1. INTRODUCTION

Ablative materials are used to protect vehicles from atmospheric reentry, to protect rocket nozzles and ship hulls from propellant gas erosion, as protection from laser beams, and to protect land-based structures from high heat environments. Ablative heatshield materials protect a vehicle from excessive heating, and can also act as a structural component. Thermophysically, ablation involves the elimination of thermal energy by sacrifice of surface material. Some principles of this heat and mass transfer process are melting, conduction and storage of heat in the material substrate, absorption of heat by gases, and exothermic and endothermic chemical reactions. The three groups of ablative materials are subliming or melting ablators, charring ablators, and intumescent ablators. Subliming ablators, eg, Teflon, carbon and composites, act as heat sinks until the surface reaches the sublimation or melting temperature, then removing heat from the insulation material. In the charring ablator, the surface material acts as a heat sink, up to the reaction temperature, where endothermic forms. Intumescent ablators form a foam like region on exposure to heat, resulting in improved insulation performance.

2. HISTORY OF ABLATOR

The first known ablative materials were meteorites. These thermally degraded bodies coming to us from space were indeed a subject of great curiosity, since they had a demonstrated in principle the utility of aerodynamic ablation for a thermal protection of atmospheric entry object. Perhaps in these bodies were hidden the secretes to successful re-entry of man-made vehicles. The search was thus initiated to discover their composition and construction. The information obtained was interesting, but it provided few usable clues. Furthermore, it was an apparent that the stony and iron meteoritic materials were not suitable for application to man-made thermal protection systems. Hence the search for workable ablative material was continued.Man-made ablative materials were discovered only a decade ago. Various technique were being explored to protect and insulate structural metals while exposed to a hot rocket exhaust. Certain reinforced plastics and ceramics were seen to exhibit remarkable durability during short time hyper thermal exposure. In addition, the high temperature of the environment was restricted to the surface region of the ablative material. These thermal barrier materials apparently and great potential for solving the high temperature problems associated with re-entry heating and rocket propulsion system. In the next several years, thousands of different material compositions and construction were characterised and evaluated on a trial and error basis. The high temperature facilities used were various combustion torches, small rocket motors, arc plasma jets, pebble bed heaters, arc imaging furnaces and others. Environmental simulation was often impossible to obtain these facilities, but there use permitted the generation of the test condition in which much useful materials information was obtained. Composite construction was most versatile, since unique properties of individual components were incorporated into a single material system.

