Heat Flux Reduction by Transpiration-Cooling of CMCs for Space Applications

Space flight requires high performance materials with high strength to density ratio over a wide temperature range. The definition of high temperature, however, depends on the application. In terms of satellite systems, for example, high temperature can be anything above 450 K, while this is perceived as a low temperature in rocket engine technology or structural response at hypersonic flight. Ground based high temperature applications, include firefighting, where the thermal emergency conditions exceed 600 K, power plants and incinerators operating at up to 1100 K, and modern gas turbines which are run at temperatures of up to 2500 K for increased efficiency and reduced emissions. Extreme environments are also faced during high-speed flight or in engines and rockets. The wall temperatures currently endured are roughly 3500 K in the hot combustion zone and a few hundred Kelvin in the cooled exterior of the chamber wall and approximately 2000 K on short-term exposed vehicle leading edges. To shield hypersonic vehicles ranging from atmospheric re-entry vehicles to hypersonic cruise vehicles from extreme heat loads, thermal protection systems (TPS) are required. However, access to space is extremely expensive so that overdesign of the TPS must be avoided. Besides being high temperature capable, material and more specifically the hot structure needs to be damage tolerant and provide stability versus combined loads over a large number of load cycles. The critical components include leading edges, control surfaces, acreage TPS, and seals, as illustrated in Figure 10.1. Both, atmospheric re-entry vehicles and hypersonic cruise vehicles need a thermal protection system. Figure 10.2 shows what happens at atmospheric re-entry without a heat shield. ATV (Automated Transfer Vehicle), the five European expendable supply spacecraft have each been brought to a destructive re-entry in the atmosphere on purpose. As the density increased upon re-entry, the heat load reached the maximum temperatures the materials could withstand and it separated into three major fragments. It is supposed that the equipment bay first collapsed, followed by the propellant bay and lastly the cargo bay. The part that survived the longest time was the Russian docking adapter, a massive titanium component.

Space flight requires high performance materials with high strength to density ratio over a wide temperature range.The definition of high temperature, however, depends on the application.In terms of satellite systems, for example, high temperature can be anything above 450 K, while this is perceived as a low temperature in rocket engine technology or structural response at hypersonic flight.Ground based high temperature applications, include firefighting, where the thermal emergency conditions exceed 600 K, power plants and incinerators operating at up to 1100 K, and modern gas turbines which are run at temperatures of up to 2500 K for increased efficiency and reduced emissions.Extreme environments are also faced during high-speed flight or in engines and rockets.The wall temperatures currently endured are roughly 3500 K in the hot combustion zone and a few hundred Kelvin in the cooled exterior of the chamber wall and approximately 2000 K on short-term exposed vehicle leading edges.
To shield hypersonic vehicles ranging from atmospheric re-entry vehicles to hypersonic cruise vehicles from extreme heat loads, thermal protection systems (TPS) are required.However, access to space is extremely expensive so that overdesign of the TPS must be avoided.Besides being high temperature capable, material and more specifically the hot structure needs to be damage tolerant and provide stability versus combined loads over a large number of load cycles.The critical components include leading edges, control surfaces, acreage TPS, and seals, as illustrated in Figure 10.1.
Both, atmospheric re-entry vehicles and hypersonic cruise vehicles need a thermal protection system.Figure 10.2 shows what happens at atmospheric re-entry without a heat shield.ATV (Automated Transfer Vehicle), the five European expendable supply spacecraft have each been brought to a destructive re-entry in the atmosphere on purpose.As the density increased upon re-entry, the heat load reached the maximum temperatures the materials could withstand and it separated into three major fragments.It is supposed that the equipment bay first collapsed, followed by the propellant bay and lastly the cargo bay.The part that survived the longest time was the Russian docking adapter, a massive titanium component.For the choice of the TPS high temperature material, ceramic matrix composites (CMCs) are becoming increasingly important for high temperature applications.New production methods for processing carbon fibers use innovative machines which offer a wide range of lay-up possibilities as well as process development of infiltration methods or of metal-ceramic-joining enable increased use of these composites.In terms of space propulsion, for example, nozzle extensions are being developed and tested (2010), such as the Vinci extendible nozzle by Snecma in France (Safran Group).Airbus Defence & Space has also been developing reinforced silicon carbide nozzle extensions for the ARIANE 5 first and upper stage motors as well as ceramic combustion chambers of smaller motors (Schmidt et al., 2004).The European missile company MBDA (France) and Airbus Defence & Space and their research arm Airbus Group Innovations are developing a fuel-cooled CMC combustion chamber for dualmode ramjets (PTAH-SOCAR).It makes use of woven carbon fibre to produce the shape of the structure wanted with subsequent densification and, as fibrous structures are permeable, a sealant application process is used to seal the pre-form.
Depending on the boundary conditions of a hypersonic mission, the heat load may still exceed even a CMC's material limitation and passive or active cooling approaches can be favorable.A semi-passive cooling concept, for example, utilizes heat pipes.Heat pipes are a standard device in satellite thermal control and can provide cooling of stagnation regions by transferring heat to locations aft of the stagnation region.Heat-pipe-cooled wing leading edges were studied in Space Shuttle leading edge concepts in the 1970s and during the U.S. National Aerospace Plane (NASP) program in the 1980s (Glass et al., 1999).The leading edges had a reinforced carbon carbon structure with Mo-Re heat pipes as the container, shown in Figure 10.3 a).Material compatibility and thermal stresses are two of the major challenges with heat pipes embedded in composite materials (Glass et al., 1999).Active cooling systems are of special interest for use in severe thermal environments where the passive systems are inadequate.Metallic actively-cooled leading edges have also been fabricated, and the challenges include hydrogen interaction and multi-cycle life.A photograph of the internal portion of an actively cooled leading edge proposition is shown in Figure 10.3 b) (Glass et al., 1999).Other active cooling systems include film cooling with slit, showerhead, effusion, and transpiration configuration.Deutsches Zentrum für Luft-und Raumfahrt (DLR, German Aerospace Center) has been investigating CMC rocket motors since the 1990s (Hald et al., 2005;Ortelt et al., 2013).Figure 10.4 shows the ceramic thrust chamber during a demonstration test.The experience with this technology has led to the transpiration-cooling of other hot structures, such as thermal protection systems (TPS) for atmospheric re-entry manoeuvres.Consequently, with AKTiV (Aktive Kühlung durch Transpiration im Versuch), the first transpiration cooling flight experiment was flown on SHEFEX II in a sub-orbital re-entry mission in 2012 (Böhrk, 2014).The subsequent project SHEFEX III is ongoing and a launch is planned for 2018.A cooled leading edge is required for SHEFEX III.The approach investigated at DLR focuses on the application of transpiration cooling using permeable CMC material.Figure 10.5 shows a sketch in which surround flow and transpiration flow of a leading edge are indicated.This chapter gives an overview of the application of novel permeable composites for the transpiration-cooling of space-related high-temperature structures.Section 1.1 gives an overview of the loads on hypersonic vehicles, putting into relation the two typical hypersonic flight situations atmospheric re-entry flight and hypersonic cruise flight, stressing the immense heat load on sharp leading edges.The mission of SHEFEX II is introduced, providing the platform for AKTiV, the first experiment to prove transpiration-cooling in hypersonic flight.In section 1.2, an approach of determining the thermal structural response of hot structures for their proper layout is introduced.This approach is realized in the program HEATS which serves later to interpret flight data.Section 1.3 gives a brief overview over reinforced ceramic matrix composites often used as structural material for hypersonic applications, such as the SHEFEX II leading edge.Based on this material, section 1.4 gives flight data from AKTiV on SHEFEX II.The data and evaluation shows that AKTiV has successfully demonstrated transpiration-cooling during atmospheric re-entry.The next step is thus to use the gained knowledge for leading edges subject to severe heat flux.This is planned for the leading edge of SHEFEX III.

