Advanced Ceramic Materials for Sharp Hot Structures: Material Development and on-Ground Arc-Jet Qualification Testing on Scaled Demonstrators L. Scatteia, G. Tomassetti, G. Rufolo, F. De Filippis, G. Marino CIRA-Italian Aerospace Research Centre- Via Maiorise 81043 CAPUA(CE)-ITALY Abstract. This paper describes the work performed by the Italian Aerospace Research Centre(CIRA. S.c. P.A. )in a technology project focused on the applicability of modified diboride compounds structures to the manufacturing of high performance and slender shaped hot structures for reusable launch vehicles. A prototypal multi-material structure, which couple reinforced diborides to a C/SiC frame, has been built with the aim to demonstrate the applicability of an innovative concept of nose cap to the fabrication of real parts to be installed ant subsequently tested on the flying test bed currently under development at CIRA. Particular relevance is given to the on-ground qualification test of the nose. ap scaled demonstrator which is underway at CIRA Arc-Jet facility SCIROCCO. Considering the specific typology of materials investigated, up to date, a consistent tests campaign at laboratory level has been order to create a complete materials data base. The measured materials properties have been then used as input for the design phase that also used as inputs the aero-thermal loads associated with a reference re-entry mission. Our major preliminary findings indicate that the structure is thermally fully compliant with the environment requirements and shows local mechanical criticalities in specific areas such as the materials interfaces and hot/cold joining parts. INTRODUCTION Thermal protection systems represent the key issue for the successful re-entry of a space vehicle (Behrens Future concepts for space launchers foresee sharp aerodynamic profiles as conventional aircrafts(McKenzie This kind of architecture offers several advantages with respect to current blunt shapes: maneuve improvement, decrease of electromagnetic interferences and communication black-out and drag reduction di ascent phase. As a drawback, aerodynamic heat flux increases dramatically(reaching 650-800 KW/m2). State of hot structures materials cannot withstand the thermal requirements of future slender-shaped RlVs The Sharp Hot Structures Project (SHS) is focused on the applicability of modified diboride compounds to the manufacturing of high performance and slender shaped hot structures. Zirconium and Hafnium diborides/silicon carbide composites are under investigation: those compounds are actually addressed as the sole materials that can b niently employed at temperatures above 2200K(Bellosi, 2000) SHS project activities are performed within a research network, managed by CIRA, involving Centro Sviluppo Materiali S.P.A(CSM), University of Rome"La Sapienza, the Institute of Science and Technology for Ceramic of the Italian National Research Council(CNR-ISTEC), Fabbricazioni Nucleari(FN) and University of T The main objective of this project is to provic ground qualified advanced technology products identified in critical parts of re-entry vehicles such se cap and wing leading edges to be then tested and validated in flight The final products of the research are, on the short term, an innovative nose cap, and on the long term, a wing leading edge. These components will be respectively tested on the sub-orbital re-entry mission and in the hypersonic flying test of the USV-Flying test-bed n 2 CP746, Space Technology and Applications International Forum--STAIF 2005, edited by M. S. El-Genk C2005 American Institute of Physics 0-7354-0230-2/05/S22.50
Advanced Ceramic Materials for Sharp Hot Structures: Material Development and On-Ground Arc-Jet Qualification Testing on Scaled Demonstrators L. Scatteia, G. Tomassetti, G. Rufolo, F. De Filippis, G. Marino CIRA – Italian Aerospace Research Centre – Via Maiorise 81043 CAPUA (CE) – ITALY Abstract. This paper describes the work performed by the Italian Aerospace Research Centre (C.I.R.A. S.c.P.A.) in a technology project focused on the applicability of modified diboride compounds structures to the manufacturing of high performance and slender shaped hot structures for reusable launch vehicles. A prototypal multi-material structure, which couple reinforced diborides to a C/SiC frame, has been built with the aim to demonstrate the applicability of an innovative concept of nose cap to the fabrication of real parts to be installed ant subsequently tested on the flying test bed currently under development at CIRA. Particular relevance is given to the on-ground qualification test of the nosecap scaled