Spaceplane

SpaceNut · 2016-02-17 21:13:41

From an old friend:

Shaun wrote:

Skylon is gradually moving forward.
The European Space Agency (ESA) has granted approval for the Skylon jet plane that will be able to take us anywhere in the world in four hours to be equipped with the revolutionary Sabre (Synergistic Air-Breathing Rocket Engine) technology.
I still LOVE the look of this beast!
What's not to like about a single-stage-to-orbit (SSTO) vehicle which might take off from a level runway, move 15 tonnes to LEO, and then land on a level runway again?

kbd512 · 2016-02-18 00:06:57

RobertDyck wrote:

Looking at propellant for a commercial spaceplane. I suggested using jet fuel for take-off/landing, and acceleration to speed required to start the next engine. Then air as propellant for a nuclear jet engine to gain speed and altitude. Then steam distilled water as propellant for a nuclear thermal rocket for final push into orbit. Nuclear engines require respect, but as long as the reactor remains sealed, vehicle processing is simple. And I suggested Americium-242m as the nuclear fuel. Americium-241 is the isotope in smoke detectors, so it's safe to handle as well. Plutonium is extremely poisonous, Uranium metal is toxic, Uranium oxide less so but must not be ingested, however Americium is in smoke detectors sold for use in homes. Carried propellant is either normal jet fuel, or water. That just leaves RCS thrusters for manoeuvring in space. One option was the "green propellant" being developed by NASA. AF-M315E is monopropellant with Isp=257 seconds, which really doesn't compare well to N2O4/MMH bipropellant. What could we use as an easy to handle propellant?

Am-2421 or Am-242m is safe to handle? Where are you getting this from? Am-242m is more radioactive than P-239, with a half life of 141 years. Smoke detectors are "safe", acknowledging that "safety" is a human brain construct that has no applicability to reality beyond the ears of the person who thought it so, because there's so little Am-241 inside. The radioactive material is also not normally anywhere near the people occupying the building it's installed in. I promise you that anyone who mistakenly believes that any significant quantity of highly radioactive material can be handled without precautions to prevent ingestion is not long for this world.

As stated elsewhere in other threads, I believe Am-242m is our best fissile material for development of extremely compact and light fission reactors for space power applications due to the reduction in fissile inventory required to achieve criticality. All fission reactors, fissile materials that have undergone fission, and fission products require heavy metals to attenuate gamma rays and hydrogenated materials to absorb neutrons. That's a lot to ask of a SSTO spacecraft that would require a GW class reactor to produce the heat required.

To paraphrase the designers of the original nuclear powered aircraft:

"That the design was impractical, if not foolish, mattered less than the technology development opportunity afforded by the project."

RobertDyck wrote:

What about LOX/propane? It has Isp=361.9 seconds in vacuum with 100:1 nozzle-to-throat area ratio, fuel density 582 kg/m^3, bulk density (fuel + LOX) 905 kg/m^3 at Near Boiling Point (231°K). At 100°K which is just 10° above freezing, fuel density is 782 kg/m^3, bulk density 1014 kg/m^3. Combustion temperature 3734°K. As comparison, methane Isp=368.3 seconds in vacuum, fuel density 423 kg/m^3 at NBP, bulk density 801 kg/m^3, combustion temperature 3589°K. Propane is liquid at room temperature at very mild pressure, which is why it's used for backyard BBQ. Pressure for a propane tank at +25°C is just 10 bar (1 MPa). Of course the tank requires a safety margin above that, but we're talking about the pressure of a backyard propane tank at room temperature. That's easy to handle.

Didn't Lockheed already solution this problem. The use of composite tanks that weren't ready for prime time in the 1990's was what ultimately killed Venture Star, but there was nothing technically infeasible that prevented the project from going forward. With today's no-leak composite tanks and high strength aerogels for insulation, I see no reason why we shouldn't simply use the aerospike J-2 derivatives developed for the project.

RobertDyck wrote:

Shuttle RCS thrusters and Apollo service module thruster quads used N2O4/MMH. Apollo used R-4D, Isp=312 seconds, expansion ratio 164:1. So LOX/propane has a higher Isp. The issue is keeping LOX cold without boil-off. Is that reasonable?

Why not use integrated vehicle fluids- LOX/LH2 RCS?

GW Johnson · 2016-02-18 09:44:47

As I understand from reading about Skylon, its max speed on airbreathing propulsion is about Mach 5. Then it goes rocket with stored oxygen. And it leaves the atmosphere pretty quickly upon that transition. Speeds like that are on the hairy edge of cutting the wings off with the shock shed from the tip of the engine compression spikes. It's called shock impingement heating, and it is very fatal to all known structures and materials by about Mach 6-ish speeds. It is why no entry spacecraft has ever had adjacent nacelles of any kind, not in all these decades.

To me it looks like Skylon might successfully survive the ascent. But, I would be very interested to know how they plan to survive reentry on return. Those compression spikes are still there, shedding strong shocks that hit right on the adjacent wing leading edges. Leading edge heating during entry is tough enough, without any shocks impinging at all.