A Space Application: Hypersonic Flight
Hypersonic velocities are encountered during atmospheric re-entry flight or during sustained hypersonic flight.Both are connected with flight in the atmosphere, where the large velocity of the vehicle causes a shock or a bow shock wave ahead of the vehicle.Through this shock, the atmospheric gas is compressed and the kinetic energy is dissipated into internal energy of the atmospheric gas.In the case of an Earth re-entry, an air plasma state is formed in the surrounding gas which is characterized by high temperatures and dissociated and partly ionized particles.Temperatures can reach >20,000 K with a compression ratio of several hundred.
Hypersonic velocity is generally defined as velocities of Mach 5 or greater and implies effects such as narrow shocks, with shock distances so small that the shock layer may interact with the boundary layer, or high temperature effects in which the atmospheric gas is decomposed by dissociation of the molecules and ionization of flow particles.It was found in the 1950s by H.J. Allen that the shock distance δ to the vehicle surface increases with the nose radius which drastically decreases heat load.Oertel Jr. (1994) gives the correlation δ=R LE 0.143e 3.24/M 2 (10.1) for hypersonic flight with Mach number M and the leading edge radius R LE .A few vehicles having performed successful re-entry are shown in Figure 10.6.Typical entry velocities v range from 7.9 km/s for an Earth orbital re-entry and over 11.3 km/s for a return mission from another celestial body.The kinetic enthalpy amounts to approximately 31 MJ/kg, and 63 MJ/kg, respectively.Re-entry vehicles are, in principle, deceleration systems, the most prominent one being the Space Shuttle (Hirschel and Zarchan, 2006).Therefore, all of these vehicles are blunt in design in order to achieve large drag and a large shock stand-off distance ahead of the vehicle.In the distance between the shock layer and the vehicle surface, the energy is dissipated into internal degrees of freedom of the atmospheric gas, i.e. dissociation, vibration and rotation, causing a reduction of the sensible temperature of the gas for the vehicle surface.Among the re-entry vehicles, ballistic and lifting vehicles are differentiated, comprising ballistic and lifted trajectory strategies, respectively.The challenge of the atmospheric re-entry is decreasing the vehicle's velocity from 7.9 km/s (28,000 km/hr) in flight to 0 km/s on ground only by deceleration in terms of drag and friction.
This can be achieved by a ballistic re-entry trajectory, in which the vehicle falls at high velocity back into the atmosphere and thus encounters strong deceleration loads and a high heat flux onto the vehicle's protective material over a short period of time.Deceleration can also be achieved by a lifted re-entry, in which the strategy is to decrease velocity at a higher altitude, where the atmospheric density is still relatively low, and thus encounter only a reduced amount of heat flux.In Figure 10.7, both re-entry strategies are sketched.However, when staying at elevated altitude, the mission time is increased and the vehicle has to withstand the reduced heat flux over a longer period of time.Figure 10.8 shows the heat flux over time for the EXPRESS capsule mission, launched in 1995, in comparison to the heat flux for winged re-entry body X-38, a project cancelled in 2002 two years short of completing its flight test phase.The lifting body was designed to lower heat load but over a longer period of time.Capsule-type re-entry vehicles typically have an axis-symmetric shape and fly at a negative angle of attack, resulting in weak aerodynamic performance, the measure of which is the lift-to drag ratio L/D.This parameter governs mission capabilities, such as down and cross range.Figure 10.9 shows cross range capabilities of hypersonic vehicles.
Wingless re-entry vehicles generate aerodynamic lift from the shape of their bodies.With research missions of the X-24A in the 1960s, it was demonstrated that hypersonic vehicles like the Space Shuttle could land on conventional runways by only gliding.The X-24B was intended to test lifting body concepts at very high lift-to-drag ratio.Lifting re-entry bodies have flat lower sides and fly at large positive angles of attack, resulting in an aerodynamic performance L/D between 0.5 and up to and above 2.0.
Figure 10.9: Cross range capabilities for hypersonic lift-over-drag ratio for various hypersonic flight vehicles.
As opposed to re-entry vehicles, hypersonic cruise vehicles must, like aircraft, fly with efficient propellant consumption.These vehicles are currently operational.The technology however is evolving.Hypersonic cruise vehicles have to have minimum aerodynamic drag in order to be propellant efficient, which for hypersonic flight necessitates a slender configuration.Equation (1.1) shows that a shock stands ahead of blunt shapes but may be attached to pointed shapes, showing that leading edges of hypersonic cruise vehicles see very high heat load.
A formulation developed by Marvin and Pope is widely used for determining heat flux (Marvin and Pope, 1968) (10.3) for the stagnation region of a hypersonic vehicle with flight enthalpy h, efficient leading edge radius R eff , total pressure p tot and a gas specific constant K, which is 0.368 kW (MJm) -1 (mPa) -½ in the case of air.The heat flux to the leading edge is dependent on the leading edge radius proportional to a root function, as depicted in Figure 10.10.From an aerothermodynamic point of view, sharp leading edges can cause problems.They promote local stagnation areas with very high temperatures which materials previously have not withstood.During the 1990s, the development of ceramic composites and ultra high temperature materials for TPS applications led to a renewed interest in sharp edged configurations (Longo et al., 2005) such as the waverider concept DLR F8 (Strohmeyer et al., 1998), the DLR-ONERA project JAPHAR (Eggers et al., 2001) or the lifting body concept HL-20 and the SHARP project, both from NASA (Reuther et al., 2001;Arnold et al., 2001).Adding to these issues is the slender aerodynamic shape of minimum drag which provides thin cross-sections at elevated temperatures.One of the primary thermomechanical challenges results from large thermal gradients, for example at attachments of hot structures to an airframe, as was described in the former chapter.
The DLR program Sharp Edge Flight Experiment (SHEFEX) comprises a series of platforms for re-entry experiments.The faceted concept of SHEFEX intends to reduce the design and manufacturing cost of the CMC thermal protection system (Weihs et al., 2008(Weihs et al., , 2002)).Figure 10.11 shows the geometry of SHEFEX II with a forebody of ogive form and octagonal cross-section.The forebody is symmetrically divided into eight identical facets 1 through 8 in circular direction in all five segments A through E, as marked in Figure 10.13.It performed sub-orbital re-entry from a 177 km apogee.The total flight time was roughly 500 s comprising 52 s of experimental time for the atmospheric re-entry between 100 and 30 km.Within the experiment time, the vehicle accelerated from 2559 m/s and Mach 10.2 at 101 km altitude to 2791 m/s and Mach 9.3 at 30 km.The trajctory is shown in Figure 10.12.(Turner et al., 2013).
The 1.5-m-long forebody of SHEFEX II has a sharply pointed nose tip made from solid C/C-SiC, the vehicle's leading edge.As discussed above, sharp leading edges are subject to severe aerodynamic heating since the shock attaches to the pointed structure.The challenge for the designer is thus the determination of the thermomechanical load onto the structure, the choice of the cooling method, material choice, attachment, and instrumentation.In case of SHEFEX II, it is sufficient for the nose and canard to be radiation cooled and a recession of up to a few millimeters is tolerated.
One key SHEFEX II experiment was the AKTiV transpiration-cooled experiment.The experiment is also visible in Figure 10.13 with a black permeable sample, while in Figure 10.11 the sample appears lighter than the surrounding panel.The next section will introduce a heat balance approach to the structural thermal response while the material choice for AKTiV is explained in the subsequent section.

