Figure 1. OREX thermal-protection concept (courtesy of NASDA)
The dissipation of the kinetic energy of atmospheric entry or from vehicles flying hypersonically results in significant aero-thermodynamic heating of all external vehicle surfaces. Typically, the maximum heating rates of vehicles ascending or returning from low Earth orbits remain below 1 MW/m ² , whereas for lunar return or some planetary missions such levels are greatly exceeded.
To maintain appropriate temperatures for equipment, payloads and structures in this hostile thermal environment, an efficient Thermal Protection System (TPS) is required. In practice, two fundamentally different concepts are applied:
This article focuses on high-temperature insulations used in re-radiative TPS concepts, and reports on their current development status in Europe.
Re-radiative thermal-protection systems are primarily applied on reusable vehicles. Depending on the expected temperature levels, two basic material systems are used for the TPS of the US Space Shuttle Orbiter, as well as for Russia's 'Buran" spaceplane:
The rigid tiles are made from high-purity (99.8%) silica fibres, which are rigidly interconnected in a high-temperature sintering process. Tiles with densities of 144 and 352 kg/m ³ have been realised. The exposed tile surface is covered with a thin glassy coating, which is black for the space vehicle's hottest areas and white elsewhere. This coating extends to tile surfaces in the gap between the tiles, thereby leaving only a small tile area uncovered. The tiles are bonded via a layer of nylon felt to the aluminium airframe with a silicone adhesive. The nylon felt acts as an elastic strain-insulation pad.
The flexible blankets (FRSI) were originally made from a Nomex felt. A white silicone elastomer coating provided the required thermo-optical properties and the water-proofing. The blankets were bonded directly to the aluminium airframe with a silicone-based adhesive.
More than 60 successful Orbiter flights and the Buran flight have validated this thermal-protection concept. The RSI has evolved during its use on the Shuttle. Firstly, rigid ceramic tiles with greater mechanical strength and higher application temperatures were developed, by adding alumino-boro-silicate fibres to the silica fibres before the sintering step. Secondly, advanced flexible blankets (AFRSI) have been realised by replacing the Nomex felt by a silica felt. The thermal endurance was significantly enhanced in this way, from ³70 to some 650° C. The AFRSI blankets replaced the original rigid tiles for large leeward surface areas of the Orbiter, resulting both in reduced fabrication/ installation time and costs, and lower TPS weight.
The recent Japanese OREX-capsule flight also used re-radiative thermal-protection tech-niques, with rigid ceramic tiles for the front shield, C/C for the nose cap, and C/C shingles covering a high-temperature ceramic-fibre insulation elsewhere (Fig. 1).
One of the major goals of the OREX flight was the experimental validation of TPS materials and constructions for later application on the planned Japanese spaceplane. However, as a spin-off, it demonstrated in addition that re-radiative TPS could be applied on capsules returning from low Earth orbit. Such alternatives to ablative solutions were proposed by NASA/JSC for the Assured Crew Return Vehicle (ACRV), and are also being studied by ESA.
Another benefit of the OREX flight was its feasibility demonstration of an advanced RSI concept with C/C shingles covering a high-temperature insulation. Such an approach, but with C/SiC shingles and ultra-lightweight insulation packages, was investigated by analyses and ground testing in the framework of ESA's Hermes development programme. Known as ` Rigid External Insulation' (REI), it was considered for the European spaceplane's lower surface and for the cabin region (Fig. 2).
During trade-off and development studies, it was found that high-temperature multilayer insulations - also known as Internal Multiscreen Insulations (IMI) - provide the most thermally efficient high-temperature insulation package beneath the C/SiC shingles.
Figure 1. OREX thermal-protection concept (courtesy of NASDA)
Figure 2. Hermes' thermal-protection architecture
IMI-type insulations promise an evolutionary improvement in the high temperature insulation package between C/SiC shingles (or other hard covers) and the aluminium primary structure. The reason is that heat transfer by radiation becomes dominant in low-density ceramic-fibre insulations above some 600° C. Therefore the implementation of reflective screens (Fig. 3) produces a strong reduction in heat radiation. The IMI concept relies accordingly on a core of stacked reflective foils separated by low-density ceramic-fibre fleeces. For handling purposes, the core might be enclosed in a bag, which consists of ceramic fabric if the IMI cannot be integrated directly with the shingle.