demonstrator which is underway at CIRA Arc-Jet facility SCIROCCO. Considering the specific typology of materials investigated, up to date, a consistent tests campaign at laboratory level has been performed and concluded in order to create a complete materials data base. The measured materials properties have been then used as input for the design phase that also used as inputs the aero-thermal loads associated with a reference re-entry mission. Our major preliminary findings indicate that the structure is thermally fully compliant with the environment requirements and shows local mechanical criticalities in specific areas such as the materials interfaces and hot/cold joining parts. INTRODUCTION Thermal protection systems represent the key issue for the successful re-entry of a space vehicle (Behrens, 2003). Future concepts for space launchers foresee sharp aerodynamic profiles as conventional aircrafts (McKenzie, 2003). This kind of architecture offers several advantages with respect to current blunt shapes: maneuverability improvement, decrease of electromagnetic interferences and communication black-out and drag reduction during the ascent phase. As a drawback, aerodynamic heat flux increases dramatically (reaching 650-800 KW/m2). State of art hot structures materials cannot withstand the thermal requirements of future slender-shaped RLVs. The Sharp Hot Structures Project (SHS) is focused on the applicability of modified diboride compounds to the manufacturing of high performance and slender shaped hot structures. Zirconium and Hafnium diborides/Silicon carbide composites are under investigation: those compounds are actually addressed as the sole materials that can be conveniently employed at temperatures above 2200K (Bellosi, 2000) . SHS project activities are performed within a research network, managed by CIRA, involving Centro Sviluppo Materiali S.p.A. (CSM), University of Rome “La Sapienza”, the Institute of Science and Technology for Ceramics of the Italian National Research Council (CNR-ISTEC), Fabbricazioni Nucleari (FN) and University of Turin. The main objective of this project is to provide on-ground qualified advanced technology products identified in critical parts of re-entry vehicles such as nose cap and wing leading edges to be then tested and validated in flight conditions. The final products of the research are, on the short term, an innovative nose cap, and on the long term, a wing leading edge. These components will be respectively tested on the sub-orbital re-entry mission and in the hypersonic flying test of the USV-Flying test-bed n.2 129
The PRORA USV Space Program In the framework of the Italian National Space Research Program(PRO. RA), supported by the Italian Ministry of Education and Research (M.I. U R), the Italian Aerospace Research Centre(C I.R. A is conducting the aerospace USV (Unmanned Space Vehicle) Research program(Russo, 2002; Marino, 2002). The USV program is aimed at developing and validating, up to flight tests, key technologies for the next generation of reusable space transportation vehicles. The program embraces the following main area of interest: aerothermodynamics, structures and Materials, propulsion, guidance, Navigation, and Control. Technological project are currently ongoing at Cira in each of the aforementioned research branches. Together with R&D activities, the development of a family of experimental vehicles(FTB: Flight Test Bed) is underway. These vehicles will be employed to perform fo missions at increasing complexity: Dropped Transonic Flight Test (DTFT), Sub-orbital Re-ent Hypersonic Flight Test (HFT), Orbital Re-entry Test(ORT). Each mission is conceived to test the de technologies in One of the main projects of the structures and materials research area is the Sharp Hot Structures project, which is the subject of this paper and will be described in details in the following paragraphs SHS PROJECT STRUCTURE AND LOGIC The Sharp Hot Structure Project is articulated into the following phases: a)a basic research on selected UHTC materials conducted in parallel with the related manufacturing processes assessment; b) hot structures thermo- mechanical design, which drive the materials and process assessment; c) Hot structure scaled demonstrator manufacturing; d)On-ground qualification test at CIRA Scirocco Plasma Wind Tunnel; e) In-flight validation test of the component, in the Sub-Orbital re-entry mission of the USV Flying Test-Bed. MATERIAL CHOICE RATIONALE The Sub-Orbital Re-entry Test of PRO. R. A-USV FTB-2 vehicle will be characterized by very high thermal loads that conventional CMCs such as C/C and c/sic, although