Hot and Cooled Structure Thermal Response
In order to design a thermal protection system, the thermal response of the material to the load has to be known.Depending on the mission, transient wall temperature determination is essential in order to optimize TPS mass.However, for horizontal hypersonic cruise flight at likely invariant flight and atmospheric conditions, a steady state of the TPS is approached.
The heat equation in two dimensions, with density ϱ, heat capacity c, temperature T and thermal conductivity λ in x and y direction (10.4) can serve as an approach to describe the temperature development at any point within the TPS material.In the case of a steady-state problem, . The boundary condition is the heat balance at the surface and a boundary condition of second order can, for example, be used at the rear side and the edges of the TPS-panels (10.5) i.e. assuming that these sides are adiabatic.Figure 10.14 a) illustrates a capacitively cooled thermal protection system that manages the heat by a thick wall of material that has high heat capacity.This is used for low heat fluxes and relatively short mission times.Radiation of the vehicle surface can be modelled by a gray body assumption, q • rad =σεT 4  (10.6) with Stefan-Boltzmann-constant σ and the wall material's emissivity ε.However, for capacitive cooling, surface temperatures are usually low and thus, radiative heat flux on these structures is marginal.If a heat sink structure is heated for a long period of time, enough heat could be absorbed to overheat the structure.Figure 10.14 b) shows a radiatively cooled system which is used for example in the EXPERT (European Experimental Re-Entry Testbed) nose cap or the SHEFEX TPS.A view of the SHEFEX TPS is given in Figure 10.15 (Böhrk et al., 2011).Thermal radiation, as a mechanism to remove the heat, becomes more efficient with higher temperature.The emissivity has strong influence on the surface radiation cooling.Figure 10.16 shows the radiative heat flux emitted away from the vehicle surface for some vehicles according to Equation (10.6).In contrast to a heat sink structure, hot structures with significant radiation cooling can be used for a higher heat load, allowing the structure to reach steady state conditions.Again, the heat is both radiated away and conducted inward.The objective of the insulation is to minimize the amount of heat reaching the structure.Figure 10.14 indicates that a small amount of heat is conducted through the insulation to the structure.Boundary condition (1.5) now applies at the rear of the insulation.
The heat balance for the surface in the case of the capacitively and radiatively cooled systems assumes that all incoming heat flux q • HG−S from the hot gas (HG) to the surface (S) is either conducted into the structure, q • cond , or radiated away, q • rad , so that (Böhrk et al., 2010) (10.7) Heat flux from the flow to the wall can be convective heat flux, radiation, or chemical effects such as catalysis.In the present approach these effects are neglected and only convective heat flux is looked at.The thermal energy input rate  (Böhrk et al., 2010;van Driest, 1952van Driest, , 1956) ) The boundary condition at the surface S is (10.11) If a hot structure has to withstand an extreme heat load for long times, passive radiation or heat sink cooling is not sufficient anymore.The X-15, for example, used Inconel as a heat sink for thermal protection, and suffered skin penetration due to high temperatures generated from shock-shock interactions (Glass, 2008).One possible solution to reduce temperatures at the surface is the backface cooling of the hot structure.This can even be realized in a semi-passive thermal protection system using a heat pipe.Heat is transferred by a working fluid to another region of the heat pipe where the heat is radiated away.Figure 10.17 shows the heat transfer mechanisms for the heat pipe.All processes remain the same as with radiative or capacitive cooling, but the boundary condition between hot material and working fluid is now with index rear−C indicating convective heat transfer from the wall to the coolant.Heat flux q • rear ̵ C can, again, be estimated as Stρ c c p,c u c (T rear -T r,c ).Active cooling is required for even higher heat fluxes and for longer mission times.
Convective cooling of the hot surface, in which the coolant heats up and carries the heat away, is also utilized for a high heat flux and long times.The structure operates hot but is maintained within its structural limits with respect to temperature by the active cooling.Two kinds of convective cooling, film and transpiration-cooling, are shown in Figure 10.18., for example the surrounding atmospheric gas, heats up the film which in turn heats the wall over a length Δl.At the same time, cool flow with flow enthalpy h in is flowing in from upstream and is heated up before it exits the film of film thickness δ F on the downstream side with increased flow enthalpy h out Turbulent flow is characterized by a strong vorticity, leading to a mix between the two gas flows, hot and film, reducing the cooling efficiency.In the heat balance of the coolant film as of eq.(1.14), the term of convection between hot gas flow and film can be replaced by a term accounting for the entraining hot gas mass into the film and the subsequent heat adjustment between the fluids (Böhrk et al., 2010).Equation (10.14) becomes + ((ρu) in δ in + ṁ) h out (10.15) with growing film thickness δ F .The factor ṁ describes the mass of air entering into the film (Goldstein, 1971).
In transpiration-cooling, a coolant is fed through a porous and permeable wall structure to the outer mold line of a hot structure and there it forms a film.In contrast to the discrete locations used in film-cooling (a slit or hole), this is done over large areas.The structure is cooled by both convection, as the coolant passes through the pores, and by the film that has formed on the hot-gas side.Thus, transpiration-cooling is caused by two physical phenomena: the porous structure being convection-cooled by the coolant and the coolant layer on the outer -hot -surface lowering the heat transfer from the high-enthalpy environment to the vehicle surface.A coolant film may also provide oxidation protection for the structure, which may be an important issue in case of carbon-based materials.
The heat balance on the hot side, according to eqs.(10.14) and (10.15), remains for both laminar or turbulent flow conditions.The cooling of the inner structure of the porous material is carried out by convective heat transfer between the material and the coolant.Radiation within the porous structure is negligible.According to Glass, the heat equation (10.4) for the wall material depending on coolant temperature T c is extended to (Glass et al., 2001) (10.16)An additional balance for the coolant yields (10.17)When only the pressure gradient in flow direction is taken into account, the above equation yields (10.18) with the volumetric heat transfer coefficient α V , wall temperature T, coolant mass flow rate ṁ and transpiration-cooled area A.
The computer program HEATS (Heat Exchange Analysis for Transpiration Systems) serves to solve the heat balances of re-entry vehicle structures and uses the set of equations given above (Böhrk et al., 2010).HEATS determines transient wall temperature throughout an entire re-entry trajectory at short calculation time.It can thus be used as a lay-out tool for cooled or uncooled structures for both, in-flight and ground testing, so that overdesign can be minimized.This way, the thermal response of the structure is determined for the uncooled case as well as for coupled film-and convection cooled conditions.It has been validated by comparison to experiments in an arc-heated wind tunnel under laminar in-flow (Böhrk et al., 2010, Sep. 2012).HEATS will be referred to in a later section of this chapter when data is shown and evaluated.