There's an interval there from Mach 25 down to about Mach 5 that is incredibly unsurvivable due to that shock impingement heating. I'm not sure that even a one-shot / throwaway silica-phenolic heavyweight heat shield piece would withstand the insult. This is the phenomenon that nearly cut the tail off of one X-15 carrying a scramjet engine test article on its ventral fin stub at a peak Mach 6.67. They had even put a white ceramic coating on the bird, and it didn't help.

GW

SpaceNut · 2017-05-26 16:13:01

Forget supersonic: Hypersonic is the US military's new speed

Boeing Co.’s XS-1 (Experimental Spaceplane), which the company dubs “Phantom Express,” got a green light this week by the Defense Advanced Research Projects Agency, or DARPA. The XS-1 is designed to quickly lift satellites as heavy as 3,000 pounds into orbit for $5 million or less, launching from the ground, deploying a small upper-stage module, and then landing like a traditional airplane—the key to reuse and lower operating expense. DARPA also has a separate program aimed at launching 100-pound satellites for less than $1 million per launch, using conventional aircraft. The Phantom Express will be powered with an Aerojet Rocketdyne Holdings Inc. AR-22 engine, a newer version of the main engine trio that served on NASA’s Space Shuttle. Boeing will design and build the aircraft through 2019, including 10 engine ground firings over 10 days, followed by 12-15 flight tests in 2020.

Seems we have some real old topics coming up again....

SpaceNut · 2017-07-13 19:14:07

Shaun wrote:

Hypersonic flight (Mach 5 and above) in the atmosphere is fraught with difficulties, as we know; not least of which is the heat generated at the leading edges of the aircraft. Anywhere from 2000-3000°C.
Apparently one of the best materials we've had to resist the severe oxidation and ablation that occurs at that speed is zirconium carbide, but it's really not good enough.
According to THIS ARTICLE, scientists and engineers think they may have solved the problem:
Researchers at The University of Manchester, in collaboration with Central South University (CSU), China, have created an innovative ceramic coating that could revolutionize hypersonic travel - for defense, space and even air travel purposes.
... the University of Manchester's and the Royce Institute, in collaboration with the Central South University of China, have designed and fabricated a new carbide coating vastly superior in resisting temperatures - perhaps even up to 3,000 C - taking air travel that much closer to achieving Mach five and beyond.
... the carbide coating developed by the teams is proving 12 times better than the optimal conventional UHTC [Ultra-High-Temperature-Ceramic], Zirconium carbide - an extremely hard refractory ceramic material commercially used in cutting tools.
Someone should tell Reaction Engines about this! If they haven't already.

GW Johnson · 2017-07-13 19:51:45

Well, there's a bit more to heat protection in hypersonic flight than just a ceramic coating on the metal. If you just do that, the metal is almost as hot as the coating, precisely because the coating is a dense, high thermal conductivity material, and it is very thin.

You can do that for actively-cooled turbine blades at no more than 1200 C, in gas somewhat hotter than that. If the blades are not cooled inside, then the gas and material temperatures are just about the same, and 1200 C is just about the max. Highly stressed designs must run cooler than that. The adjacent air stream is that hot at only Mach 5.3 in the cold stratosphere.

"2000-3000 C" is way above the meltpoint of all known feasible structural metals, which casts great doubt as to whether the press release is a real article or just publicity hype. I really suspect there is no real airframe design approach with this ceramic yet.

If you can find an acceptable surface treatment or coating that is radiationally "black" (emissivity > 0.80), you have the option of passively cooling a superalloy structure by radiation to the environment to about 870 C. At only Mach 5, this even works down in the stratosphere around 50,000 feet or so. The speed limit is much slower in the warmer air lower down.

The more loads the hot structure must bear, the cooler it must equilibriate in order to be strong enough, and the more severe this speed limit due to heat transfer equilibrium will be. If the structure need not bear loads, then there are a couple of 300-series stainless steels that will serve. If the structure bears loads, then something like a Haynes 188 or 230 is required.

Speed limits are higher in the thinner air above 100,000 feet, in spite of the fact that the air there is warmer than in the stratosphere. This is because film coefficients and convective heat loads are more or less dominated by low air density.

Sorry, it's complicated and it's difficult to do. But above 100,000 feet, steady flight speeds around Mach 6 or 7 might be feasible. Much beyond about 150,000 feet, lifting flight starts getting problematical, just because the air and its forces have reduced so low, while weight is the same as at sea level.

To bear loads as a lightweight structure, the material must stay very much cooler. That requires a low-density insulation, or an ablative, to limit conductive heat transfer into the interior. You cannot zero that heat conduction! You must do something with that heat for steady flight. What works in reentry is transitory: they literally heat sink their way through a peak-heating event that is only 2-3 minutes long.

GW

Last edited by GW Johnson (2017-07-13 20:08:33)

GW Johnson · 2017-07-18 11:30:24

Spacenut inserted a press release article for Lockheed-Martin's SR-72 that USAF has just funded into "unmanned probes", as "running on empty - NASA launches on a wing and a prayer". It really belongs somewhere else, maybe here. I posted there a couple of things regarding hypersonic aircraft in response to it. It's pretty clear that, as publicized, this is a gravy-train project funded by USAF for Lockheed-Martin. It won't go anywhere, not as-proposed.