The Hermes-driven high-temperature insula-tion development effort also included use of the Flexible External Insulation (FEI) concept for large areas of the spaceplane's leeward surface. FEI thermal protection is an assembly of quilted blankets that are bonded to the substructure of the spaceplane or capsule by a silicone adhesive. As shown in Figure 4, each individual blanket is composed of a silica-fibre-fleece insulation core, embedded between outer silica-/ABS- and inner glass-fabrics and fixed by a square sewing pattern. Silica-, ABS- or glass-threads are used to realise this sewing pattern.
The FEI is quite similar to the AFRSI which has already performed well on the leeward surfaces of the Shuttle Orbiter and on Buran. Several US and European studies identified that flexible external insulations could also be applied to the rear sides of expendable capsules.
In the following, the development status of both high-temperature insulations - FEI and IMI - is described in more detail.
Figure 3. High-temperature multilayer insulation concept
Figure 4. Concept of a flexible microfibre insulation blanket
The expected temperature range to be endured provides the first criteria for the selection of materials/components (Fig. 5). Products made from S-glass can be applied up to 380° C. In the temperature range between 380 and 800° C, silica fibres, threads and fabrics are preferred due to their favourable low density and low thermal conductivity. For temperatures beyond 800° C or in contaminated environments, other ceramic materials are a better choice. As a general rule, the temperature durability increases with the alumina content of the fibre, thread or fabric.
Extensive characterisation testing for threads and fabrics has demonstrated the superiority in strength terms of ABS material for temperatures above some 500° C. However, their thermal properties are inferior to those of silica products, and consequently it was confirmed that silica-based products provide the best compromise up to some 700° C under moderate mechanical loads.
The thermal-insulation function is dominated by the microfibre insulation core. The total heat transfer in this core consists of heat radiation and conduction both by the enclosed gas and by the fibres. However, for the core densities envisaged (80-120 kg/m ³), the conduction via the fibres is almost negligible. The core's resistance to radiative heat transfer depends on fibre diameter, material and orientation. Silica fibres of roughly 2 microns diameter would provide maximum resistance (Fig. 6). In practice, silica fibres in the 2-3 micron and around 9 micron size range are easily available from suppliers.
Figure 5. Material/component selection guide
Figure 6. Mass-related resistance to heat radiation as a function of the diameter of the fibres used for the blanket's core
For capsule applications, alumina-based fibres could become necessary. Such fibres are available with diameters of 3 to 5 microns.
In Figure 7, the theoretical contribution of the gas conduction in different fibre fleeces is compared with experimental values at different gas pressures. Generally, it falls dramatically with reduced pressure. Again, 2 micron fibres are found to be thermally more effective.
As the FEI forms the outer surface of the re-entry vehicle, it is exposed to dynamic forces from static pressure, aero-acoustics and aero-elastic loads (Fig. 8). The bobbin threads and the outer fabric are obviously the most heavily loaded elements in this application.
In view of the fatigue behaviour, it is considered favourable to keep the bobbin thread under tension load at all times. This is achieved by compressing the insulation core to the specified density during the manufacturing process. The resulting compression stiffness of the core and the membrane stiffness of the outer fabric are the major parameters affecting the response to dynamic excitations.
During FEI blanket development, problems with fatigue under high acoustic loading (OASPL 168 dB) had to be overcome. The first blankets failed under long-duration acoustic loading caused by breakage of the sewing threads. Selection of more suitable ABS materials (Fig. 9) for blankets exposed to the highest mechanical loads has allowed them to sustain these extreme loading for more than 15 min, which exceeds the envisaged cumulative acoustic exposure time for the spaceplane's design life of 30 flights.