Material Choice
As mentioned above, thermal protection systems (TPS) consist of structural parts, hot structural parts, attachment components, insulation and possibly a system infrastructure (such as gas or a liquid reservoir for active cooling).Flight vehicle substructures are commonly frames with good thermal conductivity, so that hot spots are avoided where hot parts are attached.The ideal candidates for the hot side of the TPS are ceramic materials, specifically carbon fiber composites, which are lightweight and have high specific strength and stiffness.They also have high emissivity at relatively low thermal conductivity which is necessary when insulated hot structures are used.The materials of the hot structures should typically have little to no erosion rate, e.g.due to oxidization or nitridation, and low surface catalycity, so that recombination of atomic species at the material surface is suppressed.Finally, they have to have compatible thermal expansion with their substructure, possess a high tolerance for cyclical loading and to be able to form complex shapes.Their purpose in hot structures is to handle the forces and temperatures experienced by leading edges in atmospheric re-entry and sustained hypersonic flight.An example is shown with the SHEFEX II leading edge attachment described in this section (see Fig. 10.22).
Carbon fiber composites have weak fiber-matrix bonding and higher matrix modulus than fiber modulus.This results in improved damage tolerance due to fiber pull-out, as shown in Figure 10.19.Since carbon fiber composites are lightweight and have stable properties even at high temperatures with respect to Young's modulus, strength, stiffness and tensile stretch, they are often applied for thermal protection of outer skins.The fibers increase damage tolerance and tensile strength of the ceramic while still being temperature and wear resistant.They also reduce the thermal expansion of the material.This is useful in structures that are subject to large thermal gradients, so thermal stress does not affect the component greatly.As opposed to monolithic ceramics, it is possible to manufacture ceramic parts that are even thin-walled, large or complexly shaped with the structural properties of these materials.Carbon fiber reinforced silicon carbide production techniques include liquid polymer infiltration (Schmidt et al., 2004), chemical vapor infiltration and liquid silicon infiltration.One example for a ceramic matrix composite, DLR's C/C-SiC, is shown in Figure 10.20.The material is based on carbon fibers with a carbon and silicon carbide matrix.The fiber lay-up is typically a stack of 0/90° planar laminae so that the material is orthotropic and directional.Within the laminae planes and parallel to the fibers are the 0° and 90° direction, while the material can also be oriented at any other angle.It has a relatively low porosity of 3% and a low overall density of 1900 kg/m 3 .The thermal expansion is low, since fiber-dominated, with 2×10 -6 K -1 in laminae direction and higher, since matrix-dominated, with 6×10 -6 K -1 perpendicular to the laminae.The material has high specific strength at high temperatures, is thermo-shock resistant and withstands loads up to temperatures of 1840 K. Since it is made with carbon fibers it has lower oxidation resistance.C/C-SiC has been qualified in plasma wind tunnel testing and during real re-entry flight, such as in the FOTON 9 (Hald, 1994), EXPRESS (Hald, 1994(Hald, , 2001)), FOTON-M2 (Reimer, 2006), and SHEFEX I and II (Weihs, 2013;Gülhan et al., 2006) missions.
The fiber ceramic C/C-SiC is manufactured in house at DLR (Heidenreich et al., 2014) with liquid silicon infiltration.The process is described here as an example for the manufacturing of CMC components.Firstly, a green body is made from carbon fiber reinforced plastic.Secondly, the green body is treated by pyrolysis under a pure nitrogen atmosphere at 1170 K and, thus, the matrix is carbonized to C/C.In the second step, a micro crack pattern is formed within the carbon matrix.This intermediate stage C/C has intriguing permeable properties.The good permeability of the C/C produced during DLR's pyrolysis step, is used in transpiration-cooling in which a cooling fluid is fed through the structure into the boundary layer, as described in section 1.2.The micro-cracks in the low density carbon matrix serve as open pores for the fluid transfer.The pores distribute the coolant evenly over the hot facing surface to keep it cool and enable it to withstand exposure to high temperatures.The permeability of the C/C used for the AKTiV hypersonic in-flight transpiration-cooling experiment in 2012 is on the order of 10 -13 m 2 .The porosity is relatively high at 12%.A coolant film may also provide oxidation protection for the structure.Since material properties are adjustable by choice of rovings and stacking of plies, both conductivity and permeability can be optimized to the values necessary for the application.The third step towards C/C-SiC contains the liquid infiltration of the C/C-body with silicon under vacuum at 1870 K. Silicon carbide is formed, starting at the phase interfaces between silicon and carbon, until the entire matrix is transformed to silicon carbide.Meanwhile the fibres remain unaffected and a content of free silicon of approximately 2% remains.