Boeing got a semi-gravy train funding from DARPA for its XS-1. This is a hypersonic rocket glider that could actually be built and flown. Given a big enough rocket booster, it could be a spaceplane, but not as a single stage, no. It is a reprise of the old X-20 Dyna-Soar that USAF funded at Boeing in the late 1950's. That one was killed in 1963 with the first 3 examples near the end of the production line.

My how history repeats.

GW

Last edited by GW Johnson (2017-07-18 11:38:48)

Oldfart1939 · 2017-07-18 11:42:37

History may not exactly repeat itself, but does rhyme.

RGClark · 2017-12-28 11:39:40

GW Johnson wrote:

Spacenut inserted a press release article for Lockheed-Martin's SR-72 that USAF has just funded into "unmanned probes", as "running on empty - NASA launches on a wing and a prayer". It really belongs somewhere else, maybe here. I posted there a couple of things regarding hypersonic aircraft in response to it. It's pretty clear that, as publicized, this is a gravy-train project funded by USAF for Lockheed-Martin. It won't go anywhere, not as-proposed.
Boeing got a semi-gravy train funding from DARPA for its XS-1. This is a hypersonic rocket glider that could actually be built and flown. Given a big enough rocket booster, it could be a spaceplane, but not as a single stage, no. It is a reprise of the old X-20 Dyna-Soar that USAF funded at Boeing in the late 1950's. That one was killed in 1963 with the first 3 examples near the end of the production line.
My how history repeats.
GW

GW, do you have software that can do simulations of reentry, to supersonic, to subsonic flight for different wing airfoils?

I was thinking of ways to do guidance in near vacuum for small amateur built stages. Early in the space program spin stabilization was used for stages at high altitude where fins were ineffective. But this meant you had little control of the final orbital parameters. I was thinking of ways you could get this fine control at low cost and light weight for amateur built stages.

Fins are effective at low altitude even with their relatively low size compared to the size of a rocket due to the dense atmosphere. I thought then if you made the aerodynamic surface a hundred times larger for air density a hundred times less, this should again be effective. This space shuttle wings cause significant slowdown even at reentry speeds for example because of their large size.

So how to add large aerodynamic structures to a small stage? You could use wings but this would be high drag and high disturbance at low altitude that would have to be counteracted especially for solid rocket stages that leave the launch pad at high speed.

To avoid this I thought of the curved cylindrical surface of the rocket stage that could open up like a clamshell:

In our scenario though the stage itself would not open up revealing the interior but the extra aerodynamic surface, call them clamshell wings, would be attached to the exterior of the stage. They would be closed up around the stage during low altitude flight, and opened at high altitude.

It occurred to me then that this also might be able to be used for reentry. If you can make this extra surface be lightweight then you would get low wing loading. The importance of low wing loading for reentry for spaceplanes is discussed here:

Wings in space.
by James C. McLane III
Monday, July 11, 2011
http://www.thespacereview.com/article/1880/1

At the end of the article there is this passage:

Wing loading (the vehicle’s weight divided by its wing surface area) is a prime parameter affecting flight. The antique aluminum Douglas DC-3 airliner had a big wing with a low loading of about 25 psf (pounds per square foot of wing surface). At the other end of the spectrum, the Space Shuttle orbiter has a high wing loading of about 120 psf. This loading, combined with an inefficient delta-shaped wing, makes the orbiter glide like a brick. A little Cessna 152 private plane features a wing loading of about 11 psf and modern gliders operate down around 7 psf. A space plane with huge lifting surfaces and a very low wing loading might not require any external thermal insulation at all. Building a space plane with a wing loading of, say, 10 psf should not be an impossible proposition. Perhaps some day it will be done.

{emphasis added}

Moreover, because of their curved shape they should be even more effective at slowing down the descent during reentry, like a parachute.

I estimated the wing loading using this clamshell wing idea for the new Falcon 9 FT first stage, assuming they added a proportionally small amount to the weight. I used the specifications here:

Falcon 9 FT (Falcon 9 v1.2).
http://spaceflight101.com/spacerockets/falcon-9-ft

The dimensions given there are listed as 42.6 meters long and 3.66 meters in diameter, at a dry mass of 22,200 kg.

Regarding the stage horizontally, you would have to put the swing points along the sides, rather than at the top, so that the clamshell wing on each side could open without blocking the opening of the clamshell wing on the other side. This means the wing area would be half that of the full surface area. So the surface area is (1/2)*Pi*3.66*42.6 = 244.9 m^2, 244.9*3.28^2 = 2634.86 ft^2.

The dry mass is 22,200 kg, 22,200*2.2 = 48,840 lbs. So the wing loading is 48,840/ 2634.86 = 18.5 pounds per square foot(psf). This is not 10 psf, but it is significantly better than the shuttle, and with the reduction in descent due to the curved surface this might still be enough to require minimal thermal shielding.

Also, we might be able to get additional wing area by putting clamshell wings on the upper surface, though not the same size as the lower ones so that all can open fully.

