Update 10-11-13:This material was incorporated into a paper on ceramic entry heat shields that GW presented at the 16th Annual Mars Society Convention in Boulder, Colorado, in August 2013. That paper was very well-received. That presentation and many others were videotaped for publication on YouTube. Those videos are now available. Be aware they take a while to load, being 20-to-30 minutes long.
GW's presentation at the 2013 Mars Society convention on a lightweight thermal protection ceramic material is available on Youtube:
Reusable Ceramic Heat Shields - GW Johnson - 16th Mars Society Convention.
(Control+click to follow link)
Using yttria-stabilized zirconia instead of alumino-silicates could increase the operational surface temperature limit to something on the order of 4000 F (2204 C), making LEO entry shields far more feasible at attractive ballistic coefficients. It remains to be seen whether available fibrous zirconia products could be successfully processed into a low-density re-radiating heat shield with redundant retention and enhanced structural strength. However, I am now looking at these materials for another combustor liner application. No answers yet.
The spaceplane designer’s dilemma is that stagnation point heating is dramatically reduced by flying dead-broadside during entry, but, that same broadside attitude causes extreme air loads that rip off wings. Plus, highly-swept wings cause bad landing characteristics.
This dilemma is not a solvable problem in fixed geometry, no matter how highly-swept the wings might be. The real solution lies in some kind of variable geometry.
Of many possibilities, the one not involving jettisoning major hardware might simply be folding the wings into the wake zone behind a fuselage. That fuselage then enters dead broadside, and thus as blunt as possible, like a space capsule. This is not a new idea, having been proposed unsuccessfully by Lockheed in the 1950’s for what eventually became the X-20 Dyna-Soar project.
The results of my bounding analysis study implied a folding wing spaceplane, with straight wings of subsonic airfoil section. It unfolds long after entry, while hanging from a parachute, at entirely subsonic speeds. The wings are unfolded during ascent, for purposes of abort.
In order to obtain characteristics more like those of an ordinary aircraft instead of a traditional spacecraft, a robust non-ablative heat shield is needed, so that periodic refurbishment or replacement is unnecessary. The same simplified entry bounding calculations indicate that this might be possible with a non-ablating ceramic heat shield for dead-broadside entry attitudes, if the ballistic coefficient can be kept low enough, and the bluntness high enough.
A low-enough stagnation point convective peak heating makes a low-density re-radiative ceramic heat shield feasible, because the equilibrium skin temperature for re-radiation of the heat load is low enough to stay within material limitations. One example of such a material is the ceramic tiles used on the Space Shuttle. The black ones on the windward side were useful at higher heating rates than the white ones on the leeward side. That is the influence of surface color upon effective or average emissivity.
The entry environment at Mars is generally less demanding than that for entry from Earth orbit, primarily because the vehicle velocities are far lower there. The low-density re-radiative ceramic heat shield would be even more feasible there. The main advantage is the very low weight of such a heat shield, something very important to any vehicle designs intended to be used at Mars, and especially to those intended to fly more than once at Mars.
One such material already available is the low-density ceramic tile used on the Space Shuttle. Another version of this same low-density ceramic material that is less susceptible to damage, and less expensive to install and maintain, than Shuttle tile, would be very highly desirable. Such a material may actually be available, although it is still quite experimental.
In 1984, I constructed a low-density, fabric-reinforced ceramic composite material, and used it as an experimental low-conductivity, non-ablating monolithic liner insert in a very small ramjet combustor. This was, in turn, part of a device called “Warm Brick”, whose purpose and characteristics are not germane here.
Entry Conditions Comparison
Entry from low Earth orbit (LEO) was modeled as a surface-grazing elliptical transfer from a circular orbit at 600 km altitude. From low Mars orbit (LMO), the same kind of surface-grazing transfer was assumed from a circular orbit at 200 km. These provide very shallow trajectory angles at the entry interface point, as demonstrated in the article dated 8-10-12 and titled “Big Mars Lander Entry Sensitivity Study” on this site.