The octahedral pyramid shape of the SHEFEX II leading edge was first bonded and machined with the carbon-carbon before it was siliconized.The tip radius was determined after siliconization to approximately 1 mm.The tip was attached to the vehicle by a mount from thin-wall C/C-SiC material, as shown in Fig. 1.22.The mount was composed of two in-situ joined half shells and an inlay.The nose was attached to the mount using four fasteners, which were screwed into four screw-bores in the inlay.The thin-wall mount, in turn, was attached to the aluminum substructure by the CMC stand-offs (Böhrk et al., 2014).Three thermocouples were mounted in the center of the tip to measure the material temperature as close as possible to the front, from where the aerodynamic shock rises.The thermocouple beads were located approximately 20 mm along the vehicle axis and 5 mm below the surface.Additionally, pressure is measured on each of the eight sides via slender bore holes leading to the aft side of the leading edge.Figure 10.21 shows the pressure tubing of the tip.10.12) results in the simulated temperature response shown in Figure 10.24 (Böhrk et al., 2014).The data are in good concurrence with the in-flight measurement.The differences between the simulation and the measurement are attributed to uncertainties in the atmospheric model and assumptions of constant material properties.A sensitivity analysis with respect to specific heat has shown that in the temperature regime present in the return flight of SHEFEX II, thermal response is affected by this property by approximately 5%.Moreover, bond thickness between the structure and the thermocouples is unknown.Another carbon fiber composite widely used is reinforced carbon-carbon (RCC), which was also used for the Space Shuttle Orbiter leading edges.It is manufactured by pyrolizing a nylon cloth, converting it to graphite.The graphite is then impregnated with a phenolic resin and cured in an autoclave.The cured material is pyrolized again to transform the cured resin to carbon.The part is then densified by impregnating it with furfural alcohol in a vacuum and pyrolizing a third time for carbon transformation.The densification step is repeated three times until the final RCC part has achieved a density of 1440 to 1600 kg/m 3 (Kasen, 2013;Curry et al., 1983).
An example of an RCC leading-edge is the X-43 Mach 10 vehicle leading edge, which was designed to reach nearly 2450 K during the short 130-second flight (Glass, 2008).The X-43 was a series of scramjet propulsion test vehicles for an unmanned experimental hypersonic aircraft.After one failure in which the vehicle had to be destroyed, two other missions successfully flew in March and November 2004, making it the fastest free flying air-breathing aircraft in the world until beaten by the X-51 in 2010.For the X-43 Mach 10 vehicle leading edge, high thermal conductivity fibers in an unbalanced weave were used help conduct heat away from a leading edge, thus reducing the maximum temperatures.The fibers were woven in a 75/25 unbalanced plain weave to obtain a high thermal conductivity in the chordwise direction.
The primary environmental durability challenge with RCC is oxidation resistance, which has a major impact on mission life.The number of cycles required under combined loads, inspection and repair, and the ability to predict mission life are the present challenges (Glass, 2008).A way to prevent oxidation is therefore to coat the structural parts.Silicon carbide-based coatings increase the oxidation resistivity of a structure.They have a high emissivity and can be used up to 2000 K.This is good for blunt re-entry missions that impose only a short exposure time, such as the EXPERT nose cap.On air-breathing vehicles, however, temperatures exceed the 2000 K due to high dynamic pressure, velocity and exposure time.Coatings can also be ultrahigh temperature ceramics coatings or oxydic.Oxydic coatings are used for resistance to high temperature oxidation and corrosion.Figure 10.26 shows an iridium coated leading edge from (Glass, 2008).When a coating is used, thermal expansion of the coating and the substrate have to be matched to prevent the coating from spalling, resulting in oxidation of the substrate.(Glass, 2008).
Ultra-high temperature ceramics (UHTCs) such as zirconium diboride and hafnium diboride ZrB 2 or HfB 2 , among others, have melting points at temperatures on the order of 3000 K (Wuchina et al., 2007).They have been investigated in the early 1950s as nuclear reactor material.For the time being, oxidation resistance still needs to be improved for reusable or long term TPS leading edge applications, for example by adding around 20% SiC.Since the material is monolithic, it has low thermal shock and fracture toughness and is better applied as a coating onto reinforced material.Extensive research is still ongoing to properly establish material properties in order to validate their use for aero structures including composite approaches with prepregging and infiltration processing to increase mechanical robustness (Levine et al., 2003).Moreover, for fracture toughness, a composite approach is required, the UHTCMC.Current investigations include reactive melt infiltration to manufacture a fibre reinforced ZrB 2 matrix.In this approach; a boron containing fibre preform is infiltrated with a liquid zirconium alloy which forms ZrB 2 on infiltration (Kütemeyer, 2015).Other research focuses on silicon carbide fiber and hafnium diboride filled SiC precursor slurry (Leslie et al., 2014).
Ultra-high temperature materials and their production are, however, still connected with several physical challenges and limitations.Therefore, the approach to encountering extreme temperatures presented here, concentrates on active transpiration-cooling.The cooling is currently realized on the basis of the C/C material previously presented, which represents an intermediate stage in the production process of C/C-SiC, and has compelling permeable properties.