3. ABLATIVE MATERIALS

AEROFAST is a typical research and development project funded by European Union, which aims at preparing a demonstration mission for aerocapture at Mars, as well as increasing the TRL of Aerocapture technology, in order to prepare for future missions such as Mars Sample Return (first potential mission to use the aerocapture technology).The aerocapture is a new technology for Solar System exploration that uses a single pass through a planetary atmosphere to decelerate the spacecraft and achieve a targeted orbit. Such manoeuvre saves a significant amount of mass with regard to a more conventional technique of insertion using propelled braking.Because of the high heat flux, aerocapture requires a thermal protection system to shield the spacecraft from aerodynamic heating, as well as the use of a guidance system to assure that the spacecraft leaves the planetary atmosphere on the correct trajectory. This has inferred the key objectives for the whole project, among which the development and modelling of an innovative TPS material.Relying on demonstrated performance of Astrium Norcoat-Liege for atmospheric entry probes, it has been chosen to develop advanced cork-based composites within this study. This is undertaken by Amorim, leader in the cork industry, and supplier among others of the P45 and P50 aerospace thermal protections.After reminding the main considered requirements, the general rationale and the selection process of the different formulations, the proposed paper will present the achievements of this development.The material must be able to withstand the severe front shield aerothermal environment. Numerous formulations have been investigated using a parametric combination of cork granule size, resin type/ratio, reinforcement fraction, fillers and the mixing and agglomeration processes. A basic (thermo-mechanical) characterization and qualitative analysis allowed for a first selection of the 4 most promising candidates. Thesecandidates have been tested in the inductive plasma facility (COMETE) of Astrium, for an aerothermal environment similar to the AEROFAST mission.In parallel, a 3D ablation and charring material model has been implemented in the finite element program SAMCEF, and successfully validated during the AEROFAST project. The numerical model consists of three sets of equations, namely the transient heat balance equation, the steady state mass balance equation and the charring equations. For the charring of the material we use a multi-species Arrhenius model with the species densities as degrees of freedom. The ablation is modelled by a surface imposed and temperature dependent ablation speed, followed by an in volume mesh deformation. On the basis of these preliminary experiments, additional efforts were also devoted to the modelling of the thermal, swelling and ablative behaviour of the selected cork based material (developed within this project).It was stated a few years ago that no existing European ablator was suitable for sustaining the very high heat fluxes while coping with the stringent mass requirements. As this component is clearly an enabling technology for the considered mission, Astrium then initiated the development of the low density ablator named ASTERM.ASTERM concept is based on a low density carbon phenolic composite, which was shown to be a quite effective solution by the US PICA. Beyond its primary use for Sample Return missions, such a solution can indeed be used for various applications: ISS return (ARV), Mars (cf MSL) and eventually Venus exploration.The Astrium ST R&T effort enabled elaborating and validating the development process, as well as anticipating the exploration of various versions. Quite mature solutions were thereforeproposed for evaluation within the ESA TRP DEAM study (Development of European Ablative Materials) conducted in parallel.Relying on the promising test results obtained from the above-mentioned activities, ASTERM was then selected by ESA for the delta development starting within the DEAM2 activity.In order to be ready for subsequent mission development phases, a TRL (Technology Readiness Level) of 6 is the current target for 2014. The associated tasks are obviously not limited to material elaboration, and they include the following aspects:Manufacturing trials to consolidate fabrication approach and process Elementary characterisation Validation of the behaviour under aerothermodynamics conditions representative of the atmospheric entry of the envisioned future vehicles. Elaboration of a thermal and ablation modelThe success of an asteroid sample return mission such as MarcoPolo-R or Mars Sample Return (MSR) mission heavily relies on availability of a light European heat shield technology able to withstand the very high energy Earth re-entry conditions of the Earth Return Capsule (ERC), resulting typically in peak heat fluxes in the order of 8-20 MW/m2.World wide experience for space exploration and high heat flux entry is very limited, and all missions to date have selected tape-wrapped carbon-phenolic composite Thermal Protection System (TPS) with density in the range of (1.3 - 1.4) supported by a cold structure. These missions are Pioneer-Venus (PV) and Galileo entry probes from NASA, and Hayabusa entry probe from JAXA.

Its selection is due in particular to the fact that:

All of sample return mission with high energy ERC entry vehicles require a robust heat shield TPS, in particular for "category 5 restricted return" such as for MSR, with very challenging reliability requirement for the re-entry capsule. Carbon phenolic composites of density in the 1.3-1.4 exhibit highest heat resistance and lowest conductivity, guarantee of the lowest ablation (authorizing less TPS thickness - i.e. minimum weight - and minimum aero-shape changing - i.e. better aerodynamic control). Denser material are furthermore more robust with regard to the MMOD impact.

Although carbon-phenolic is definitely the best robust TPS candidate, the high density of the material could end up to a weight penalty in the overall mass budget of the probe. This has been highlighted in both MarcoPolo and MSR Earth Return design studies. Nevertheless the concept Sepcore proposed for many years by Herakles (HKS), and based on carbon phenolic joined to a CMC hot structure, allows to manufacture a light heat shield, but at the same time robust and reliable.

3.1 ABLATIVE MATERIALS THERMAL PROPERTIES ESIMATING

In many practical situations it is impossible to measure directly thermal and thermokinetic properties of analyzed composite materials. The only way that can often be used to overcome these difficulties is indirect measurements. This type of measurements is usually formulated as the solution of inverse heat transfer problems. Such problems are ill-posed in mathematical sense and their main feature shows itself in the solution instabilities. That is why special regularizing methods are needed to solve them. The general method of iterative regularization is concerned with application to the estimation of materials properties. The objective of this paper is to estimate thermal and thermokinetic properties of advanced materials using the approach based on inverse methods. An experimental-computational system is presented for investigating the thermal and kinetics properties of composite materials by methods of inverse heat transfer problems and which is developed at the Thermal Laboratory of Department Space Systems Engineering, of Moscow Aviation Institute (MAI). The system is aimed at investigating the materials in conditions of unsteady contact heating over a wide range of temperature changes and heating rates in a vacuum, air and inert gas medium. In estimating temperature-dependent properties of modern composite destructive materials the most effective are methods based on solution of the coefficient inverse heat transfer problems. The most promising direction in further development of research methods for destructive composite materials using the procedure of inverse problems is the simultaneous determination of a combination of material's thermal and kinetic properties (thermal conductivity, heat capacity, heat capacity of charring gas, thermokinetics and some other parameters). Such problems are of great practical importance in the study of properties of composite materials used as destructive surface coating of spacecrafts.The experimental equipment and the presented method could be applied for estimating of material's seven characteristics; the availability of a few specimens of the material allows us toprovide uniqueness of the solution. The application of the considered technique for real thermoprotective materials is presented. The results of temperature measurements inside the specimen are assigned as necessary additional information to solve an inverse problem. To construct an iterative algorithm of the inverse problem solving a conjugate gradient method was used. In the approach being developed the methods of calculus of variations are used for calculation of the minimized functional gradient.For partially decomposed materials the model of heat conduction with temperature-dependent thermal characteristics is approximate, and characteristics are effective, since the heat transfer in such material is provided not only by heat conduction but also by different transformation processes depended on conditions of heating. A deviation of calculated and experimental temperature values in the experiments did not exceed 8 K, that confirms the possibility of using, for the given material, a model of heat conduction with the effective thermal characteristics. But the presented method can be used only for determining the effective thermal characteristics of composite materials for particular heating conditions.

