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You guys weigh in, tell me which propellant combinations might actually prove practical to manufacture on Mars. I simply don't understand propellant manufacture well enough to pick the winner. The range of lander masses that I got was 34 tons for LOX-LH2 all the way up to 90 tons for NTO-UDMH (see the "exrocketman" posting I cited above in #399). I got 68 tons with LOX-CH4 and 78 tons with LOX-RP1. All were sized for a fixed payload mass of 3.2 tons. Metric of course.
It makes quite a difference, to both the mission transit design, and to what has to be maintained as a reusable vehicle. Whatever we use on that first manned visit to Mars should stay there, and continue to be used with locally-manufactured propellants. The landers can even be used suborbitally for very long-range transport of critical items, between sites very far apart.
GW
This one has been inactive for a while now. But, I wanted to let everyone know I have looked at a reusable single-stage chemical Mars lander, and found it to be feasible. I posted this in a short form over at http://exrocketman.blogspot.com in a posted dated 8-31-13, and titled "Reusable Chemical Mars Landing Boats Are Feasible".
As for the ceramic heat shield material in my last post above in this thread, I determined it really will work, and presented that outcome in a favorably-received paper at the 16th convention in Boulder. It was part of my reusable lander feasibility study, with PICA-X ablatives as a backup.
GW
Cookies-or-not, this pattern is very consistent. Second attempt login is successful, as long as I have not looked around on the site.
GW
I'm just guessing that a glass made of basaltic material will be a hard, opaque product somewhat similar to natural volcanic glass here on Earth. There would be a variety of uses for such a product, ranging from tools and dishes to building material.
There is a definite need for a transparency material. As far as I know, that is silica glass. There is silica sand on Mars, just "not everywhere" in concentrations that would be useful. The places where it is, should be identified. That's where you put your transparency factory.
Ultimately, surface transport by truck or train will be the most practical, but until that infrastructure exists, you will have to fly in order to transport people and product. The air is too thin for airplanes, but suborbital rocket flight would work. So, the same landing craft that put people on Mars could be used for this, if (and only if) they were designed from the outset to be reused with a long service life. You just displace some propellant weight with more payload weight for suborbital flight (that is, provided your craft was designed with the extra volume to contain the extra payload).
You will not get that kind of service out of something with a 5% structural inert weight fraction.
GW
Hi Bob!
I saw your post on this yesterday. I quite agree that SSTO is feasible, especially at around 5% inert-weight structural fractions. It's all in the price reduction for 1 stage vs the cost increase for 2-stage vs the rather-sharp unit price increase for the lower available payload fraction with 1 stage. In point of fact, that's what NASA is doing with its Senate-mandated SLS design. Theirs is a one-stage core (with some SRB's) that essentially pushes a huge mass to LEO. The upper stages they are looking at are mostly for departure from LEO. Or for GTO work. It's just that there are no commercial payloads yet that are big enough to need a rocket that size. The next commercial step is Spacex's Falcon-heavy at 53 metric tons to LEO out of Canaveral.
I don't think 5% inert fraction rockets will ever be practically reusable with the materials we have, despite what Spacex is attempting with its Grasshopper test vehicle and reusable-Falcon concept. To date, I know of no liquid-propellant stages that have ever survived the tumbling airloads on striking dense atmosphere supersonic/hypersonic. That's why Spacex has never yet recovered a Falcon-9 first stage in a salvageable condition: just too fragile not to break up. They did try, at least initially. And that doesn't even consider ocean impact loads, no matter how big the parachutes.
Whether Spacex can control the attitude of a Falcon stage for rear-end-forward thrusting recovery is problematical, and even they admit that. The combination of thrust and airloads entering like that is far beyond anything the lightweight stage was ever designed for in its ascent. That is one tough row to hoe. I wish them well, but I am not holding my breath about success. Neither is Musk.
I rather doubt that survivability can be obtained at 10% inert fractions. I think minimally-survivable (relatively short service life) inert fractions will look a lot more like 20+%, and that's using composites to-the-max everywhere the aeroheating won't damage them. Done all-metal, such inerts would fall in the 30-40% range. The longer the design service life, the more structural weight is required. The X-15 was right at 40%. Most high-speed bombers are around 50%, with Navy birds pushing 60%. These materials are all only so strong. You have to employ enough of them to take the punishment your mission environment dishes out. That's heavy, no way around it.