For attitude control we allow the swing points to be moved up or down.

Bob Clark

Last edited by RGClark (2017-12-28 11:53:26)

GW Johnson · 2017-12-28 21:12:39

Bob:

I'm afraid all I have relevant to high altitude hypersonics is the old scale-height constant-coefficient method developed for warhead entry in the 1950's, and I have that as a spreadsheet. It's an oversimplified 2-D model for nonlifting ballistic entry only. Anything else would be nothing but pencil-and-paper straight out of a book.

Low "wing loading" is less applicable for entry than is "ballistic coefficient", which is mass divided by coefficient x blockage area. It's like wing loading, but the area is different.

What happens at low ballistic coefficient, all else equal, is much higher deceleration gees, with the deceleration occurring higher up where the air is thinner. Heat transfer coefficients vary (crudely) as density ^ 0.8, so convective heating is lower in thinner air. If the exposed structure can cool by radiation, that balance occurs at a lower surface temperature in the thinner air.

If cooling by radiation (or active cooling) cannot be had, the structure essentially eqilibriates its temperature with the effective temperature of the air. Ignoring the difference between recovery and total temperatures (rather small in turbulent flow), then you can estimate the equilibrium temperature as the total temperature, up to ~ Mach 6.

For Mach 6 at 130,000 feet in a standard atmosphere model, this is 3680.2 R = 2044.5 K. So you can see the material problem. Even the ultra high temperature ceramics won't help, as they will get that hot, and you have to hang onto them somehow. But, with what?

For higher speeds, a rough estimate of effective "air" temperature (plasma, really) is T, K ~ V, m/s. At 130,000 feet and Mach 15, velocity is roughly 15,570 ft/s = 4746 m/s. The plasma would heat you as if its effective temperature were ~ 4800 K. This is far beyond any possible ultra ceramic.

So, in both cases, some sort of cooling is an utter necessity. This can be radiation of energy (to 300 K Earth temperatures, not the 4 K of deep space), some sort of active cooling (a real problem with fluid choices, pressures required to stop phase change, and coking from thermal breakdown), or else ablation. Or some combination of these. As far as I know, there are no other physics to invoke here.

They solved this on the shuttle with fragile low density ceramic tiles and ceramic blankets on the lateral and leeward surfaces that equilibriated by radiation, and with ablative carbon-carbon on the leading edges and nose tip that also re-radiated as well as burnt slowly away. Nose tip and leading edges would re-fly a tiny handful of times. The tiles lasted longer, except for chronic impact damage and bondline failures.

On the space capsules, they used one-or-another ablative composite, beginning with things like silica or carbon phenolic materials. Some today claim Apollo's main windward heat shield was Avcoat, but I have a hard time believing that. I always thought that Avcoat was the heat shield ablative on the lateral sides. On Mercury and Gemini, the lateral sides were metal re-radiating to space.

I think you might get away with a metal winged craft given ablative leading edges and nose tip, if you can keep the wing loading / ballistic coefficient low enough, and you can still stand ~ 10+ gees deceleration with a low ballistic coefficient. But I doubt you could fly the thing twice. I looked at a folding-wing design that entered dead broadside with a low-density ceramic heat shield. It looked feasible for reusable flight, but such a thing requires feasibility testing to confirm even that. I folded the wings to keep them light, but without ripping them off while broadside to the flow.

Ascent heating is just not at all that bad, because most rockets are only doing around Mach 2 at 100,000 feet (~410 K total temperature), and maybe still under Mach 5 at 150,000 feet (~1600 K). You can get away with folding metal structures for aerodynamic control at such conditions, especially if they can cool by radiation.

I'm not really sure it's a solution, though, because it would be inherently heavy as large surfaces. Why not just do attitude control jets? That's what they did on the old "Scout" launcher, and it worked great for over 3 decades. Monopropellant decomp thrusters, deadband control scheme. All on a 4 stage solid. Low tech could be pressurized superheated water as a steam rocket thruster. One small heavy thermos bottle, plus some sheet steel thruster items.

GW

Last edited by GW Johnson (2017-12-28 21:31:57)

RGClark · 2017-12-29 17:11:23

Thanks for that. BTW, thinking about the clamshell wings a little more, and testing with a paper cylinder, I think we can get the full surface area, not just half, by making the swing points be on the top surface. That is, you can get the two wings to open fully without them restricting each others movement.

So the wing-loading would be in the 10 psf range mentioned in the "Wings in Space" article. Also, that calculation was for the first stage which won't even get to orbital speed and will experience much reduced reentry heating anyway. So it won't even need the psf to be so low.

But interestingly, the wing-loading for the upper stage, which of course does reach orbital speed, will be even better. Again use the specs here:

Falcon 9 FT (Falcon 9 v1.2).
http://spaceflight101.com/spacerockets/falcon-9-ft

The length is given as 12.6 meters and the diameter as 3.66 meters, and the dry mass as 4,000 kg, 8,800 lbs. So the surface area is PI*3.66*12.6 = 144.88 m^2 = 144.88*3.28^2 = 1,558.65 ft^2. And the wing-loading is 8,800/1,558.65 = 5.64 psf. I wonder what would be the heating for wing-loading that low.