reliable and well tested are not able to sustain. Aerospace research is moving towards ceramic systems based on hafnium, zirconium and titanium borides, in account of their high configuration stability(ablation resistance) in the presence of high velocity dissociated air, high thermal shock and thermal fatigue resistance (Levine, 2002) The fundamental concept at the basis of a successful use of these systems in oxidizing atmosphere lays in the growth of a protective oxide layer on the surface of the component. Between the Diboride family of ceramic, Zirconium diboride is particularly promising with respect to Hafnium and other compounds thanks to its lower densi Nevertheless, ZrB2 alone cannot be conveniently employed. Actually, in an high oxid temperatures boron oxide is formed: its high volatility for temperature higher than totally hinders all the advantages of ZrB2 melting temperature(>3000 C). The formation of a stable and pro re oxide layer is doesn't take place and potential high temperature use of this material is impaired (Marino, 2003) For this in the material selection phase of the Shs project a ZrB2-Sic compound was chosen. ZrB2-SiC compound forms, in high temperature oxidizing environments, a boro-silicate glass based surface layer, which protects the bulk from further oxidation. High aerodynamic performance sharp leading edge components, obtained by sintered ZrB2-SiC composites, were successfully tested by NAsa in the frame of the sharP project and other researches, thus evidencing the technical feasibility of this solution and the effectiveness of the protective action of the silica based layer. Nevertheless the production processes experimented up to date are limited to sintering techniques showing dimensional limitations and unsuitable for the realization of protective coatings. During the SHs project, the non-conventional Plasma Spray Deposition technique (Valente, 2000) was selected in order to obtain thin protective ZrB2-Sic coating on a structural C/Sic long fiber composite frame, as described in the following paragrap 130
The PRORA USV Space Program In the framework of the Italian National Space Research Program (PRO.RA), supported by the Italian Ministry of Education and Research (M.I.U.R), the Italian Aerospace Research Centre (C.I.R.A.) is conducting the aerospace USV (Unmanned Space Vehicle) Research program (Russo, 2002; Marino, 2002). The USV program is aimed at developing and validating, up to flight tests, key technologies for the next generation of reusable space transportation vehicles. The program embraces the following main area of interest: aerothermodynamics, structures and Materials, propulsion, guidance, Navigation, and Control. Technological project are currently ongoing at CIRA in each of the aforementioned research branches. Together with R&D activities, the development of a family of experimental vehicles (FTB: Flight Test Bed) is underway. These vehicles will be employed to perform four flight missions at increasing complexity: Dropped Transonic Flight Test (DTFT), Sub-orbital Re-entry Test (SRT), Hypersonic Flight Test (HFT), Orbital Re-entry Test (ORT). Each mission is conceived to test the developed technologies in actual flight re-entry conditions. One of the main projects of the structures and materials research area is the Sharp Hot Structures project, which is the subject of this paper and will be described in details in the following paragraphs. SHS PROJECT STRUCTURE AND LOGIC The Sharp Hot Structure Project is articulated into the following phases: a) a basic research on selected UHTC materials conducted in parallel with the related manufacturing processes assessment; b) hot structures thermomechanical design, which drive the materials and process assessment; c) Hot structure scaled demonstrator manufacturing; d) On-ground qualification test at CIRA Scirocco Plasma Wind Tunnel; e) In-flight validation test of the component, in the Sub-Orbital re-entry mission of the USV Flying Test-Bed. MATERIAL CHOICE RATIONALE The Sub-Orbital Re-entry Test of PRO.R.A-USV FTB-2 vehicle will be characterized by very high thermal loads that conventional CMCs such as C/C and C/SiC, although reliable and well tested, are not able to sustain. Aerospace research is moving towards ceramic systems based on hafnium, zirconium and titanium borides, in account of their high configuration stability (ablation resistance) in the presence of high velocity dissociated air, high thermal shock and thermal fatigue resistance (Levine, 2002). The fundamental concept at the basis of a successful use of these systems in oxidizing atmosphere lays in the growth of a