This situation is alien to most of the rocket designers out there today, SRB's excepted (and they just "lucked out", they didn't really know). No one has ever seriously tried to survive reentry with liquid booster stages before. The usual rules-of-thumb for rocket design don't apply, it's a completely different environment, and way far more challenging. The SRB guys sort-of survived parachute recovery at sea (a lot of these segments were too damaged to reuse) only because those were 900 psi pressure vessels, of inert weight fractions 10+%.
GW
Last I heard about quantum entanglement, you could send a signal, but not understandable information. Not sure anybody really knows, not yet. Edge-of-the-known physics stuff is always very iffy and subject to reversals-of-opinion.
GW
You'll have to use the airbreather for that portion of the trajectory where speeds and densities are appropriate.
A good supersonic-inlet (subsonic-combustion) ramjet design, aimed at high speed capability, will give useful thrust and impulse between takeover at around 0.5-0.6 km/s velocities, and its top speed of about 1.5 to 1.8 km/s. Those are predicated on stratosphere speeds of sound (295 m/s). The net effect on a depressed HTO trajectory so that the airbreather supplies 100% of the thrust during its operation is an acceleration through 1.1 km/s out of the required 7.7-plus-gravity-and-drag-losses. Drag losses are simply enormous on depressed trajectories, however.
If you do it as ramjet assist on a vertical launch trajectory, you go for the low-speed design with pitot inlet and lower performance capabilities. The rocket leaves the sensible atmosphere at around 80,000 feet (25-ish km) at around Mach 2-ish (around 0.6 km/s), long before the first stage burns out at 3-ish km/s in vacuum. That's where you stage off the airbreather, which becomes useless as the air thins to nothing. That kind of airbreather ignites subsonically at about Mach 0.7-ish (0.2-ish km/s), and might supply at most 20% of the ascent thrust.
To burn at speeds resembling 4 km/s (Mach 13-ish), would require scramjet. Scramjet is simply not yet "ready-for-prime-time", despite the 2 (of 4) successes achieved by the X-51 test vehicle. Those flights demonstrated engine burns in minutes at Mach 5. ASALM-PTV the ordinary ramjet accidentally (!!!!) flew Mach 6 3 decades ago. It was only designed to cruise Mach 4, but we had a throttle runaway on one flight test. Subsonic combustion dump combustor, kerosene-like fuel, supersonic inlet and nozzle, technologies considered well-developed since WW2-to-late-1960's.
I'm not sure vertical launch ramjet assist is really worth the trouble, but that answer remains to be found. For HTO spaceplanes, ramjet looks pretty promising, if you can stage near Mach 5 to Mach 6, at about 60,000 feet (just under 20 km), and manage to pull up to about 40 degree path by adding in some rocket thrust somewhere, all on a short transient. Your first stage is then a fully-reusable rocket/ramjet-equipped airplane. No new technologies need be developed to build it, this is all "prime-time-ready" stuff.
The delta-vee required of the second stage is (unfortunately) near 5.9 km/s, so the propellant fraction is inherently pretty high. At reusable inert fractions, payload fraction is inherently low, which is bad for economics. However, there is always a niche for this capability, down in the low payload mass range, where the first stage aircraft is not too large.
GW
Josh:
I just found out something interesting regarding the log-in difficulties. If I go straight to log-in without looking around on the forums, it logs me in on the second attempt, even on this Windows 8 laptop. If I look around first, I cannot log in no matter how many attempts. The clue was logging in on the second attempt after leaving the website and then coming back. It is the second attempt, always, but it does work.
GW
Hi Midoshi! Sorry I didn't recognize who you were. I saw the MAVEN presentation at the convention. Very impressive. Keep me posted, and good luck on the launch!
I don't know about you probe launchers, but we in the rocket missile business used to laugh and joke about sacrificing a virgin (or a reasonable facsimile thereof) to the test gods.
Whatever it takes, I suppose. None of the lady engineers or technicians ever volunteered. Yet we still usually had successes, even in development work.
GW
Spacenut's link to the older thread is indeed interesting. Glass-as-a-substitute-for-cement in a "concrete". Hmmmm. Has anyone ever actually made such a material?
What about means to concentrate surface calcium sources into the 10+% range so that glass-as-we-know-it becomes feasible on Mars in those locations where SiO2 is available? I'm looking for clear glass panels for habitats and greenhouses.
GW
Can't do probes forever, unless we never send men. There seems to be little interest in sending the "right" probes: it's boiling down to either/or sending ISRU demos vs. sample return, not both. This is not looking very good.
