far lower on the forums.
GW
These Falcon-9 components have a stage inert weight fraction in the vicinity of 5%. With 9 engines, that leaves very little weight for the tankage, actually. So the stage is a very fragile item structurally.
What they learned from the ocean landing attempts so far is two-fold: (1) a need for redundant attitude control during powered descent in the atmosphere, and (2) a need for a hard surface upon which to land. That last comes from the very soft touchdowns in the water, that they've already pulled off while retro-thrusting. The tankage broke up and sank on impact when it flopped over on its side. Which also proves my point about structural fragility, and why parachute water landings and 5% inert fractions have no "solution space" available.
They added the grid fins to maintain stability during atmospheric descent, even in an engine-off condition. What they found is that engine restart was impossible if it tumbled. The pinpoint landing on a platform "solves" the flop-over breakup problem after touchdown, hopefully. If they can hit the target, it might work.
I'm impressed that they are already successfully using hypersonic/supersonic retro-propulsion, uncanted. Nobody else seems to have picked up that. I'm not sure yet how they're handling the entry heating, which should be in the vicinity of Mach 8 to 10 when it first hits air.
GW
Scramjet geometry is quite different from ramjet, in spite of some inlet components whose appearance is externally similar. The shock wave geometry is different at higher speeds and forces the external compression features to fall further ahead of the cowl. So, even the similar-looking part is really quite different.
In a scramet, there is no subsonic diffuser and duct with a terminal shock that moves from one diffuser area to another to match backpressures. Scramjets must have an isolator duct with a weak shock train to provide stability, or they tend to blow up. The kind of thing tested as a short dummy unit on the X-15 decades ago won't work: no isolator duct. This kind of inlet passage functioning and shape is nothing at all like the subsonic combustion ramjet inlet.
There is no convergence to a ramjet throat. In fact the basic combustor is slightly divergent, and the expansion nozzle is more strongly divergent. This kind of combustor and nozzle geometry is nothing at all like that of subsonic combustion ramjet.
The most difficult problem has been fuel injection. Injecting a fuel stream into a supersonic flow always causes a shock wave around the injected stream, and this is very susceptible to a full shockdown to subsonic flow. That event has all kinds of fatal outcomes, if it happens.
As I have said earlier, the min takeover speed for stable scramjet operation has long been known to be right at Mach 4. So far with hydrocarbon fuels (X-51) we have demonstrated stable burn for a couple of minutes at Mach 5, but without demonstrated airbreathing acceleration. With hydrogen (X-43), this is just under Mach 10, and only for a few seconds.
Inside these devices, the convective heating is just awful, even in the inlet ahead of the burn. To be successful as a propulsive device, this massive heat flow has to be recycled back into the combustion release. As near as I can tell, no one has done that yet, not to the significant degree that is required. The tests have been "heat sinkers", by and large.
GW
There's actually two design ranges for ramjet. Both are very simple devices, and usually fixed geometry, and extremely lightweight.
One is something I call "low speed range design", which feature a simple pitot/normal shock inlet, and a convergent-only nozzle that often runs unchoked. Nozzle area is pretty near 65% of combustor flow area. You have to size the pitot scoop correctly. These can light at subsonic speeds, start producing useful thrust (at greater than solid rocket Isp) around Mach 0.5 to 0.7, and peak out at about Mach 2.5 to 3, depending upon the drag of the airframe they are installed in. They are rarely used above flight speeds of Mach 2, and peak Isp occurs somewhere near Mach 1.1-or-so. That peak Isp with kerosene or gasoline fuels is near 900 sec if well designed.
The other is probably more like what folks would want to use for ideas typical of these forums: what I call "high-speed range design". These have external compression features like spikes or ramps just ahead of the capture cowling, features that generate shock waves that need to fall correctly upon other adjacent structures. That's part of proper design. Inlet capture area has to be properly matched to ramjet throat, for the specific characteristics of the inlet components.
The nozzle is always choked, has a throat near 65% of the combustor area, and a mild supersonic expansion ratio near about 1.3. Peak Isp is usually near 2.5 at near 1300 sec with hydrocarbon fuels running full rich, and max thrust about Mach 3-ish. Takeover speeds are usually between Mach 1.8 and 2.5, usually quite near 2.
There are no attached-shock solutions for the inlet compression features at speeds like Mach 1.4 to 1.5, which means there is no air ingestion there, and no thrust. Thrust decreases gradually with Mach as you speed up from the max thrust point, although airframe drag is increasing like a power function. Max speed depends more on airframe drag than anything else, and usually falls between Mach 4 and Mach 6, where the thrust and drag curves vs Mach cross.
You can replace chemical heating with nuclear heating by some means, which will remove the 65% throat area limit (a combustion stability item). Flameholding features can also be eliminated, which reduces internal losses. This resizes all the components, of course. The main problem is the same as is faced in rocket engines: once the heated gas ionizes significantly, you cannot convert further heat energy additions into velocity with a C-D nozzle. Recombination does not add velocity to the stream. This is true of chemical or nuclear systems.
GW