BTW, I mentioned your Mars Society conference presentation on lightweight thermal protection again in the latest post to my blog, exoscientist.blogspot.com. The density you estimated for your material of 0.03 specific gravity, is only a third that of the famous shuttle aerogel underside tiles, which were legendary for their high insulation at low weight.

You mentioned in the video your material had the lightness of styrofoam, which would indeed put it in the 0.03 range. Could you weigh a sample you have on hand to confirm that? If confirmed that would be a radical improvement over current materials.

Bob Clark

GW Johnson · 2017-12-29 18:08:19

Hi Bob:

Sorry, I last made that stuff in 1984 at what was then Tracor Aerospace in Austin, Texas. I retained no samples when I left, only stole some photographs and some data documents. I did confirm with the manufacturer before the presentation at the convention that the paste and cement coat materials are still available, as is the fire curtain cloth I used as a reinforcement.

I'm not at all sure, but I believe the tooling geometry and processing procedure that I used forced the moisture out as steam, wormholing through the laid-up composite, and reducing its density below what is normally obtained heat-curing a pipe insulation paste. The paste I used was Cotronics Corp 360M, with a finished surface coating of Cotronic 901 cement as a sealing surface paint. The reinforcing fabric was Nextel 312 fire curtain cloth. These are all alumino-silicate materials.

When I checked on availability, these materials no longer show in the Cotronics catalog, but they are still available. I called them to verify availability, and Cotronics thought I could blacken the 901 cement with carbon black to achieve IR emissivity ~ 80%. Nextel 312 woven fabric with yarns about 0.030 inch diameter is still available from 3M, but is considered obsolete as well. I think the product name is AF-30.

I'd have to buy the stuff and fabricate tooling to recreate the material.

GW

RGClark · 2017-12-30 18:03:17

Thanks for that info. As you described its development in the video, this would certainly be patentable. Even if you don't have samples you could still patent the material, describing the process to make it.

In addition to the use as reentry thermal protection. It would also have use as combustion chamber insulation. This would have widespread applications since in this post I discussed the case of the Star 48B for which the insulation counts as a significant proportion of the dry mass of the stage.

In the video you said it weighed less than PICA-X that SpaceX uses but more than the space shuttle underside tiles. But actually if your material really is 0.03 specific gravity then it is 1/3rd the weight of the space shuttle tiles. Perhaps because you didn't realize how much lighter it would be than currently existing tech is why you didn't see the urgency in confirming its properties and patenting it.

Bob Clark

Last edited by RGClark (2017-12-31 03:33:22)

GW Johnson · 2017-12-31 14:00:35

Bob:

I did that work as an expedient to experimentally test a hot gas source and infrared emitter. It was a shoestring IR&D project aimed at feasibility of an alternative to an aircraft flare. I tried it on a guess and it worked. No resources were left to do anything else.

By the way, the decoy was also feasible. Although nobody has ever built one since. I built a thing with the IR power and color temperature of an F-100 Super Sabre tailpipe at full military power. It would blister your face with heat 20 feet away. It was only about the size of a coffee cup. It would inherently fool a two-color seeker with its spectrum. That the decoy worked at all was also a lucky guess, but it did.

GW

RGClark · 2018-01-04 12:54:12

What was the chamber pressure when you tested it as insulation in ramjet combustion chambers?

Bob Clark

GW Johnson · 2018-01-05 13:08:57

It was rigged as a subsonic ram combustor, immersed in a stream from about 0.5 Mach to about 0.9 Mach. Max engine static pressure in stable burn operation was about 10 or 15 psig. Most of the time it was less. The light-up and flameout transient times were on the order of 50-100 msec ("blink of an eye" stuff, not measured).

When I drove it into rich blowout instability, the size of the pressure oscillation was very roughly 1 atm above steady state, at a audio frequency somewhere between 50 and 100 Hz. That's a peak pressure near 2 atm, on a cycle time about 0.1 sec long. From "low" to near 2 atm in about .005 sec is a pressure rate (ramp rate) in the gross vicinity of 400 atm/sec.

That's way too fast for the porous material to fill and drain, so it was actually demonstrated capable of holding pressure difference 2 atm while hot. I showed a tad of surface melting, so the very surface of it was 3200-3300 F, while the side facing the steel shell was nearer 215 F. Being a nominal quarter inch thick, that's in the vicinity of 12,000 F/inch thermal gradient at the test conditions. That is how I know it really was low density.

I never weighed the silly liner (and I should have), but it felt to the hand as if it were carved from one of the denser grades of commercial styrofoam. It was alternating layers of paste and fire curtain cloth wound onto a mandrel, and inserted into the combustor shell. I oven-cured it at 215-220 F in an oven, above the steam point for the water that cooks out of it during cure. The path for the steam to escape was obstructed by the shell and the mandrel, which is why I think steam wormholing is a part of the low density feature.

The paste used as pipe insulation is cured out in the open, with a heat gun, so it stays below the steam point, and that is much denser than styrofoam. Its thermal conductivity is some 5 to 10 times higher than what was evident in my tests (something in the vicinity of 0.01 to 0.02 BTU/Hr-ft-F from a long-after-the-fact calculation).