GW
Hi Terraformer:
I'd be surprised if a single-stage vehicle with a 20% inert weight fraction was actually robust enough to be long-life reusable. The X-15 was very reusable at 40% inert, but did not take off by itself. Most high-speed (subsonic!) bombers are around 50-60% inerts with all-metal construction, and have been since WW2.
If you replace around half the structure with composites, you might achieve around 27% inert in what was 40% all-metal. You can't replace it all, because composites have such a low tolerance for heating, which happens in-spades on entry, and somewhat (or more) on ascent.
Work-in a "reasonable" payload fraction, like 5 to 10% to make the economics easier to achieve, and there is (with the 27% inerts) about 63 to 68% of the vehicle's gross weight available to be propellants. Those are mass ratios of 2.7 to 3.1, too low for SSTO with chemical rocket propulsion, even as a vertical launch rocket (for min gravity and drag losses). It's worse with the depressed trajectories HTO spaceplanes fly: very large gravity and drag losses.
The airbreather might help, but its higher Isp is not constant, and only available over a small portion of the trajectory. Further, you have to blend it with rocket thrust, in order to climb, especially in the thin air around 60,000 feet. There's not much frontal thrust density available from the airbreather above that altitude. That's why I doubt that ramjet might make SSTO spaceplanes possible, when you restrict the design to larger payload fractions for the economics. TSTO, maybe.
GW
Midoshi:
Sorry I missed you. We (wife and I) were there at the Saturday banquet with Josh Friedman.
GW
Sounds to me like glass might be fairly easy to make in certain locations on Mars. All one needs is an appropriate furnace for the melt. The rest should be like here, except maybe the casting and solidification. The cooling rate might need to be retarded at Martian conditions, to keep the glass from cracking or turning back to crystalline (sand or chunks) form. We cool it fairly slowly here to keep it in the amorphous state that is glass.
This does apply only to locations where SiO2 sand is actually available, though.
GW
Hi Terraformer:
I took a look at the blog posting via the link in your post just above. Here's my response:
Well, I think arguments based on mass ratio and “average” Isp are too crude to get you anything useful for ramjet (or any other airbreather). You need a real cycle analysis, which should be a subroutine in a real trajectory code, or which you can use for point performance calculations over a flight envelope for empirical correlation. (You still need a real trajectory code.) I’ve done both, they’re both effective approaches to first order. Fixed averages are not. Sorry, that’s a simple fact-of-life.
You do need to understand thrust and drag accounting, because if you don’t, it is really easy to leave out some very important drag forces in your force balance. I’m not talking about basic ram drag here (the “airbreather’s burden”), I’m talking about things like additive drag, spillage drag, diverter drag, and bleed drag. These are quite important, both at takeover, and at very high flight speeds. These are neither trivial to understand, nor trivial in their effects. You need some training in propulsion aerodynamics for these. This isn’t basic physics textbook stuff, and never will be.
All this stuff will be in my ramjet book, which is not yet ready for publication. Its intended audience is engineers working in ramjet propulsion, whether for missiles, or for launch vehicle work. I’m still trying to “rough-write”-down all the topics, but I think I have most of them documented in rough form, just not all of them. All this stuff is currently very rough first-draft stuff, and will need extensive re-organization and re-write, before it is book-ready. But, I really am working on it.
There are two speed ranges for ramjet design, “low” and “high”. Low speed range designs have simple pitot (normal-shock) inlets, convergent-only nozzles, and can be ignited at subsonic speeds. They will show nacelle thrust greater than drag down to very low speeds, but will have specific impulse lower than composite solid rocket, below about half a Mach number. Peak specific impulse potential is at about Mach 1.1 or so, at about half or 2/3 the max Isp potential of supersonic designs. Max useful speed is about Mach 2, or maybe Mach 2.5 at the very outside. With hydrocarbon fuels of almost all types, about the biggest nozzle throat/combustor area ratio is 0.65, limited by flame-holding considerations. Performance at lower area ratios is inherently lower.