Transpiration Cooling
The previous sections have shown that the outer surface of hypersonic vehicles are subjected to severe aerodynamic heating.It was also shown that especially for sharp edges at the stagnation point, the loads are immense.However, stability of the outer mold line is important since it impacts performance.Sharp leading edges at an airbreathing propulsion system generate shocks which are necessary to maximize airflow into the engines.Therefore, leading edges should not ablate, even when they are expendable parts.Instead, they may have to be actively cooled.
Eckert and Livingood showed in the 1950s that transpiration-cooling is an effective method compared to conventional cooling (Eckert and Livingood, 1954).At that time, however, it was only possible to manufacture porous metal sponges for surface cooling that melt at a relatively low temperature and are, thus, not adequate for space applications.The studies focused mainly on aircraft propulsion, but were soon applied to blunt body cooling of re-entry vehicles as well.
As mentioned above, DLR has been investigating CMC rocket motors since the 1990s (Hald et al., 2005).Modern regeneratively cooled cryogenic combustion chambers rely on metallic designs and are highly optimized.Within a regeneratively cooled rocket engine, the fuel is usually routed around the outer surface of the inner liner and opposite the hot gas flow towards the injector faceplate before being injected.Thereby, sufficient heat can be exchanged to protect the predominant inner copper wall.However, in an environment of roughly 3500 K in the hot combustion zone and a few hundred Kelvin in the cooled exterior of the chamber wall, thermal fatigue problems occur due to the high thermal expansion of metals, thus, limiting the engine cycles.Regenerative cooling is also accompanied by a high pressure loss in the cooling channel structure, which is a substantial disadvantage, because the feed system must compensate for the pressure loss.
Figure 10.27 shows a radial cut view through a transpiration-cooled combustion chamber.The primary structure, a reinforced plastic, provides the structural integrity of the chamber by absorbing the mechanical loads induced by the chamber pressure.At the same time, it is a diffusion barrier for the hydrogen.The porous ceramic inner liner is cooled by a fraction of the cryogenic fuel passing through the material (Ortelt et al., 2013;Greuel, 2013).Cooling efficiency was proven by subscale testing.Both structural integrity as well as thermo-chemical resistance are given, showing operational safety.Beyond that, new technology approaches have been developed to introduce CMCs into regeneratively cooled as well as radiatively cooled systems.For example, a CMC injector could bring advantages related to high temperature injection (Ortelt et al., 2013).A similar concept was investigated at Ultramet (US), using three components: the outer jacket for structural support, an intermediate permeable core (here a SiC-foam) and a porous inner liner, a mixed molybdenum disilicide-silicon carbide (Sözen and Davis, 2008).This technology, in which a coolant is forced through a permeable wall component by a pressure gradient, has recently been used for re-entry load cases.Concepts based on porous transpiration or effusion-cooled CMC structures have been and are currently investigated within the programs IMENS (Kuhn and Hald, 2008), ATLLAS, and the DFG Research Center TRR40 (Langener et al., 2012).A variety of mass flow rates and porous materials are under investigation.Also, different coolant media, such as helium, nitrogen, air and argon, are tested.Wolf states that the amount of a gas-coolant required to provide a certain amount of transpiration-cooling of a body is approximately inversely proportional to the heat (or cooling) capacity of the gas according to (Wolf Jr. et al., 1966).In further investigations, liquids are used as a coolant (Wang et al., 2014).They have the advantage that the heat of vaporization can be used as an additional cooling mechanism.Water is a candidate frequently studied because it has a high heat of vaporization of 4168 J/kg K and is cheap and available.However, care must be taken so that the carbon fibers do not oxidize in the water vapour.
There is yet to be a fully integrated actively cooled TPS design, withstanding both thermal and mechanical loads.However, AKTiV, the first hypersonic flight experiment of a transpiration-cooled TPS element, was flown in the suborbital atmospheric re-entry body SHEFEX II in 2012 (Böhrk, 2014).
The AKTiV transpiration-cooled experiment was a key experiment of SHEFEX II, monitoring an actively transpiration-cooled thermal protection panel during a hypersonic return flight.In order to interpret the cooling effect of the transpiration, AKTiV was fitted with thermocouples, as well as pressure sensors and resistance thermometers.The experiment, using a porous and permeable C/C ceramic, was located onto a dedicated panel (C3, as shown in Figure 10.13).A non-pressurized reference set-up was mounted on the opposite panel (C7) where the same ambient flow conditions were expected.Although this location was not subject to extreme heat load for the SHEFEX II trajectory, the findings serve to support theories for future structural design with transpiration-cooling.
Figure 10.28 shows the AKTiV experiment set-up.A porous C/C 5 mm-thick sample of 61×61 mm 2 was inserted in the center of a 7 mm-thick thermal protection panel.The coolant, here nitrogen, was run through the porous sample, which was flanged into the surrounding TPS-panel by a pressure reservoir and riveted ceramic fasteners (Ortelt et al., 2001;Böhrk and Beyermann, 2010).The reservoir itself is made of stainless steel.The panel and sample are sealed against sneak flows by graphite Sigraflex felt.
The system consists of a pressurized tank, a pressure regulator, a valve on the vehicle side of the experiment and another pressure regulator, a mass flow controller (sonic nozzle), sensors and data acquisition hardware on the payload side of the experiment.A photograph of the entire set-up is shown in Fig. 10.29.The hot panel of AKTiV is instrumented with thermocouples as shown in Fig. 10.30.The thermocouples shall serve to monitor the effect of the film-cooling redundantly, specifically downstream of the porous sample, but also locally resolved, as designated by the white numbers in Fig. 10.31.The reference module on panel C7, monitoring the uncooled behaviour of the set-up, is indicated by black numbers in Fig. 10.31.A direct comparison will allow for assessment of the experiment's cooling efficiency, defined as (10.19) Cooling was switched on via telecommand at 431 s into the SHEFEX II flight.It can be seen in Fig. 10.32 that when the coolant began to flow, the temperatures of the panel decreased, showing the effect of the transpiration-cooling.This is demonstrated by measurement on the porous sample itself, but also in the film-cooling region downstream of the sample.The temperature difference with respect to the uncooled reference setup was largest at location K38 with 87 K, which, according to Eq. (10.19), corresponds to a cooling efficiency of 58%.Downstream of the sample, the temperature was effectively reduced by 74.5 K at K33, resulting in a cooling efficiency of 42%.The transverse temperature distribution in the downstream region of the cooled sample is shown in Fig. 10.33.The second excentric thermocouple K28 failed.However, the lowest temperature, i.e. the highest temperature reduction, was measured at the centermost thermocouple while the outermost thermocouples registered the highest temperatures.This corresponds well to the expected coolant film behaviour, where flow from the vicinity is entrained into the coolant film.Figures 10.34 and 10.35 show that the cooling is reproduced by the computer program HEATS, described in section 1.2, with only small deviations on the porous sample.Figures 10.32-10.35also show the interesting effect that the temperatures increase again after being cooled down from the time at which the coolant was turned on.This shows that the structure is not over-cooled and still reacts to increasing heat flux.This is supported by the simulations in HEATS.The measurements of AKTiV at t=485.12 s are plotted over distance x from the panel edge in Fig. 10.37.The starting temperature chosen for the simulation with HEATS was the average of the panel temperatures at 431 s.The result for the uncooled panel C7 shows good agreement with the measurement for the data from C/C-SiC, i.e. the panel into which the sample is embedded.Far higher deviations are observed at the C/Csample of AKTiV (0.7445 m -0.8055 m), which are attributed to the heat sink effect of the stainless reservoir.In the HEATS calculation, the temperatures of the C/C are higher than those of the C/C-SiC because of both lower thermal conductivity and a thinner sample thickness.However, in the flight measurement, the temperatures are reduced by the set-up with the reservoir.
Cooling efficiencies according to Eq. (10.19) are shown in Fig. 10.38 for a reference coolant temperature of T coolant =300 K. Figure 10.32 had shown that upon return from the apogee (see Fig. 10.12 for the trajectory), the sensor data of AKTiV deviated from that of the reference setup on the opposite vehicle side by an average of 8.5 K.This deviation is not yet explained, but it cannot be attributed to aerodynamic effects, because it originates at altitudes higher than those influenced by the atmosphere.This yielded an apparent cooling efficiency with an average of η ini =8.7%, which had to be subtracted from the efficiencies and finally resulted in the efficiencies displayed in Fig. 10.38.The corrected cooling efficiency η korr =η-η ini , thus, is 51% and 58%, up and downstream of the porous sample, respectively.Downstream the porous sample, it decreases with sensor distance from the sample with 42% and 30%.A cooling effect on the order of 11% can even be noticed upstream of the sample, and it is assumed that heat conduction caused this cooling effect.