3.2 WORKING OF ABLATIVE MATERIALS

The most commonly used ablative materials are the composites, i.e., materials consisting of a high-melting point matrix and an organic binder. The matrix can be glass, asbestos, carbon or polymer fibers braided in different ways. In some cases, a honeycomb construction can be used, filled with a mixture of organic and nonorganic substances and possessing high heat-insulating characteristics (as used, for instance, on the space vehicle Apollo).

Figure1.Schematic model for the destruction of an ablating composite material.

Shown in Figure1 is a schematic model of the destruction of a composite material from a high-melting point matrix and an organic binder. The characteristic property of such heat-shielding coverings is the presence of two fronts or zones, to be more exact, in which physicochemical transformations take place. In convective heating, a viscous melt film can be formed on the surface of such composite materials. Despite its thinness, the film strongly affects the destruction process. In particular, the coalescence of particles of the surface layer prevents their erosion blow-off by the flow. The melt film also reduces the rate of oxidation of chemically-active components of the material by the incoming flow of gas.Further into the surface lies a comparatively thick layer of charred organic binder reinforced by high-melting fibers. Still deeper is the thermal decomposition zone, where a mixture of volatile and solid (coke) components is formed. The volatile components filtered through the porous matrix are injected into the boundary layer of the incoming gas flow. An intensive sublimation of glass or other oxides which form high-melting fibers occurs on the surface of the melt film. The fraction of gaseous ablation products in the total ablation mass can, therefore, be high. The particles of coke are practically pure carbon; thus, at the melting temperature of glass they remain solid. The spreading film of glass breaks out the porous structure of the charred layer and carries away the particles of coke. The later, in turn, affects the flow of the melt, increasing its effective viscosity.At high temperatures, the coke particles in the melt film are not inert components they interact actively both with glass and with any oxidant present in the gas flow. Tens of various strongly interacting components can exist in the boundary layer over the surface of the composite heat-shielding covering. The choice of a theoretical model for the destruction process of such materials, presents considerable difficulties. However, on the basis of extensive experimental and theoretical studies of thermophysical, thermodynamic and strength phenomena which attend the process of the incident flow effect, we have succeeded in creating a schematic model or a mechanism for the destruction of a heat-facing layer. Such a mechanism has been designed only for some classical representatives of the range of composites (see Sublimation, Melting). At the same time, advances in chemistry and materials technology extend the possibilities of selecting improved ablation materials. In this context, a demand arose for some unique parameter to compare various types of ablative materials convenient for both theoretical and experimental studies. One such parameter is the effective enthalpy of destruction, symbolized as heff.The effective enthalpy defines the total thermal energy expenditure necessary to break down a unit mass of ablative material. The problem of comparing numerous ablative materials is most easily demonstrated for a quasi-stationary destruction (see Heat Conduction) when the velocity of all isotherms or destruction fronts inside the material coincides with the velocity of the outer surface displacement. In this case, the temperature profile inside the heat-shielding covering is described by a set of exponents, and the heat flux spent on heating inner layers does not depend on the material thermal conductivity .Let us first consider a destruction process under conditions of exposure to convective heating. The thermal balance on a destructing surface (Figure2) can be written as follows:

Figure2.Destruction process with convective heating.Here, (/cp)0 is the heat transfer coefficient, and he and hw are the enthalpies of the gases in the incoming flow and the wall, respectively. In contrast to a nondestructing ablative facing, the convective heat flux supplied from without is expended not only for heating the material ( ) and by radiant re-emission of the four heated surfaces ( T4W ) but also for the surface (with mass loss rate and bulk (with mass loss rate physicochemical transformations, whose thermal effects are evaluated as Qw and Q. If a melt film is formed on the surface of a heat-shielding covering, then , where is the mass loss rate of a substance in a molten form. The total thermal effect of the bulk failure Q contains not only the heat of matrix melting, but also the thermal effect of the thermal decomposition of an organic binder, the heat of heterogeneous interaction between the glass and coke inside the charred layer, etc. In a similar manner, the thermal effect of surface destruction Qw must account for the thermal effect of evaporation of a melted film and the burning of the coke particles in the incoming flow of gas.Gaseous ablation products which penetrate into the boundary layer cause a reduction of a convective heat flow due to the so-called injection effect. We can evaluate the blocking action of the injection effect by a linear approximation.Here, is the dimensionless coefficient of injection ( < 1), which in the general case depends on flow conditions in the boundary layer (laminar or turbulent) and the ratio of molecular masses of the gas injected and the incoming flow. Unlike other effects influencing the absorption of the heat energy supplied, the injection effect rises steeply with the increasing velocity or temperature of the incoming flow and finally becomes predominant.If we denote the share of gaseous ablation products in the total mass loss of the substance by ( = / ), then we can obtain a generalized characteristic of destruction power, namely, the effective enthalpy of destruction, heff:The effective enthalpy determines the amount of heat which can be blocked when breaking down a unit mass of covering material (whose surface temperature is Tw) through physicochemical processes. The higher the effective enthalpy, the better the heat-shielding material. We place emphasis on the independence of the effective enthalpy from the geometrical dimensions or the shape of the body. Actually, as distinct from a heat flux whose value, with the given parameters of the incoming flow (pe, he), is inversely proportional to (where RN is the typical dimension of the body; for instance, the radius of curvature in the vicinity of the critical point), the effective enthalpy is unaffected either by the shape or the dimension of the body. This qualifies it as a parameter for relating laboratory and real heat-loading situations.We can see from the definition of effective enthalpy that in all cases when 0, it must increase substantially with the rise in the enthalpy of the stagnated flow he. The parameters of the incoming gas flow (pressure Pe and enthalpy he) can effect heff through changes in the temperature of the destructing surface Tw, the fraction of the ablation which is in gaseous form and the thermal effect of surface processes Qw. The effect of surface temperature Tw on heff can be considered to be rather limited. A typical dependence of Tw, and heff on enthalpy he and pressure Pe in breaking down glass reinforced plastics in an air flow is shown in Figures 3, 4 and 5. The flow condition (laminar or turbulent) in the boundary layer determines the injection coefficient (see Heat Protection), which affects radically the dependence of heff on he (Figure6 ). If the ablative material does not contain oxides, then, as a rule, the share of gasification is close to unity. For graphite-like heat-shield covering, in particular, = 1. In this case, however, the thermal effect of surface processes Qw varies from a negative value on carbon burning C + O2 = CO2 to a positive value upon its sublimation. An extra liberation of heat upon burning brings about surface overheating relative to the equilibrium value of the temperature for a heat-insulated wall. In this case, the effective enthalpy becomes negative and the notion of heff loses practical sense. The dimensional rate of destruction is often used as an alternative parameter for generalizing the experimental and the design data.Its advantage is that the function (he ) is always positive and besides, the temperature of the destructing surface Tw and the emissivity are not warranted. Typical dependences of on the stagnation enthalpy he for Teflon, glass-reinforced plastic and graphite breaking down in air flow are shown in Figure7.

Figure3.The share of gasification as a function of stagnation enthalpy of incoming gas he.

Figure4.

Figure5.

Figure6.