High speed-range designs feature external compression features like ramps or spikes that protrude ahead of the inlet cowl lip. They also have almost-zero thrust potential below about Mach 1.6 to 2. But, they work just fine to about Mach 5-or-6, depending far more on vehicle drag characteristics, than anything about the ramjet engine design. With kerosene fuels, peak Isp potential is around 1200-1300 sec at about Mach 2.5-ish, lower slower, and lower faster. Nozzles are C-D, but exit “bell” area ratios are closer to 1.5-max, than anything to do with the expansion ratios one sees in rockets. With hydrocarbon fuels of almost all types, about the biggest nozzle throat/combustor area ratio is 0.65, limited by flame-holding considerations. Performance at lower area ratios is inherently lower.
These things can be very lightweight, depending upon whether it has to be re-usable or not. The “best” designs have been one-shot missile designs, with an ablative combustor liner, for missile speeds up to about Mach 4. External heat protection is also an issue, from about Mach 3 on up for reusable designs, even with steel construction. There are air-cooled perforated liner designs from the 1940’s and 1950’s that would actually work to Mach 6 on a transient, exclusive of external heat protection problems. There are ablatives that would work externally to Mach 6 on a transient, but these have replacement issues. Missiles generally always use ablatives inside, and maybe outside, if needed.
There is my oddball ceramic-ceramic composite combustor liner material, which offers considerable potential for a re-usable combustor. It might also serve as external heat protection, for a fully-re-usable design. This is still an experimental material, though.
Ramjets require boosters to reach takeover speed: about Mach 0.5 to 0.8 for “low speed” designs, and about Mach 1.6-to-2 for “high speed” designs. For one-shot missile applications, the best choice has proven to be the “integral rocket-ramjet”, wherein a solid rocket booster is cast or loaded within the ramjet combustor. This requires an appropriate ejectable booster nozzle nested within the ramjet nozzle, and some sort of inlet duct obturator, usually ejectable or frangible port covers. Re-usable launch applications might well be “best” with parallel-burn rocket and ramjet engines in the same airframe. It really helps if the rocket and the ramjet use a common fuel.
From a flame-holding standpoint, I think the dump combustor has “way-to-hell-and-gone” more potential than the V-gutter, or can, or “colander” (or any other type of obstruction-type) flameholder. Dump combustors have very little sensitivity to dump plane speeds, compared to any of the blockage-element types. Variable speeds at the dump are inherent with launch accelerators, whether vertical-launch or horizontal takeoff. Almost no textbooks describe dump combustors. My book will.
Ramjet liquid fuels can be any kerosene (or kerosene-like synthetic), or any liquifiable hydrocarbon. The early engines with subsonic ignition used mainly low-grade gasoline. Today, in supersonic-inlet designs, JP-4, JP-5, JP-7, Jet-A, Jet-B, Jet-A1, RP-1, K-1 kerosene, a synthetic variously known as RJ-5 or Shelldyne-H, and even liquefied methane, are all very attractive candidates.
I have even used propane, but it and LPG are not all that attractive, for their inherently-heavy fuel storage considerations. LCH4 will require extra care to insure full vaporization, and extra care with flame-holding issues. RJ-5 is a synthetic that resembles kerosene, except that its density is substantially higher. It was used in ASALM-PTV, with one test that reached Mach 6.
I hope the book might be available in a year or two. It’ll be the ramjet analog to the famous (or infamous) “drag bible” written long ago by Hoerner.
GW
Sorry, Bob. I was responding to Terraformer's question about a reusable SSTO spaceplane, presumably horizontal takeoff.
Actually, I quite agree with you about one-shot throwaway SSTO rockets. The original Atlas was very nearly one. The Titan II first stage certainly was one. You can have a higher payload fraction as TSTO, but at a higher launch price. Whichever flings payload to orbit at lowest overall $/payload mass is the choice most folks will select (especially if they fail to consider the effects of not flying at full load). The 1960-on engine technologies and the modern stage structural fractions in the 5-10% range make SSTO vertical launch rockets quite feasible.
A reusable vehicle is another ball game entirely. There is just no way, even with modern materials, to survive entry and fly again with structural fractions in the 5-10% range. I'd be very surprised (and pleased) to see a design at 20% inerts that actually proves-out reusable with a "reasonable" life. I think real vehicles like that will fall in the 25-35% structural inert range.
Not even the miracle composites we now have are going to help, because of ascent and descent aeroheating, tumbling airloads aside. All the organic composite matrices are debris and powder by 290 F (143 C). Tumbling airloads just make that problem immensely worse, or else you must carry the weight for thrusters and drogues to control vehicle/stage attitude very carefully.
And we haven't even begun to talk about ocean or dry-land surface impact loads with any practical chutes. There's a reason 1/4 inch steel plate is considered to be "wastepaper trim material" by marine/shipbuilding engineers, and it is a bloody good reason, too.
