There are too many (!!!) solutions of these problems, so you need to formulate your question more concrete and mention also other requirements to the materials you would like to apply for ultra-high velocity spacecraft thermal protection system.

Maybe metal such as inconel combined with water cooling(emerging water steam would be discharged)... But you need to think of weight of such system too. It is expensive to take additional dead weight to space.

Very big balloons as a way of gentle aerobraking were discussed some time ago...

Peter,

Transpiration cooling is indeed an old and (somewhat) tested technology for TPS uses. I believe that the ex-HoToL people at Reaction Engines are baselining that for Skylon.

However, a more pertinent question to ask Mr Lu, is that generally the constraint is not so much the temperature, as the heat flux. One can always design a craft that results in a survivable temperature, for a given speed. Peter's reference to the 'ballute' is one such example.

For high temperature non-ablating designs I would look at diborides of hafnium and other refractories.

UHTC materials are excellent candidate for your application(HfB2, HfB2-SiC, ZrB2, ZrB2-SiC,...) although they are so expensive. other choices are MAX Phases. they are new materials with special properties. Ti3SiC2 and Cr2AlC and Ti2AlC can be used for your application. be lucky with your researches

Sorry, but most of MAX phases are not stable at higher temperatures...

yes you are right, MAX phases have transformation at high temperature.Under vacuum and inert gasses, they dissociate to TiC, SiC,... and under oxidation situation they make a thin film layer made from TiO2, SiO2 and other oxides that protect them from more oxidation. I have tested  them for 1300 degree of centigrade with no problem. But  before making any decision, you must consider all of the criteria. First, you should determine the maximum temperate, service life, corrosive media, and velocity that may destroy the protective layers. After determining these parameters, you can make choice.

In stead of MAX phases you can use hetero-modulus ceramics (or it is also called as high-E - low-E composites). The melting points of TiO2-SiO2 compositions are very low. For advanced thermal protection systems you need to achieve at least 1600-1800 oC...

You know, I'm not exactly expert for this application, but ceramic such as that which was used on space shuttle was protection you are asking for. But it has its drawbacks...  toughness, brittleness... And when you are going from orbit down to earth with high supersonic speeds of around 25Mach it generates temperatures of 1600-1700°C and even if your material can withstand that temperature, heat must not reach superstructure of spacecraft. That is why I mentioned active cooling together with temperature resistant metal.

Recently I read about very low density rockets which could soar like balloon to 70-100km and then use engine for reaching orbital velocity, but I find it unrealistic.

"...25Mach it generates temperatures of 1600-1700°C..." It is not correct as in reality it is much higher, even sometimes more than 5000-6000°C...

I took info about temperatures from internet, therefore I can be wrong. http://en.wikipedia.org/wiki/Space_Shuttle_thermal_protection_system

 http://science.ksc.nasa.gov/shuttle/technology/sts-newsref/sts_sys.html

If there is 5-6000°C, then I don't see a way without ablative cooling, or active cooling

I think before answering this question, we need more data about your problem. please determine the maximum velocity, maximum service life of spacecraft that you want to make insulation for it. I think, the temperature of 5000-6000 is unbelievable, It is needed to be calculated this temperature versus velocity exactly.

I read a paper under title of Materials Design of Hetero-Modulus Ceramics for Extreme Environments Based on Ridge Effect Phenomenon from professor Shabalin, I think these materials are attractive and may be the answer, although it seems , it is needed more researches for applying them as insulation for spacecrafts.

In our experimental works we have observed the surface temperature of graphite samples 5x5x25 mm even at 5 Mach in air plasmotron (laboratory device) in the ranges of 4000-4500 oC with the permanent rate of ablation of about 1 mm/s. In re-entry conditions for anti-ballistic missiles the velosity can be more than 20 Mach.

It seems that UHTCs are hot structure ceramics. It is not realistic to use this kind of materials as large areal thermal protection. Comparing with ceramic tiles used on space shuttle orbiter, UHTCs are not weight efficient.

Active cooling, eg. tranpiration cooling is a good idea, but as Dr. Peter said, weight efficiency is a important factor to consider. I don't know whether transpiration cooling agents can really transpire in reentry conditions, as a possitive atmosphere pressure is there. 

Be sure, the proposed re-entry conditions for "shuttles" and "anti-ballistic missiles" are very different. So, the approaches to the thermal protection systems for these two types of high-velosity spacecrafts are completely different as well... However, you don't emphasize what is your design purpose!

Professor Shabalin, could you please talk more about  "anti-ballistic missiles" thermal protection system, and thermal destruction materials as well? I need to learn more basics on this type of TPS materials. Thank you very much. 

Sheng,

"I don't know whether transpiration cooling agents can really transpire in reentry conditions, as a possitive atmosphere pressure is there."

The pressure after the stagnation shock for a typical STS entry profile is well under 1bar (I recall numbers around 15kPa). It is not difficult to imagine a pumped system for transpiration cooling.