Supposedly, the pipe insulation is good for about 100-200 psi normal pressure, but that's the higher-density form. The low density form has to be weaker because of the higher porosity. I never tested that. It held up in rich blowout instability, and that was "good enough" for my decoy testing purposes.

Immersed in a 190 F airstream, steady-state shell temperature was just enough to boil spit (about 215 F). That's with 3300-3400 F hot gas inside the chamber. The shell matched the dimensions of 2 inch schedule 40 iron pipe. I had the nozzle sized to choke down combustor Mach number to under 0.100 or thereabouts (never actually verified in any way, I just found a reduced nozzle size that let it light and burn).

Somewhere I've got a Seider-and-Tate model of the combustor film coefficient, and a similar one for very turbulent flow along the outside of the 300-series stainless shell. The steady state heat flow balanced at k ~ 0.01-.02 BTU/hr-ft-F, ignoring radiation effects, at the Mach 0.9 test condition.

We never took digital data. There were a couple of thermocouples fed to a data logger, and the pressure taps were connected to a water manometer bank, and read manually. Very simple, very unsophisticated. None of those records survive from that time.

And that's about all I know.

GW

Last edited by GW Johnson (2018-01-05 13:10:13)

elderflower · 2018-01-06 05:56:44

Was your fire curtain cloth asbestos based back then, GW?

GW Johnson · 2018-01-06 23:50:05

Hi Elderflower:

No, it was alumino-silicate. No asbestos. Ordinary commercial/military fire curtain cloth. Mine was state of the art then, rather obsolete now. It was 3M's Nextel 312 ceramic fiber woven as their AF-30 fabric product.

A pretty good description of what I did is given in posts 62, 64, and 66 just above, taken together. I was bloody lucky, twice over, doing that project. Not only did the composite insulation work fine (paste alone shattered under the rich blowout instability violence), the same basic alumino-silicate materials also behaved as a very non-gray emitter, making the decoy work unexpectedly well.

I used the fire curtain cloth and the ceramic cement-as-a-coating as my infrared emitter, sort of like a Coleman lantern mantle, heated by the gas from the ram combustor. Because it was sharply long-wave-selective in its spectral emissivity I had a 3000 F object with its spectral peak looking like only 1200 F. Yet the coffee cup-sized decoy radiated the same total infrared power (in any wavelength band you care to name) as a jet engine tailpipe about a cubit in diameter and a cubit long.

Because of the distorted spectrum, this thing would fool even a two-color infrared seeker, something a mag flare just inherently cannot do.

GW

elderflower · 2018-01-10 04:07:11

Thanks, GW. I was just wondering. Asbestos was still in use in the early days of my engineering career. For use on Mars it might still be an option for ascent engine or heat shield components - nobody is likely to breathe in the dust. So far I seem to have escaped its dire effects.

GW Johnson · 2018-01-10 13:25:55

Lots of us were exposed to asbestos, and lots of them never knew it.

Asbestos incorporated into an item like a linoleum or a paint or a wallboard/ceiling tile presented little or no danger, as long as you left it alone. Most of the remediation efforts removing things like that have made airborne fiber counts worse, not better. Everybody in the remediation business knows that, and will tell you that privately.

But the "official government regulatory line" is that such asbestos-containing products must be removed. That particular emperor is as naked as a peeled egg. When remediation actually makes things worse, then the remediation regulations are patently stupid. And they are. Similar for lead paints that aren't peeling.

The shipyard and refinery workers spraying asbestos fiber or installing raw mats of asbestos fiber had very lethal exposures. In those workplaces, often you could not see 50 feet for the dense cloud of fibers in the air. Those are the people who died/are dying in droves from asbestosis and asbestos-induced lung cancer.

The other big exposure was asbestos-organic brake shoes and pads. Brake dust used to be full of asbestos fibers. These days, there are organic pads and shoes with some fibers other than asbestos: organic, metal, and ceramic. That danger is now gone. However, since the air in automotive repair shops was never so full of fibers as the shipyard and refinery workers, those workers never died in droves. Cases of asbestosis or asbestos-induced lung cancer were actually rate among those workers. I have done my own brake work for 4+ decades now.

In the defense industry, the best solid rocket motor liners were asbestos fiber-reinforced rubbers. Those have been replaced by kevlar fiber reinforced rubbers. Not quite as good, but no asbestos fibers get spread by liner ablation into the rocket plume. Again, that danger was minimal, for both test personnel, and for rubber-formulating personnel, because the air was never clouded with airborne fibers. There might have been a few cases, but no droves of folks affected. I was a rocket motor tester for 2 decades, among other things.

That kind of fiber-reinforced ablative rubber liner was and is very common in solid rocket motor cases (not the nozzles). In fact, it's about the only thing that makes any sense to do. Liquid rockets have never really used ablative chamber liners, although such would be possibly feasible for one-shot items. Not reusable items. The nozzles would require fiber-reinforced phenolic plastics (carbon, and silica fibers might work, and perhaps glass, but not as well. Asbestos worked fairly well, but silica worked better), with graphite nozzle inserts.