GW
Seems to me that someone identified what may be dust-covered glaciers or other massive ice in the northern lowlands. That would be in the ancient sea-bottom, near the old shoreline. I don't remember exactly where this was, and it was reported a few years ago. I never heard how this discovery panned out, but I do remember photos of what looked like dirt-covered chunks of sea-surface pack ice. I don't remember the scale of the photos, but my impression was that these chunks were very large, perhaps km in dimension.
If this interpretation could be confirmed, then we have have a promising landing site with scads of minable water not covered with monolithic rock. We'll need to include a deep drill rig and probably some kind of bulldozer and backhoe with the ISPP/ISRU plant that we land. There's your hydrogen (and some oxygen).
In the northern lowlands, it might not be that hard to freeze-out some dry ice, precisely because it is cold. Once solid, it's easy to pressurize CO2 to many bars by confined (solar or waste) heating, no moving parts, nothing to regenerate. Even easier than that absorption thing, and maybe lighter and more reliable, too. There's your carbon. LOX-methane is a nice propellant combination. Better than kerolox Isp, and almost as dense.
I wonder whatever became of that identification of dirt-covered ice, and if there is any way to confirm/deny it with the data we have, and the assets already orbiting Mars? That would seem to be a very important piece of info that we need for planning any manned landing.
GW
Louis:
I haven't seen this stuff since I was a kid! Many thanks! What memories it brought back!
These were the proposals for spaceflight coming from Von Braun's group in the late 1940's, some of which was built upon ideas they came up with in the late 1930's and early 1940's over in Nazi Germany. The rockets bear a close kinship to the V-2 (actually A-4), its two-stage cousin the A-9/A-10 ICBM, and its winged cousin, the A-4b (which actually did fly in combat). But, of course they do.
Add to this the early 1950's thinking about cesium-propellant ion engines, and you have the genesis of the 1956 Disney "Tomorrowland" short films on spaceflight, the last of which dealt with a slowboat trip to Mars in a fleet of giant ion-propelled ships (resembling umbrellas) with rocket landers. I think I can get those old films on DVD now. What a hoot. And, some of those ideas are still good today, but not all.
In those Disney films, the Earth orbit ferry ships , the wheel space station, and the moon ships, were all exactly these same late 1940's Von Braun stuff.
GW
This is for RobertDyck, Bob Clark, Louis, and several others in this thread:
RobertDyck: I think you and I are not arguing for things poles apart, which I seem to have given you the impression that I am. I am sorry that I was not clearer. Like you, I think the typical NASA mission proposal is bloated, cost-wise and mass-wise, by around a factor of 10, if not more. NASA also suffers from a very bad case of NIH (not-invented-here) attitude, and has since around 1960-or-so. (No one else on these forums seems to have noticed that.)
All: I am not suggesting that ISRU/ISPP won’t work, I am suggesting that it is not ready yet to bet lives upon it, and for reasons that have less to do with the technology than what we know about exactly where all the resources are, underground, on Mars. I detect from your (plural) comments that all (or at least most) of you feel the same way I do.
All: RobertDyck is exactly correct in pointing out that without ISRU/ISPP, the “typical” mission costs are simply too high to be feasible for a manned mission yet, because of politics. USA, Europe, doesn’t matter. Yet, if the mission is not led by NASA (or ESA, or any other governmental agency), those same costs might actually become quite feasible!
Louis: of course, I am proposing solar PV or solar thermal for powering any compression scheme, whatever it might be. Why not use the energy that is already there, why ship it from Earth when that is so very bloody heavy?
All: that last being said, there is great ease, simplicity, low mass, and low cost associated with the notion of trapping mined water (or dry ice) inside a container barely larger than the mined solid volume, and heating it with simple solar thermal, as a batch process. As the solid vaporizes, its vapor pressure very rapidly builds into the 1-2 atm range, making further compression by more-or-less standard electro-mechanical means simple, small, and lightweight (and very little different from here at home).
All: Once in that 1 atm range, then all of the Sabatier reaction (and many other proposals) start to look for Mars exactly like their experimental counterparts here at home. But, until you get vapor near 1 atm to compress, nothing about those processes looks anything at all like their experimental counterparts here at home. It’s vapor compression, not the specific process we are discussing, that is the show-stopper/crew-killer, for an ISRU/ISPP technology we can bet astronauts’ lives upon. That’s because the size, mass, cost, energy demand, and complexity of the compression machine (whatever it is) depend more upon inlet density than any other variable. Simple fact-of-life.