Perspective and Concluding Remarks
The presented study shows the potential of advanced composite materials for the development of hypersonic flight through transpiration-cooled CMCs.However, transpiration-cooling applied to sharp leading edges promoting stagnation areas, could cause strong temperature gradients, and thus stress in the permeable material.Moreover, since the heat load is most severe in the tip of the leading edge, it will become necessary to direct more coolant there than anywhere downstream of the tip.This is also important because the stagnation area is not film-cooled by coolant from an upstream region.Because this is the case downstream, less transpiration-cooling is needed, there.In order to make the best use of the upstream injected gas and to protect the tip from the heat load, graded permeability can be used to counteract the pressure gradients.However, transpiration-cooling must be thoroughly investigated to see if it can serve efficiently in a TPS for orbital or higher atmospheric velocities.Liu et al. (2010) have investigated a transpiration-cooled wedge geometry made from stainless steel with a nose radius of 8.72 mm and a porosity of 42.8%.They used various gaseous media (air, N 2 , He, Ar, CO 2 ) for the cooling.They showed that the temperature gradients along the wall surface and cooling efficiency differ from those along a curved wedge geometry.The performance of water transpiration-cooling of a leading edge with phase change was investigated by Wang et al. (2014).The porous leading edge with 2 mm radius was made of sintered 316L alloy.The absolute porosity was 34%.They showed that a configuration with the thinnest wall thickness at the stagnation point provides enhancement of the cooling there.The setup of the experiment in the hot wind tunnel of the University of Science and Technology of China is shown in Fig. 1.38 Prediction tools for three-dimensional transpiration-cooling are needed in order to design the SHEFEX III leading edge.This includes the transient computational simulation of the cooled leading edge's thermal response.The simulation must be capable of considering strong gradients of all state variables.Moreover, it must compute the coolant distribution in the cooled wall at thermal non-equilibrium between coolant and structure and at wall pressure gradients.Also, special emphasis is to be placed on investigation and exact simulation of the film-cooling region.
The next step is the application of this technology to the SHEFEX III leading edge (Dittert et al., 2014).Figure 10.39 shows its current design, with porous material at the chime and film-cooled top and chin panels.Challenges include the prediction of the necessary coolant mass flow distribution over the profile, including cooling efficiency in turbulent flow, the realization of a graded permeability and directed mass flow rate, attachment of the structure to the vehicle, tightening of the reservoir towards the porous structure and many other aspects.