GW

Last edited by GW Johnson (2018-01-10 13:32:06)

RGClark · 2018-01-13 10:21:14

GW, about reusable orbital stages, SSTO's or upper stages, I was surprised in most discussions of them it is assumed they are able to reenter to achieve an only 100 m/s terminal velocity. This is true whether horizontal or vertical propulsive landing is being discussed. See for example here:

http://yarchive.net/space/launchers/hor … nding.html

As the example of the space shuttle shows, this is not so surprising for winged, horizontal landing. But the thing is this is assumed even for vertical propulsive landing. This is surprising because this means for a cylindrical rocket stage, without wings, reentering broadside or a conical stage entering base forward, they can cancel out almost all of the ca. 7,800 m/s orbital velocity to get down to only 100 m/s terminal velocity from aerodynamic drag alone.

Actually, I haven't seen the argument for this. Anyone have a reference for the idea you can get down to 100 m/s terminal velocity even without wings?

But if so why all the big debate over vertical vs. horizontal landing? For either method nearly all the 7.8 km/s orbital velocity will already be cancelled out before either landing method comes into play, and even then they only have to account for a measly 100 m/s.

That is surprising though that without even needing wings you can cancel out nearly all of orbital velocity to get down to terminal velocity. Given that, on Earth reentry for an orbital reusable isn't even particularly hard. So why all the hullabaloo about how difficult it is to get full reusability because you also need to return the orbital stage from orbital velocity?

In fact orbital stage reusability might even be easier than first stage reusability:

For the first stage, you need a reentry burn to cancel out forward velocity, then a boostback burn to return to the launch site then you need the final landing burn. This results in quite a bit of lost payload because of the propellant that needs to be kept on reserve.
But for the orbital stage you just let Earth rotate beneath you until the launch site is back under you before you initiate reentry, so you don't need the boostback burn. Then aerodynamic drag cancels out nearly all of orbital velocity even without wings so you don't need the reentry burn. Then you only need the landing burn and this is for only ca. 100 m/s.

Yes, I know for the orbital case you need more thermal protection than for the first stage but this is not particularly heavy anyway. Even for the space shuttle, this was only ca. 8% of the landed weight, and SpaceX's PICA-X weighs half that to only 4%. And if GW's thermal ceramic really is as lightweight as he estimates the thermal protection would be a minor portion of the vehicle weight.

Then when you consider the either wing weight or propellant weight needed for the final 100 m/s of the landing would be 5% or less, the total lost payload for the orbital stage would be less than 10%. Compare this to the 30% to 40% lost payload for the two stage Falcon 9 or BFR.

Taking these facts into account a reusable SSTO may actually be more economical than a reusable two stage vehicle.

Bob Clark

Last edited by RGClark (2018-01-14 10:32:47)

GW Johnson · 2018-01-14 01:17:15

There's nothing magic about "100 m/s" being terminal velocity, although numbers like that are sort-of typical for objects of ordinary shape and object density, low in Earth's atmosphere. Terminal velocity is an equilibrium thing, which takes significant time and space to approach. That's why folks get into such trouble with that concept on Mars, where the air is so thin.

You have great difficulty approaching an equilibrium before you run out of time and space on Mars, because you go transonic so deep down close to the surface. And on Mars, terminal velocities are somewhere between high subsonic and low supersonic in that thin air, for objects of ordinary density and shape.

The equilibrium is drag = weight. Weight is mass x local gravity acceleration, with gravity acceleration nearly constant over rather wide ranges of altitude. Drag is dynamic pressure x drag coefficient x the reference area associated with that drag coefficient. Dynamic pressure is 0.5 * density of the air x object velocity squared, or 0.5 * air specific heat ratio * air pressure * object Mach number squared. Up high, where density or pressure are low, velocity or Mach will have to be higher to create the same value of dynamic pressure required for force equilibrium.

An object's weight and its aerodynamic reference area are sort-of functions of its object density. This effect of density makes the denser object smaller for otherwise the same mass. The bigger the dimension, the bigger any areas and volumes are.

Here on Earth very low in the atmosphere, a human body falling dead broadside to the air has maximum coefficient x area for max drag. Terminal velocity (last few thousand feet before impact) is usually around 120 mph. Slow, but still more than fatal. That same person diving headfirst with forward-stretched arms like a diver has a lower drag, and will equilibriate closer to 250 mph. Feet first with toes pointed, arms tucked to body and hands over face, drag is lowest, and equilibrium obtains at about 300 mph. The density of a human body is very close to that of water.

Up in the thin air, equilibrium speed is just higher. When Joe Kittinger made his 100,000-foot balloon ascents and parachute descents about 1960, his free-fall speed up there was just about Mach 1, whereas below 20,000 feet it was under 300 mph, and he could survive the chute opening shock. Chute touchdown speed (24 foot dia personnel chute) was just about 20 mph, maybe 25 mph with the heavy pressure suit. Survivable, although sometimes painful.