All: So, to avoid both super-high mission costs and the very high probability of a dead crew (which NASA has correctly identified as the highest-cost item associated with spaceflight), we need to do reliable ISRU/ISPP on the first manned mission to Mars, as RobertDyck has so eloquently pointed out. Yet, this has to be done at a site with solid minable ice, as I have pointed out more than once, or we will most likely kill that crew!
All: So, we need a site where we already know there is massive buried ice for that first manned mission. Otherwise, nothing we do will reliably succeed, or save mission costs. Absolutely none of the probes and rovers we have sent so far have the capability to answer the question “where is massive buried ice?”, not even Curiosity. None of these can drill or dig more than a few centimeters beneath the surface, when the answer to the critical question requires (at the very least) drilling MULTIPLE METERS beneath the surface.
All: There is maybe the opportunity to send one or two more probes to Mars before we try to send men. If you believe NASA, we will try to send men somewhere in the 2030’s, but I believe this to be “code” for “never”. None of the probe proposals I have seen has drilling capability to 10’s, maybe 100’s, of meters, yet that is EXACTLY what we need.
All: NASA’s typical manned Mars mission “design” falls in the $500B class. I think that someone like a Musk/Spacex (not ULA or “the usual gang”) could do the same job for around $50B or so. Add in reliable ISRU/ISPP, and you might get this down to around $5-10B (or $50-100B if done by NASA/ESA/etc). THAT is where we are. Now, how do we get out of this quandary? I honestly don’t know.
GW
Hi Josh:
I’m sorry, AOA means “angle of attack”, which is aircraft engineering and pilot’s jargon. It’s the angle between your craft’s reference axis and the relative wind vector, zero when you are streamlined to the flow. For an airplane, this reference is usually the fuselage body axis or the main deck line. For a capsule or gun projectile shape, it is usually the axis of symmetry.
What you say about not “bouncing-off” into a longer-period ellipse is quite true of a two-body problem, considering only the Earth and the moon. The trouble with near-escape speed problems is that the orbital mechanics is really a multi-body problem. This situation is adequately explained as a three-body problem: Earth – moon – sun. At near-escape speeds, the sun’s perturbation can “strip” the craft free of the Earth-moon barycenter and place it into a sun-centered orbit, which would have a period grossly similar to that of the Earth-moon system in its orbit about the sun. I remember “them” (both the TV news and the NASA spokespersons) talking about this during the Apollo years.
As for educating yourself about entry dynamics and heating, do what I did. Go find the Justus & Braun EDL paper on the internet. It covers multiple related topics, and is very useful for descriptions of “average atmosphere” conditions on several bodies of interest. That paper also describes the old Julian Allen entry ballistics model, and gives lots of sample results with it. I also remember seeing that model in class when I was a college student at UT engineering school decades ago. But, those notes disappeared long ago. I had to re-learn it from Justus & Braun, and other sources on the internet.
As for Justus & Braun, their version looks OK for the dynamics, but they did not understand Allen’s heating correlation properly. They use the wrong value for the proportionality constant, and inconsistent values by orders of magnitude between the closed-form peak rate equation and the closed-form integrated-total equation. I fixed this when I put that model into spreadsheet form, and added a numerically-differentiated timeline, and then used it to numerically-integrate total heat absorbed from the instantaneous rates. I used the old 1950’s Imperial-units constant, and very carefully converted it to metric for this spreadsheet. The heating rate equation is a dimensionally-inconsistent empirical correlation. It’s pretty reliable, it was the closed form estimate equations that I could not trust.
I’ve got that stuff posted somewhere on “exrocketman”. If you want, I can email you a copy of that spreadsheet model. It’s a 2-D Cartesian thing that gets remarkably close to Apollo’s 11 peak gees when coming home from the moon at the max allowable 2 degrees below horizontal. So I presume, since I avoided all the closed-form estimates, that the instantaneous-peak and numerically-integrated total heating numbers I got for Apollo are also reasonably accurate. At least to around +/- 20%. Ballpark.