The low emissivity of the outer fabric would result in unfavourable external surface temperatures during re-entry.Therefore, a thin high-emissivity, low-catalycity coating is applied, which also improves the fabric's erosion resistance.
By employing all of the material components and developments mentioned above, medium-sized blankets with good performances can now be manufactured in varying thicknesses by DASA in Germany (Fig. 10).
Returning to the design concept presented in Figure 3, silica- or alumina-based fibres can be chosen for the spacer fleece between the reflective foils depending on the temperature range to be endured. The C/SiC shingles provide mechanical protection against the airflow. Consequently, one can choose fleeces with a much lower density for the IMI than for the FEI.
Figure 7. Gas conductance as a function of pressure
Figure 8. Schematic of the structural load system
Figure 9. Acoustic performance as a function of component material (time to first failure at 168 dB)
The thermal-improvement potential of IMI-type products is illustrated in Figure 11, where the radiative `conductivities' of an IMI-type insulation (density around 35 kg/m ³ ) and a conventional high-performance fibre felt (100 kg/m ³ ) are compared, with the thermal conductivity of air as a reference. Above some 600° C, the thermal conductivity of the IMI-type insulations is lower than those of conventional fibre felts. In addition, the IMI-type insulation's mean density is much lower. Thus the product /\.r, commonly taken as a measure of an insulation's thermal efficiency, is lower over the whole temperature range.
IMI-type insulations can be tailored for a wide range of transient re-entry load cases by judicious choice of the spacer density and number of screens. Quite a high number of reflective screens is required for low-density spacer felts, which implies a need for lightweight screens. These could take the form of noble-metal-coated ceramic substrate foils, made from short-alumina-fibre reinforced alumina.
Figure 10. Flexible External Insulation (FEI) Blanket Sample
Figure 11. Conductivities of felt- and IMI-type insulations as a function of temperature
Figure 12. The ceramic substrate foil after firing
Attention has to be paid during foil manufacture to the development of procedures that minimise surface micro-cracking. Foil masses range between 30 and 45 g/m ³ , depending on the surface-quality requirements. The impres-sive flexibility of the finished product greatly facilitates the assembly of the IMI package.
Gold and platinum were selected as the base materials for the reflective coatings. To stabilise the noble-metal films against agglomeration at high temperatures, diffusion-blocking coatings are applied. Gold is totally resistant to high-temperature oxidation and very low emissivities have been measured (Fig. 13), but its applicability is limited to about 1000° C. Platinum coatings show slight oxidation at the highest temperatures but, due to the volatility of the oxides, no surface contamination occurs. Their maximum utilisation temperature is therefore around 1450° C for long-term applications, and even higher for short exposure times.
Compressed low-density spacers tend to exhibit strong creep behaviour at high temperatures. Procedures have been developed in Europe that significantly reduce this effect (Fig. 14), which would otherwise lead to plastic deformation of the IMI after just a few reentry flights. Mechanical testing has demonstrated the very favourable damping behaviour of the IMI and shingle/IMI systems when exposed to acoustic loads.
All in all, the manufacture of IMI packages has been well-mastered in Europe, by MAN. Figure 15 shows the various elements in such a modern sample of high-temperature multilayer insulation.
High-temperature insulation for space vehicles has made enormous strides over the last few years thanks in large part to the efforts of European industry. Both FEI- and IMI-type insulations have been extensively ground-tested and can thereby be considered ` pre-qualified' for a spaceplane-type application and are reusable for several missions.
Both types of insulation have been ` mastered' to the extent that they can now be tailored to different hypersonic flight environments. Enhancement of FEI insulations for still-higher temperature applications and their potential use for capsules is presently being investigated in the framework of ESA's MSTP technology programme. NASA is currently investigating the potential of IMI insulations for its Delta-Clipper re-usable launcher.
Figure 13. Effective emissivities of gold- and platinum-coated screens (measured at ZAE in Würtzburg, Germany)
Figure 14. Spacer resilience as a function of mechanical loading before and after manufacturing improvement
Figure 15. High-temperature multilayer insulation demonstration sample
ESA Bulletin Nr. 80.