protective oxide layer on the surface of the component. Between the Diboride family of ceramic, Zirconium diboride is particularly promising with respect to Hafnium and other compounds thanks to its lower density (Monteverde, 2002). Nevertheless, ZrB2 alone cannot be conveniently employed. Actually, in an high oxidizing environment and at high temperatures boron oxide is formed: its high volatility for temperature higher than 1000°C totally hinders all the advantages of ZrB2 melting temperature (>3000°C). The formation of a stable and protective oxide layer is doesn’t take place and potential high temperature use of this material is impaired (Marino, 2003). For this reason, in the material selection phase of the SHS project a ZrB2-SiC compound was chosen. ZrB2-SiC compound forms, in high temperature oxidizing environments, a boro-silicate glass based surface layer, which protects the bulk from further oxidation. High aerodynamic performance sharp leading edge components, obtained by sintered ZrB2-SiC composites, were successfully tested by NASA in the frame of the SHARP project and other researches, thus evidencing the technical feasibility of this solution and the effectiveness of the protective action of the silica based layer. Nevertheless the production processes experimented up to date are limited to sintering techniques showing dimensional limitations and unsuitable for the realization of protective coatings. During the SHS project, the non-conventional Plasma Spray Deposition technique (Valente, 2000) was selected in order to obtain thin protective ZrB2-SiC coating on a structural C/SiC long fiber composite frame, as described in the following paragraph. 130
Nose Cap Structural Concept and Selected Material/Process Systems Figure 1 depicts a schematic of the nose cap under development. The nose is composed by: a)a bulk grap b)a truncated conical C/Sic frame manufactured by polymer infiltration and Pyrolisis process c)a ZrB2-Sic coating applied on the C/Sic frame by plasma spray deposition technique; d)a ZrB2-Sic massive conical tip produced by sintering technique. Each of the identified(material)/(manufacturing process) systems was subjected to a complete characterization test campaign, in order to provide the thermo-mechanical design with the required database of properties. Different ZrB2-X(where X is the compounding additive) compositions were considered, and depending on the composition, dense samples were obtained using hot pressing In vacuum, gas pressure sintering These samples were subjected to a full characterisation test campaign including: microstructure evaluations(XRD, SEM coupled with a eDX microanalyzer); flexural tests (4-pt fixture from R.T. up to 1500%C); Fracture toughnes (chevron notch" method on (250x25x2.0)mm3 bar); thermal expansion(up to 1350.C in flowing Argon) resistance to oxidation: (short/long term tests up to 1600.C in laboratory air). In accordance to the obtained experimental results, a diboride matrix composite including only Sic as a second phase behaves as the most promising composition from the mechanical and oxidation resistance point of view, and was therefore selected for he massive conical ti TrB2 sintered C/sic Plasma Sprain 1=200 mm FIGURE l Nose Cap Structural Concept and Constituent Materials COMPUTATIONAL FLUID-DYNAMIC ANALYSIS conditions that characterize above all the low-earth orbit part of a typical space vehicle Be a correct characte of the aero-thermal environment surrounding a structure re-entering into the atmosphere is a ke for the success of a new TPS concept design. As er of fact the existing on-ground facilities do not allow the simultaneous experimental reproduction of all nermo-fluid-dyna stagnation point heat flux and pressure. Moreover, the environment reproduced within the arter Even in the case of ve ery large facilities, as CIRA SCIROCCO, it may be difficu ce both different from the real one from a chemical point of view. Namely, the airflow strongly dissocial heater and remain frozen up to the test section. This phenomenon, commonly addressed as air vitiation, may strongly influence the material behaviour, i. e. even if the energetic level of the flow within the arc jet is the same of