The heating in that model is really intended for non-lifting craft. When you fly a capsule off-angle for L/D 0.1-ish lift, the stagnation point moves off-center of the heat shield, and the stagnation heating is a little different than you would predict from that old correlation (which is only an approximate to begin with). With a winged craft or a lifting body, things are so different that the old simple model is the wrong model to be using, mainly because the trajectory dynamics are wrong. Although, at the actual dynamics, the heating correlation still predicts a rough-but-good approximation to nose cap tip, and leading edge, heating rates. Your peak and integrated totals are wrong with the Allen model, because the trajectory is wrong.
This old Allen model was developed for warheads coming in fairly steeply, and it did a really good job. At shallow angles, you get into the mismatch between 2-D Cartesian and true spherical coordinates, and (worse) you get into the error induced by ignoring gravity as “small”. That second one means the trajectory angle at end of hypersonics (Mach 3-ish for a blunt capsule) is predicted way too small (actually, it was assumed constant); in the real world, post-peak-gees is where you start bending significantly downward.
But to zeroth/first order, it’s a pretty good model, even at shallow angles, as long as you are pretty close to zero-lift ballistic. I take its horizontal range prediction, and simply wrap that around the circular girth of the Earth, Mars, or whatever body. I just use its predicted end-of-hypersonics altitude as-is, and (for capsules) then assume a trajectory angle “near 45 degrees downward” at end of hypersonics. Rough-and-ready, but not that bad a prediction. End of hypersonics is local Mach 3-ish for blunt shapes. Closer to Mach 5 for “pointy” things.
That simple entry model is completely inadequate for aerocapture or aerobraking applications. That stuff is just too far away from “steeply-entering warheads”. For that, one needs a real 3-body computational orbital mechanics model, and a real computational trajectory program in spherical coordinates.
It can be 3 degrees of freedom for nonlifting ballistic “particles”, but you need the full 6 degrees of freedom for lifting craft, because they can roll and yaw, as well as pitch. You’ll need a sequence of aerodynamic models from hypersonic to subsonic, and in the case of lifting craft, some model of how you will “fly” your entry. That said, I don’t have any such models in my possession, although I used one like that when I worked on the “Scout” launcher at LTV Aerospace back in the 1970’s.
So, that’s why I seem clueless when aero-capture or aero-braking questions come up, yet so informed when ordinary ballistic entry questions come up. I have the tools to figure one, but not the others.
GW
What I found playing with that oversimplified LEO entry ballistic analysis is that low ballistic coefficient lets you decelerate higher up where density is lower. Doing a minimal de-orbit burn from LEO inherently gets you a very shallow entry angle, which also helps you do the peak deceleration higher up.
Peak heating rates occur close to, but not concurrent with, peak deceleration. They are directly proportional to density^0.5, inversely proportional to "nose" radius^0.5, and directly proportional to velocity^3. Choosing the right trajectory and a low ballistic coefficient get you the lower density effect. Choosing a very blunt shape gets you the "nose" radius effect, and both effects are quite strong at 0.5 power.
But that velocity effect is so strong, it kind of rules out clever peak heating reductions for near-escape entry speeds at Earth. You're basically stuck with piling on the ablatives, and trying desperately to shallow-out enough to reduce peak gees, while not bouncing-off into deep space. (Yep, I know for a lunar return, you're barely sub-escape, so you don't bounce-off on a one-way trip. But, the return time is measured in many months to a few years.)
At Mars it's different. The velocity effect is scads weaker. 5.6 km/s for the typical direct entry from an interplanetary transfer orbit, about 5 km/s escape, and about 3.6 km/s from LMO. It's a whole lot easier to get away from sacrificial ablatives there. Really tough to do that here, with escape at 11.2 km/s. (BTW, an interplanetary free return from Mars is about 15-16 km/s at entry.)
I'm kind of surprised to see 3-4 gees as the recommendation for people, when there have been roller coasters pulling civilians to 5-6 gees for decades now. The coasters are a 10-15 sec transient at peak gee, and so is the peak gee pulse during entry. Recline the seats a tad, and it's not very tough, even for the sick or injured.
I'd consider a winged spaceplane for some people with demanding circumstances, maneuvering capsules for "general use" (people and cargo). A desert capsule landing with a mile or two circular error probable is actually a very inexpensive recovery, not much different from recovering an airplane that landed in the desert. I don't think I'd attempt a very large spaceplane; we sort of found out that was a bad idea, with shuttle.
As long as the spaceplane has fixed wings, you're faced with low AOA during entry, so as not to rip the wings off, and so also the related need for ablatives on the nose and leading edges. No way around that here, not with the materials and technologies available at this point in history.