Additionally, a direct entry at Mars from interplanetary transfer was also investigated (see the article dated 8-12-12 and titled “Direct-Entry Addition to Mars Entry Sensitivity Study” on this site. This direct entry was assumed to be a “typical” higher velocity, but at the same very shallow entry angle, mostly for analytical convenience. Control of entry angle from direct transfer trajectories can be difficult, just as it was returning from the moon. These three scenarios are compared in Figure 1 (all figures are at end of this article).
The simplified entry analysis is very old and very simple: an exponential model based on a density scale height approach. This was originally developed for the early warhead entry work of the 1950’s. It derives ultimately from the work of H. Julian Allen at NACA, more recently referenced in the work of Justus and Braun at NASA. I published an item dealing (in part) with this model as the article dated 6-30-12 and titled “Atmosphere Models for Earth, Mars, and Titan” on this site. I corrected the convective heating model given in the Justus and Braun report, in the article dated 7-14-12 and titled “’Back of the Envelope’ Entry Model” on this site.
For entry velocities under about 10 or 11 km/sec, convective heating dominates the entry heating, and is usually empirically taken to be proportional to the square root of the density/nose radius ratio, and to the cube of the velocity. At hypersonic speeds, vehicle drag coefficient is essentially constant, allowing the use of a very simple constant ballistic coefficient model.
The Earth orbital spaceplane model is depicted in Figure 2. Unfortunately, that illustration includes mixed unit systems, but the key parameters ballistic coefficient and bluntness (belly or heat shield radius of curvature) are given in metric. That design concept required only a modest increase in belly radius (from the original Spaceplane1 configuration to the final Spaceplane2) to achieve a peak heating low enough for a high-emissivity ceramic shield to be feasible.
In this type of bounding analysis, the inherent uncertainty is large, while for the LEO scenario the surface temperature margin is nil. Therefore, this Earth return application of low density ceramics may, or may not, actually be as feasible as it seems. That would require far more sophisticated design analyses to resolve. Margins were definitely larger for both Mars entry scenarios, so the use of these materials is very definitely a feasible option there.
The Mars entry vehicle application is completely different from the spaceplane, being essentially a ballistic entry capsule shape of very large size (60 metric tons), to support manned landings and major cargo shipments there. This was originally envisioned as a one-use chemical vehicle, with a simple rocket ascent stage, as depicted in Figure 3, and described in the article dated 8-28-12 and titled “Manned Chemical Lander Revisit” on this site. However, the same basic shape and concept might also apply to any refueled, reusable “landing boat” or “Mars ferry” in that size class.
Key to the success of such a design is a substantial reduction in heat shield weight, and the elimination of heat shield refurbishment and replacement costs. This particular Mars lander vehicle model was used simply for analytical convenience, not necessarily strict realism.
Re-Radiation Equilibrium Analyses
For a non-ablative, re-radiating low-density ceramic heatshield, there is convective input from the entry flow, re-radiated output due to a high skin temperature, and essentially zero conduction through the material. There is essentially no radiative heat input from the shock layer at entry speeds under about 10 or 11 km/sec. The Stefan-Boltzmann law applies, and with no ablation products in the shock layer, re-radiation is essentially unobstructed. Whether you consider radiation to space at 3 K, or to the Earth at around 300 K, makes no real difference. It might as well be 0 K at this level of analysis.
The controlling factor in the balance is the spectrally-averaged emissivity of the surface. Both Shuttle tile and my experimental composite are alumino-silicate materials. These are normally white in color, and have high reflectance at visible and near infrared wavelengths, and low reflectance at long infrared wavelengths. These are most definitely not “gray emissivity” materials.
In essence, the emissivity of these plain white alumino-silicate materials is roughly 0.2 from visible wavelengths up to around 2 microns, rising to about 0.8 at around 3 microns, and then remaining high at the longer wavelengths. Both my old composite, and the white leeside Shuttle tile, fall in this category. A conservative average emissivity value would be 0.2, it could be a little higher.