Nose Cap Structural Concept and Selected Material/Process Systems Figure 1 depicts a schematic of the nose cap under development. The nose is composed by: a) a bulk graphite core; b) a truncated conical C/SiC frame manufactured by polymer infiltration and Pyrolisis process c) a ZrB2-SiC coating applied on the C/SiC frame by plasma spray deposition technique; d) a ZrB2-SiC massive conical tip produced by sintering technique. Each of the identified (material)/(manufacturing process) systems was subjected to a complete characterization test campaign, in order to provide the thermo-mechanical design with the required database of properties. Different ZrB2-X (where X is the compounding additive) compositions were considered, and depending on the composition, dense samples were obtained using hot pressing in vacuum, gas pressure sintering and pressureless sintering. These samples were subjected to a full characterisation test campaign including: microstructure evaluations (XRD, SEM coupled with a EDX microanalyzer); flexural tests (4-pt fixture from R.T. up to 1500°C); Fracture toughness (“chevron notch” method on (25.0x2.5x2.0) mm3 bar); thermal expansion (up to 1350°C in flowing Argon); resistance to oxidation: (short/long term tests up to 1600°C in laboratory air). In accordance to the obtained experimental results, a diboride matrix composite including only SiC as a second phase behaves as the most promising composition from the mechanical and oxidation resistance point of view, and was therefore selected for the massive conical tip manufacturing. FIGURE 1. Nose Cap Structural Concept and Constituent Materials. COMPUTATIONAL FLUID-DYNAMIC ANALYSIS A correct characterization of the aero-thermal environment surrounding a space structure re-entering into the atmosphere is a key factor for the success of a new TPS concept design. As a matter of fact the existing on-ground facilities do not always allow the simultaneous experimental reproduction of all the thermo-fluid-dynamics conditions that characterize above all the low-earth orbit part of a typical space vehicle re-entry path. Even in the case of very large facilities, as CIRA SCIROCCO, it may be difficult to contemporary reproduce both stagnation point heat flux and pressure. Moreover, the environment reproduced within the arc-jet facilities is quite different from the real one from a chemical point of view. Namely, the airflow strongly dissociate trough the archeater and remain frozen up to the test section. This phenomenon, commonly addressed as air vitiation, may strongly influence the material behaviour, i.e. even if the energetic level of the flow within the arc jet is the same of 131
the flight one, in the former case a large amount of energy is frozen within the fluid as formation enthalpy of dissociated atomic species. If, for instance, the material has a partially catalytic behaviour it is essential to be able to properly characterize the difference between the flight and ground environment in order to better understand which mechanism of heat release to the wall surface prevails: conductive or chemical. Moreover, when the article to be tested is scaled and/or of slightly different shape with respect to the real one, the correct reproduction of heat flux and pressure at the stagnation point do not assure in general that we are simulating the same environmen downstream of the stagnation point. In the same way, the small radius of curvature of sharp structures together with the low Reynolds flow obtainable with an arc-jet facility may give rise to undesired rarefaction effects. All of this issues strongly claims for an extensive use of Cfd both for the extrapolation from simulated flight condition to suitable operating condition of the plasma wind tunnel and for the extrapolation of the test results to flight condition. Within the framework of the prora-usv Program numerical activities have been carried out and others are currently in progress in order to characterize the aero-thermal environment that the FtB-2 vehicle will experience during the Sub-Orbital-Reentry test (SRT). In particular, CFD simulations of the flow field surrounding the fore- body part of the vehicle have been performed in correspondence of the maximum heat flux trajectory point that in the case of the SrT mission take place at an altitude of about 20Km at a Mach number of about 7.5. At this low altitude and relatively low speed the high heat flux value over the FtB-2 sharp nose(blunted cone with Icm radius of curvature) is mainly due to high pressure effects rather than to high enthalpy ones. Unit Reynolds number for the above trajectory point is about 20 so that the boundary layer will be turbulent for the most part of the vehicle surface and transition will proba ur immediately downstream the sphere-cone junction. For this reason CFD simulation have been performed lent assumptions in order to provide realistic and conservative heat flux distribution 35.6 FIGURE 2. Flow Field Around the FTB-2 Nose Iso-Contour of Mach Number The boundary layer state (laminar or turbulent) is another key factor for the on ground testing(Plasma Wind Tunnel usually are capable of low unit Reynolds). In figure 2 the flow field surrounding the first part of the FTB-2 vehicle is shown in terms of Mach Number iso-contours. It is evident how, due to the low radius of curvature of the nose the shock is very close to the body surface. Sharp nosed vehicle are commonly characterized by flying at low angle of attack along the trajectory being this a crucial factor to gain aerodynamic efficiency and than cross-range 132