You're also faced with the very poor landing characteristics of delta wings: 200-300 mph, and AOA near 40 degrees, at touchdown, while riding the rather-"iffy" leading edge vortex to prevent immediate catastrophic stall. Sink rates are inherently high because subsonic L/D is so very poor. And, basic stability is quite difficult to achieve, due to massive cp shift across the speed of sound.
There's a very good set of reasons why the F-102/F-106 fighters and the B-58 bomber were not in service very long. I just gave them to you in the previous paragraph. Shuttle suffered from the same difficulties, in addition to all the other difficulties that are better known.
GW
All the numbers I have run point to two stages with all known non-nuclear rocket and ramjet technologies and materials that we have. The numbers just aren't there for SSTO, not at practical structural fractions, and payloads big enough to be worthwhile.
Fundamentally, there's no reason why both stages of a TSTO cannot be reusable, and this is true whether you design for HTO or VTO. A practical SSTO will require some sort of propulsion breakthrough (yes, I know it can technically be done right now, just with impractically-small payload fractions). I have a lot of hopes pinned on Skylon with its Sabre engine for such a breakthrough, but I'm not betting the farm on it.
GW
Self-compressing by phase change makes a lot of sense to me, for both CO2 and H2O resources. I think the machinery would be a lot smaller, lighter, and less energy-consumptive. Your biggest expenditure would be the heat for the phase change.
Only one trouble, these solid resources (ice and dry ice) are not evenly-distributed the way the atmosphere is. Problem with that atmosphere is it's too bloody thin to use in any efficient way. It means you need to know where the buried glacier is before you land and try to use it. Same for any other local resource.
That's part of what exploration is really all about. It necessarily precedes successful colonization. As history teaches us.
GW
Hi Josh:
Now I understand, your glider is for the people. To take full advantage of the reduction in peak heating of a low ballistic coefficient / low wing loading design, you will need to do your entry from LEO. That's because the velocity dependence is cubed, and that's a huge effect going from 11.2 km/s down to 7.7 km/s. To reduce the heating, you need to hit at the inherently-very-shallow angle afforded by minimal retro burn from LEO.
Getting that deceleration into LEO is going to be some sort of tradeoff between mass ratio required (and the associated propellant logistics) and the total transit time. You could use a very minor burn to put the craft into an elongate ellipse with a perigee below 135 km in the upper atmosphere. That's for aerocapture braking, perhaps multiple passes, but the time to make that orbit is measured in the same sort of days as the lunar transit. That's a lot of days in a capsule for people, not an objection for cargo.
Winged craft have very strict AOA limits, as we saw with shuttle. Plus, hypersonic designs like that have very poor subsonic approach and landing characteristics. I kind of like the old Lockheed idea from the early 1950's of folding the wings (and tails) dorsally, and entering dead broadside belly-first, with aerosurfaces "hidden" in the wake. You could use a straight-wing design with a subsonic airfoil in a folding design like that. The people will feel the peak deceleration eyeball-down, but it might be around only 5-6 gees. That's doable, even by civilians.
The other difference with folding wings is, you no longer need ablatives to protect stagnation zones on leading edges and nose caps, if you can get the stagnation peak heating below about 25 W/sq.cm. The low-density ceramic approach would then work, even at stagnation (belly of the craft). Build a structurally-tough version (like I did 3 decades ago), and you just might have a fully-reusable heat shield that resists damage and is easy to repair.
It's easier at Mars to use this material, but there is a niche at LEO for this stuff. That's my paper next week at the convention.
GW
I honestly don't know much about incorporating aero-capture or aerobraking into any cargo transport from moon to Earth. Launching rockets off the moon could conceivably be replaced by catapults, since the delta-vee is so low.
Yet, the approach to Earth from the moon is inevitably 11 km/sec. Without men on board, you could design for aero-deceleration levels approaching 100 gees, as long as you provided a heat shield. Low-density ablatives like PICA-X offer the best near-term potential for that, which means shipment from Earth. I guess some sort of mineral-wool ablative might be made from local lunar rocks and dust, but that's a material development yet to be done.
If you enter steep enough to pull 100 gees (with the associated high heating), then you pretty much eliminate the risks of bouncing off the atmosphere into deep space. So there is an advantage to it. Whether that aero-deceleration is for aerocapture into orbit, or for direct entry, wouldn't seem to me to be much different. It's still quite hot, either way.
GW