Figure 10
Figure 10.3: a) Mo-Re tube embedded in C/C illustrating the design of a NASP heat-pipe-cooled wing leading edge (Glass et al., 1999) and b) internal portion of an actively cooled leading edge as published by Glass.

Figure 10 . 6 :
Figure 10.6: Re-entry vehicles: a) Apollo, b) Stardust, which performed hyperbolic re-entry in 2006 after having collected dust from the coma of comet Wild 2, c) European re-entry vehicle ARD (Atmospheric Re-entry Demonstrator), which completed re-entry flight from a sub-orbital mission after launch with Ariane 5, d) Space Shuttle, which operated on 135 missions from 1981 to 2011.

Figure 10 . 10 :
Figure 10.10:Heat flux dependency on leading edge radius for orbital re-entry.

Figure 10 . 11 :
Figure 10.11: Payload of the SHEFEX II re-entry vehicle on the launcher on June 22nd 2012.

Figure 10 .
Figure10.12:DMARS data of the SHEFEX II trajectory.Black lines give in-flight measurement, gray lines give the assumed subsequent flight path(Turner et al., 2013).
the hot gas to the wall through convection with the heat transfer coefficient and the recovery temperature T r .These parameters are functions of the flow properties and can be expressed by means of the ambient velocity u ∞ , gas density ρ ∞ , specific heat c p,∞ , and the Stanton number St as α A = Stρ ∞ c p,∞ u ∞ (10.9) and (10.10) with the isentropic coefficient κ.The two commonly unknown variables of Eqs.(10.9) and (10.10) are the Stanton number St and the recovery factor r.They must be determined for each flow condition, be it laminar or turbulent.This can done based on the known models of Crocco and van Driest

Figure 10 .
Figure 10.22: SHEFEX II C/C-SiC sharp leading edge design and attachment strategy.
Thermocouple data were transmitted to a ground station during the SHEFEX II flight.They are shown in Figure 10.23 and the detailed view in Figure 10.24.Thin lines indicate the assumed flight path after the data transmission ended.The maximum recorded temperature was 1122 K at an altitude of 30 km.Applying the heat balance from the program HEATS to the SHEFEX II return flight trajectory (Figure

Figure 10 .
Figure 10.23: In-flight thermocouple data for the SHEFEX II trajectory (Fig. 1.12) compared to prediction by HEATS.Thin lines indicate assumed flight path after the data transmission ended.

Figure 10 .
Figure 10.24:A detailed view of the atmospheric return flight phase of the SHEFEX II trajectory compared to the prediction by HEATS.The thin lines indicate assumed flight path after the data transmission ended.

Figure 10 .
Figure 10.25:The RCC sharp leading edge on the X-43 [NASA].

Figure 10 . 33 :
Figure 10.33: Measured temperature transverse to the panel in first and second downstream location.(Böhrk,2014)

Figure 10 . 34 :
Figure 10.34:Comparison of HEATS with up-and downstream measurement.

Figure 10 . 35 :
Figure 10.35:Comparison of HEATS with data from the cooled sample.

Figure 10 .
Figure 10.36 shows the heat flux reduction derived from HEATS for the sensor locations K37 on the sample and K33 downstream of the sample with reference to the uncooled setup according to Eqs. (10.7).The figure shows that the heat flux to the cooled set-up is greatly reduced by the film.