All of these alumino-silicate materials (regardless of form) melt near 3200-3250 degrees F, and experience a solid phase change that very seriously degrades structural properties, near 2300-2350 degrees F. Thus the material limit is only about 2300 degrees F. The mechanism of degradation is irreversible shrinkage and the resulting cracking, plus a very serious embrittlement of the affected materials.
Black-coated windward-side Shuttle tiles have a much higher average infrared emissivity, in the vicinity of 0.8 or perhaps higher. For black Shuttle tile, this was done with a glass-like surface layer. I never developed a black surface coating for my old “Warm Brick” composite material, although it could be done. Carbon black (by itself a very nearly spectrally “black” material, at average emissivity above 0.90), used as a pigment in the surface cement coating, offers one promising and easy-to-investigate avenue. So, I assumed that an average emissivity of 0.8 was possible for it, as well.
This energy balance showed that the critical value of peak stagnation heat flux (during entry from LEO) was 25 W/sq.cm, for the allowable skin temperature of 2300 degrees F, and an average emissivity of 0.8. The low emissivity white surfaces were not feasible for that mission. This value set the belly radius finally settled-upon for the Spaceplane2 configuration.
See Figure 4 for the black surface (high-emissivity) data, and Figure 5 for the white surface (low emissivity) data. Meltpoints and phase change limits are shown on these figures. If one exceeds the phase change limit at the surface, the material is definitely damaged, but will likely still protect you through that one entry, but then also likely require replacement after landing. The LEO scenario corresponds to the 25 W/sq.cm heating in these figures.
For Mars direct entry (shallow angle) at a generic 5.6 km/sec, the entry model showed a peak heat flux of about 11 W/sq.cm. From LMO, the peak heat flux was only 2.6 W/sq.cm. These were generated with the “Mars Lander Revisited” model, which uses a ballistic coefficient of 400 kg/sq.m, and a heat shield radius of curvature of 12.4 meters. These might be “typical” of any ballistic capsule shape, in the vicinity of 60 metric tons mass. These are the other two heating levels shown in the same two figures. Skin surface temperatures were well within the feasible range for the low density ceramics (of either type).
Entry Scenario Conclusions
A black surface (high emissivity) low-density ceramic heat shield might possibly be made to work for smaller (lower ballistic coefficient) vehicles coming back from LEO. In that scenario, one must be very careful balancing achievable surface emissivity against all possible reductions to peak heating.
The white surface (low emissivity) version cannot serve in the LEO scenario.
A black surface (high emissivity) low-density ceramic heat shield could definitely serve for vehicles landing directly upon Mars from an interplanetary transfer orbit. There is considerable margin below phase change limits to support a wide variety of such designs.
The white surface (low emissivity) form of the low-density ceramic heat shield is not feasible for direct entry at Mars.
Both the white surface (low emissivity) and black surface (high emissivity) forms of the low-density ceramic heat shield are feasible for a wide variety of vehicle designs entering from LMO at Mars.
In the appropriate roles, both white and black surface shuttle tile could serve in these heat shield roles, as just described. They could save considerable weight, but are subject to the same vulnerabilities, and the same high maintenance costs, as were demonstrated on the Space Shuttle. The real remaining question is whether the old “Warm Brick” ceramic composite material could also be feasible.
The Experimental Composite as an Alternative Material
The “Warm Brick” feasibility test device included as a part of its assembly a subminiature ramjet combustor, depicted in Photo A. In that photo, the inlet is to the right, and the combustor to the left, connected by a narrower inlet mixing tube. The combustor shell piece has the same exterior and interior dimensions as a piece of schedule-40 two-inch pipe, about 4 inches long.