the flight one, in the former case a large amount of energy is frozen within the fluid as formation enthalpy of dissociated atomic species. If, for instance, the material has a partially catalytic behaviour it is essential to be able to properly characterize the difference between the flight and ground environment in order to better understand which mechanism of heat release to the wall surface prevails: conductive or chemical. Moreover, when the article to be tested is scaled and/or of slightly different shape with respect to the real one, the correct reproduction of heat flux and pressure at the stagnation point do not assure in general that we are simulating the same environment downstream of the stagnation point. In the same way, the small radius of curvature of sharp structures together with the low Reynolds flow obtainable with an arc-jet facility may give rise to undesired rarefaction effects. All of this issues strongly claims for an extensive use of CFD both for the extrapolation from simulated flight condition to suitable operating condition of the plasma wind tunnel and for the extrapolation of the test results to flight condition. Within the framework of the PRORA-USV Program numerical activities have been carried out and others are currently in progress in order to characterize the aero-thermal environment that the FTB-2 vehicle will experience during the Sub-Orbital-Reentry test (SRT). In particular, CFD simulations of the flow field surrounding the forebody part of the vehicle have been performed in correspondence of the maximum heat flux trajectory point that in the case of the SRT mission take place at an altitude of about 20Km at a Mach number of about 7.5. At this low altitude and relatively low speed the high heat flux value over the FTB-2 sharp nose (blunted cone with 1cm radius of curvature) is mainly due to high pressure effects rather than to high enthalpy ones. Unit Reynolds number for the above trajectory point is about 20 millions so that the boundary layer will be turbulent for the most part of the vehicle surface and transition will probably occur immediately downstream the sphere-cone junction. For this reason CFD simulation have been performed with fully turbulent assumptions in order to provide realistic and conservative heat flux distribution. FIGURE 2. Flow Field Around the FTB-2 Nose. Iso-Contour of Mach Number. The boundary layer state (laminar or turbulent) is another key factor for the on ground testing (Plasma Wind Tunnel usually are capable of low unit Reynolds). In figure 2 the flow field surrounding the first part of the FTB-2 vehicle is shown in terms of Mach Number iso-contours. It is evident how, due to the low radius of curvature of the nose, the shock is very close to the body surface. Sharp nosed vehicle are commonly characterized by flying at low angle of attack along the trajectory being this a crucial factor to gain aerodynamic efficiency and than cross-range 132
capabilities. For the FTB-2 vehicle in correspondence of the maximum heat flux trajectory point the angle of attack is only 4deg. As a consequence of this the stagnation point lays on the spherical part of the nose and the shape of the heat flux distribution is quite similar between the windside and the leeside In figure 3 the heat flux profile derived from a full three dimensional computation (4deg of angle of attack) is compared with that obtained with a 2D axysimmetric simulation(Odeg angle of attack). It is clearly evident that, apart from the spread(that is emphasized by the turbulent state)due to the angle of attack, the axysimmetric distribution is a good approximation of the real situation In particular, over the sphere the distributions are identical. This result allowed to consider for the preliminary nose concept design a sphere-cone geometry subject to axisymmetric heat loads. The advantage of this formulation evident in terms of computational time required both for the thermo-structural analysis and for the determination of the heat loads along the trajectory, especially in an initial project phase that may require a parametric screening