Dear Sheng, I think you can use the book by Polezhaev and Frolov (please see the file attached above). The translation problems could be solved now easily on the Google web-site: https://translate.google.com/#ru/en/%D1%80%D0%B0%D0%BA%D0%B5%D1%82%D1%8B

For general introduction to the problem: http://en.wikipedia.org/wiki/Anti-ballistic_missile

Dear Sheng, How do you evaluate the book I have attached to the previous message?

Dear Professor Shabalin,

The book is written in Russian. It is hard for me to read. I studied your journal articles on Heteromodulus Ceramics. I wish to discuss with you about HMC. As you think it is prospective to use it on anti-ballistic missiles, then do you mean use HMC as leading edge thermal protection structures? If so, what is the advantages comparing with advanced C-f/C, SiC/SiC-f  hot structures and UHTCs like ZrB2/SiC-f/Graphite ?  I think the mechanical strength of HMC is moderately low, and the ablation rate would be higher than SIC-f/SiC type hot structure.  Though it is easier to process, it may be have a lower safety factor.

Very briefly:

1. There is no question on 3D- or 4D-C-f carcase, it should be in any ultra-high temperature thermal protection system. So, the main problem is connected with the correct choice of matrix composition.

2. Carbon matrix in C-f/C materials is absolutely poor against thermo-chemical action of low-temperature plasma.

3. The melting temperature of SiO2 formed on the surface of any SiC materials is too low. At higher temperatures its viscosity is neglegible, so the melt is carried away by the oncoming high-speed flow.

4. The presence of low-melting-point B2O3 in the products of oxidation ZrB2 also decreases the reasonable exploitation temperatures of this matrix. The oxidation of ZrB2 is going through the formation of monoclinic ZrO2, which are absolutely poor against thermal shock action, so the pre-oxidizing procedure in this case is not effective at all.

5. ZrC- or HfC-matrices could be pre-oxidized without any damage for nD-C-f carcase and converted (as a result of special heat treatment based on ridge-effect phenomenon knowledge) in the surface layers into thermal shock resitant oxide scales stabilized by residual (secondary in relation to the carcase) carbon and adherent to the substrate (initial matrix).

6. The common increase in strength for ZrC-C or HfC-C materials with temperature growth is much higher than for other materials considered above, so at high and ultra-high temperatures these HMCs have the strength characteristics, which are at least the same, but exactly not worse than inherent to other materials.

Additionally:

1. For details please see my publications and presentations available on my ResearchGate account/page.

2. For translation from Russian into English please use: https://translate.google.com/#ru/en/%D1%80%D0%B0%D1%81%D0%BF%D0%BB%D0%B0%D0%B2%20%D1%83%D0%BD%D0%BE%D1%81%D0%B8%D1%82%D1%81%D1%8F%20%D0%BD%D0%B0%D0%B1%D0%B5%D0%B3%D0%B0%D1%8E%D1%89%D0%B8%D0%BC%20%D0%B2%D1%8B%D1%81%D0%BE%D0%BA%D0%BE%D1%81%D0%BA%D0%BE%D1%80%D0%BE%D1%81%D1%82%D0%BD%D1%8B%D0%BC%20%D0%BF%D0%BE%D1%82%D0%BE%D0%BA%D0%BE%D0%BC (with an example of translation)

Dear Sheng

I would like to remind you that my comments above are concerning a 15-35 M flight in dense atmosphere layers.

Thank you, Professor Shabalin. Do you have the thermal conductivity data for a typical HMC? (For example, ZrC/C or HfC/C.)

If it is used as a leading edge or nose cap hot structure (for example 20mm thick), at a Mach 20 flight in thick air, then how much would be the temperature at the inner face(cold face) of the hot structure? Is  the insulation system for the intemediate temperature range(between the cold face of HMC and the metal mechanical structure of the missile) required? Particularly, the coupling method for hot structure HMC dense ceramic material and inner layer fibrous porous ceramic insulation material, is a hard problem for engineers to handle, because of different physical and chemical properties between the two material system.

We have to consider the thermal conductivity of not (!) MeC matrix, but the similar parameters of FGM (functionally graded materials) containing 2 main layers: MeO2Cx (nanoporous oxide scale with the structure stabilized by carbon) and MeCyOz (intermediate phases providing adherence between outer oxide scale and initial carbide substrate), also taking into account the counter-directed (against main heat flux) flow of released CO/CO2 gases through nanoporous outer scale. You can tailor this 2-layered structure by preliminary modeling/calculation and subsequent pre-oxidized procedure (based on the knowledge of ridge effect phenomenon). Certainly, I am considering some kind of ablative system with the operational life from several seconds to couple minutes, although the most important parameters in this case should be thermal shock resistance criteria as a heating rate can be higher than 104 K/s. This is a completely different problem compared with those, which are related to shuttle-like apparatuses, where any brake engines could be applied easily.

If you are interested in thermal conductivity and thermal expansion of HMC, please see my publications available in ResearchGate.