Figure7.Dimensionless destruction rate ( ) as a function of stagnation enthalpy (he) for various materials breaking down in an air flow.Combined radiation-convection heating of the surface of an ablative material can considerably change the mechanism of its destruction. The injection of gaseous disintegration products in cases where they do not possess high absorption coefficients, slightly reduces the intensity of the radiation component of the heat flow. As the ratio grows, the mechanism of destruction of the majority of ablative materials more closely resembles sublimation and thermal decomposition. This is due to a rapid decrease in the contribution of convective and diffusion transfer in the boundary layer while injecting gaseous products, to the ceasing of melt film flow and to the absence of burning on the destructing surface.The heat balance on the surface of an ablative material in case of high levels of radiation of heat flows is simplified as follows.Here, K, w is the absorption coefficient, which depends on the spectrum of incident radiation heat flow () and on the spectral distribution of the destructing surface emissivity ():When no mechanical cracking or melting of a heat-shielding material occurs, the total rate of ablation coincides with and the notion of effective enthalpy of the material under intensive radiation heat influence can be introduced as:Table1 shows the results of the evaluation of parameters h, K, w (in the 0.2 < < 1 m spectral range) and hR for various substances.Table1.Materialh, kJ/kgK, whR, kJ/kg

Graphite30.0000.8535.000

Quartz15.0000.275.000

Magnesium oxide15.0000.13115.000

Teflon3.0000.130.000

An analysis of the data presented in Table 1 allows us to reach a paradoxical conclusion: under the influence of intensive radiation, the effective enthalpies of destruction of graphite and Teflon become about equal. We should note that the ablation rate of graphite, as compared to magnesium oxide, does not differ so strongly as the other values of the effective enthalpies given in the table. This is associated with the fact that the temperature of graphite destruction is almost half as great, and, therefore, the levels of the reemitted energy differ by an order of magnitude. Nonetheless, the main conclusion that can be drawn in analyzing Table1 is that by decreasing the absorption coefficient of the destructing surface (K, w), we can obtain a greater efficiency of ablation than by increasing the heat of sublimation.

4. ADVANTAGES OF ABLATIVE MATERIALS High Heat Absorption and Dissipation Exceptional Thermal insulation No maximum Service Temperature Weight saving Resistance to thermal and mechanical Shock availability Design Simplicity and Flexibility Low cost Passive in Operation5. LIMITATIONS OF ABLATIVE MATERIALS Susceptible To High Mechanical Forces Service Life is Time Dependent

6. APPLICATION OF ABLATIVE MATERIALSEffective and efficient design with ablative materials has seldom been achieved, due to the newness of the materials, lack of preceding similar designs, complexity of the design factors involved, and our incomplete knowledge concerning realistic design criteria. In designing with ablative materials, consideration must be given to the environmental variables and their time dependency, availability and uniformity of candidate materials compositions and constructions, material properties and characteristics, materials formulation and fabrication, design requirements for thermal, mechanical and chemical properties, safety factors and other aspects peculiar to the design. Some degree of uncertainty exists for each of these factors, and designers have had a tendency to use an overall safety factor rather than one based on the uncertainty of each design criterion. Further research on optimum design techniques for ablative materials is required. HYPERSONIC ATMOSPHERIC RE-ENTRYOne of the most difficult and challenging problems of aerospace flight is the thermal protection of a vehicle as it enters hypersonically planetary atmosphere ROCKET PROPULSION EXHAUSTSThe containment and control of hot combustive gases in rocket propulsion systems is necessary for thrust purposes. These propellant gases constitute a severe engineering environment, since they are generally characterized by high temperature, high mechanical forces, chemical corrosion, and occasionally particle erosion.

7. CONCLUSIONThe significance of ablative materials to aerospace technology os now apparent. Our successes in solving the Re-entry heating problem and in providing light weight, high performance propulsion materials are history. The current stateof the art representsonly the generatin developments,and only asmall portion of the hyper-environmental spectrum has been investigated. Current materials deficiencies mustbe overcome, and new ablaive materials with unique properties and characteristics are necessary. Each new success in this work will permit a wider range of aerospace systems and new capabilities in aerospace technology.

8. REFERENCE

1. Duerig, T., Melton, K., Stockel, D., and Wayman, C. (Eds) Engineering aspects of shape memory alloys, 1990 (Butterworth-Heinemann, London). 2. Lagoudas, D. C., Entchev, P B., Popov, P Patoor, E.Brinson, L. C., and Gao, X.composite materials and alloys , part II: modeling of polycrystals. Mech. Mater. 2006, 38(56), 430462. 3. Otsuka, K. and Wayman, C. M. (Eds) Shape memory alloy and ablative materials, 1999 (Cambridge University Press, Cambridge).4. Patoor, E., Lagoudas, D. C., Entchev, P B., Brinson, L. C and Gao, X. Composite materials and alloys, part I: general properties and modeling of single crystals. Mech. Mater., 2006, 38(56), 391429.5. Srinivasan, A. V. and McFarland, M. D. Smart structures: analysis and design of ablative materials 2000 (Cambridge University Press, Cambridge).

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