What is true of human bodies is true of other falling objects. Only the numerical details vary. An empty rocket stage falling dead broadside falls slower than one falling end-on. Of course, the way such stages are usually constructed, wind pressures applied dead broadside are far more likely to crush the object to structural failure.

You have to do something to slow down further to survive the final landing. On Earth, parachutes work quite well for paratroopers, converting a feet-first unopened-chute descent at 250+ mph into a survivable 20 mph touchdown with the deployed chute. The Soviet military air-dropped far larger and more dense objects with parachutes: army tanks. In that case, chute descent velocity was not survivable even by unmanned equipment at something like 100 mph with the chute. They fired last-second retrorockets to bring the touchdown velocity under 30 mph.

One way of book-keeping this information is "ballistic coefficient" = mass / (coefficient x area). Only gravity acceleration and dynamic pressure are missing ingredients for the equilibrium equation, and those can be supplied. Higher ballistic coefficient objects equilibriate at far higher velocities (or Mach numbers), even in the dense air down low. And certainly faster higher up in the thin air.

Again, the key is equilibrium. Easy to achieve in a dense atmosphere like Earth's. Hard to achieve in a very thin atmosphere like Mars's.

GW

Last edited by GW Johnson (2018-01-14 01:31:49)

RGClark · 2018-01-14 12:04:06

RGClark wrote:

Thanks for that. BTW, thinking about the clamshell wings a little more, and testing with a paper cylinder, I think we can get the full surface area, not just half, by making the swing points be on the top surface. That is, you can get the two wings to open fully without them restricting each others movement.
So the wing-loading would be in the 10 psf range mentioned in the "Wings in Space" article. Also, that calculation was for the first stage which won't even get to orbital speed and will experience much reduced reentry heating anyway. So it won't even need the psf to be so low.
But interestingly, the wing-loading for the upper stage, which of course does reach orbital speed, will be even better. Again use the specs here:
Falcon 9 FT (Falcon 9 v1.2).
http://spaceflight101.com/spacerockets/falcon-9-ft
The length is given as 12.6 meters and the diameter as 3.66 meters, and the dry mass as 4,000 kg, 8,800 lbs. So the surface area is PI*3.66*12.6 = 144.88 m^2 = 144.88*3.28^2 = 1,558.65 ft^2. And the wing-loading is 8,800/1,558.65 = 5.64 psf. I wonder what would be the heating for wing-loading that low.

GW, I was reading some posts on your blog about the reentry problem for large crewed vehicles on Mars:

http://exrocketman.blogspot.com/2012/08 … orbit.html

I was interested in this passage:

The scope of this study covers ballistic coefficient ratios from those of the successful unmanned lander probes (around 100 kg/sq.m), through values appropriate to prior manned capsule designs (300 kg/sq.m), to the much larger values thought to be associated with very large landing vehicles appropriate to manned landings (500 to 2000 kg/sq.m).

I was wondering if a cylindrical rocket stage entering broadside would be enough to get the ballistic coefficient down to the range that we already know works for our unmanned probes, 100 kg/sq.m. Then we could use the final landing techniques that we already know work for our unmanned probes, i.e., parachutes plus small rockets for the final stage, just before impact.

For the ballistic coefficient of the Falcon 9 v1.2 upper stage, the hypersonic drag coefficient of a cylinder I found after a web search to be 1.33. Then for this stage entering broadside, the surface area presented to the airsteam would be half the full surface area of the stage, so 72.44 sq.m. Then the ballistic coefficient would be 4000/(1.33*72.44) = 41.52 kg/sq.m. Then if we added a capsule of 6,000 kg mass this would at about the 100 kg/sq.m. needed to be able to use the familiar final phase landing techniques.

That's just using the standard rocket stage. If we opened up the stage using the clamshell method the surface area would be doubled. The drag coefficient probably would be better also since its concave shape would likely provide more drag. So the ballistic coefficient would be less that 20 kg/sq.m., and we could use a 16,000 kg capsule and still get the ballistic coefficient less than 100 kg/sq.m.

Note what I'm envisioning here is unlike the reusable Earth SSTO case, I'm thinking now of opening up the actual stage, so it would be disposable, ejecting the capsule, then using the capsule parachutes and rocket thrusters for the final landing approach.

Bob Clark

Last edited by RGClark (2018-01-14 17:46:44)

SpaceNut · 2018-01-14 12:12:30

From what I understand Drag and lift are both surface area equations for atmospheric entry to which if you ride the fuel stage with a layer of protection down deep in the atmosphere then it does slow the entire vehicle but then again we have more mass to add to that momentum equation that we are trying to slow....

RGClark · 2018-01-15 00:11:16

SpaceNut wrote:

From what I understand Drag and lift are both surface area equations for atmospheric entry to which if you ride the fuel stage with a layer of protection down deep in the atmosphere then it does slow the entire vehicle but then again we have more mass to add to that momentum equation that we are trying to slow....

You're right. I didn't consider the lift component of the reentry. The open clamshell configuration would have significant lift. So you may want to enter at the best angle of attack to maximize this. Using lift would extend the time you stay up in the atmosphere, i.e., you would not descend as rapidly into the dense lower layers, so this would reduce the heating.

Bob Clark

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