of fifteen 1.1E+06 Turbulent- Axi symm DE+06 80E+05 DE+05 60E+05 50E+05 40E+05 30E 20E+05 X(m). Distance from nose apex. FIGURE 3. Heat Flux Profiles with Radiative Equilibrium Hypothesis. Turbulent. The anal ysis of the flight environment allow to identify the most critical condition that the material has to withstand and that has to be reproduced in wind tunnel testing. In this framework it is foreseen to perform in SCirOCCO a series of test on a full scale nose model with the aim at reproducing the same maximum heat flux occurring along the trajectory. At the moment CFD analysis has been focused on the design of a PWT test to be performed on a scaled nose model with a stagnation point heat flux lower than the target value Starting from the value of heat flux to be realized at the stagnation point, theoretical-numerical activities have been conducted in order to aid the set up of the wind tunnel operating conditions. As a matter of fact, in order to properly execute the test, it is necessary to know the value of heat flux and pressure to be realized on a calibration hemispherical (10cm dia. probe made of copper and cooled at a constant temperature of about 50C. when the desired conditions are obtained over the probe this is extracted and the model is injected into the plasma flow and the test take place for the desired time. Therefore, aim of the numerical activities in this phase is the translation of the heat flux requirements over the model to be tested into operating conditions for the calibration probe. This process is influenced by several factors that cause differences between the probe and the test article: 1) different shape, 2) different positioning within the test chamber. Due to the effects of the nozzle expansion the conditions 133
capabilities. For the FTB-2 vehicle in correspondence of the maximum heat flux trajectory point the angle of attack is only 4deg. As a consequence of this the stagnation point lays on the spherical part of the nose and the shape of the heat flux distribution is quite similar between the windside and the leeside. In figure 3 the heat flux profile derived from a full three dimensional computation (4deg of angle of attack) is compared with that obtained with a 2D axysimmetric simulation (0deg angle of attack). It is clearly evident that, apart from the spread (that is emphasized by the turbulent state) due to the angle of attack, the axysimmetric distribution is a good approximation of the real situation. In particular, over the sphere the distributions are identical. This result allowed to consider for the preliminary nose concept design a sphere-cone geometry subject to axisymmetric heat loads. The advantage of this formulation is evident in terms of computational time required both for the thermo-structural analysis and for the determination of the heat loads along the trajectory, especially in an initial project phase that may require a parametric screening of different concepts. FIGURE 3. Heat Flux Profiles with Radiative Equilibrium Hypothesis. Turbulent. The analysis of the flight environment allow to identify the most critical condition that the material has to withstand and that has to be reproduced in wind tunnel testing. In this framework it is foreseen to perform in SCIROCCO a series of test on a full scale nose model with the aim at reproducing the same maximum heat flux occurring along the trajectory. At the moment CFD analysis has been focused on the design of a PWT test to be performed on a scaled nose model with a stagnation point heat flux lower than the target value. Starting from the value of heat flux to be realized at the stagnation point, theoretical-numerical activities have been conducted in order to aid the set up of the wind tunnel operating conditions. As a matter of fact, in order to properly execute the test, it is necessary to know the value of heat flux and pressure to be realized on a calibration hemispherical (10cm dia.) probe made of copper and cooled at a constant temperature of about 50°C. When the desired conditions are obtained over the probe this is extracted and the model is injected into the plasma flow and the test take place for the desired time. Therefore, aim of the numerical activities in this phase is the translation of the heat flux requirements over the model to be tested into operating conditions for the calibration probe. This process is influenced by several factors that cause differences between the probe and the test article: 1) different shape; 2) different positioning within the test chamber. Due to the effects of the nozzle expansion the conditions X (m). Distance from nose apex. 133