The combustor shell itself was used as the insulator molding shell, as depicted in Photo B, along with the slightly-tapered wooden plug that formed the inner insulation surface. There was also a shell-and-plug combination to form the combustor nozzle from the same low-density ceramic material. These are also shown in the same photo.
The first article (not shown here), was a simple low-density molding without any reinforcement. It did not survive rich blowout instability in the combustor. For the second (successful) article, alternating layers of the low-density molding compound and a standard fire curtain cloth (both alumino-silicate) were emplaced upon the molding plugs, then forced into the molding shells, and cured all-at-once. The nozzle was made first, then used as part of the mold tooling setup for the combustor liner.
Once these articles were “cured” (by driving the water off by several hours’ exposure in a low-temperature oven at 215 F), then the surface was painted with a ceramic cement. This material is normally used to bond two molded parts together, but for this purpose, it acted as a gas-impermeable hard surface layer, same as Shuttle tile. The same low-temperature oven “cure” hardens this cement.
Photo C is a view of the second article, removed from the combustor shell, after many hours of accumulated burn time at full-strength fuel-air mixtures. The view is aft looking forward, toward the sudden-dump flameholder lip. Photo D is a view into the corresponding nozzle, looking aft.
The near-pristine condition of the ceramic is quite evident, in spite of exposures resulting in internal skin temperatures approaching the actual meltpoint of the alumino-silicate materials. After some (but not all) of these tests, some localized surface melting was noted. Strength properties were never evaluated; however, the rich blowout instability it survived was near 0.8 atm amplitude, at audio frequencies (several hundred Hertz).
A recent re-analysis of one of those tests provided a crude (but realistic) estimate of the as-built thermal conductivity of this material: 0.02 BTU/hr-ft-F = 0.035 W/m-C = 0.00035 W/cm-C. The density was never characterized, but falls in the range of commercial Styrofoam products. In both respects, this material strongly resembles the properties of Shuttle tile material.
The differences with Shuttle tile are threefold: (1) this material is substantially tougher and stronger, (2) this material was not bonded to the combustor shell at all (being a free-standing internal insert), and (3) this material can be made (and used) in very large-dimension monolithic parts.
Figure 6 shows a concept for an externally-mounted heat shield panel. It is not bonded to the baseplate metal panel, being retained instead by the ends of the fabric reinforcing material layers, wrapped around the edges of the base plate, and clamped in place. Panels like this can be simply bolted to the framing members of the vehicle. (Any vehicle capable of entering dead broadside will have substantial internal framing.) The “trick” is to lay-up and cure this item as one monolithic piece, then add the surface coating to the cured item.
Update 12-19-2013: Actually, the ceramic composite can be bonded to the baseplate as well as mechanically retained by holding onto the reinforcing fabric. That make redundant retention possible, something the Space Shuttle never had.
The actual materials used back then are no longer available, but I have recently contacted those manufacturers, and determined that modern equivalents are currently available. The molding compound provides an open, porous matrix made of flakes and fibers of alumino-silicate mineral crystals. The fire curtain cloth is woven from yarns of alumino-silicate fibers, that are each about 0.020 inch (0.5 mm) diameter.
Update 12-19-2013: Actually, the original materials are still available, but no longer show as catalog items. You have to ask for them by name and product ID number. And, the ceramic cement manufacturer agrees that, up to a point, carbon black can be used to create a black instead of white hard surface coat.
Other than actual testing, there are two things that need to be done to support application as an entry heat shield. One is a black (high emissivity) surface coating. The other is the detailed process development for constructing the illustrated panels. For the black coating, it might be possible to mix significant amounts of carbon black into the ceramic cement. How that would affect cure and processing is unknown, but it is the most obvious idea to try. The process development work for a shield panel assembly is seemingly pretty straightforward (it just needs to be done).
Note: if this kind of engineering, or this kind of novel ceramic composite material, interests you, do please contact me about consulting help. Be aware that some of this (beyond this level of detail) is subject to the ITAR restrictions, though.