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
along the axial direction are not uniform; 3) different wall temperature condition. Wall temperature of the cooled probe is constant in time and uniform in space, while that of the test article comes out from the balance of heat convected from the fluid towards the surface. heat radiated from the surface towards the fluid and heat conducted into the solid. By neglecting the latter contribution a radiative brium assumption is made that allow to decouple the external flow field simulation from the thermal computation inside the solid. 4)different catalytic behaviour of the copper probe( fully catalytic) and of the test article(finite rate catalysis) In figure 4 and figure 5 the results of the Cfd simulations of the scirocco nozzle flow and of the flow surrounding the test article are respectively shown. The above mentioned computations have been performed with conditions. Results of this anal ysis in terms of heat flux distribution over the scaled nose geometry have constituted an input for the thermo-structural analysis that will be illustrated in a following paragraph NOZZLE Type F·Ho=44MJKg·Po=3Bba 「「「「目 mac:0.511.522533544.555566.577.588599.5 FIGURE 4. CFD Simulation of SCIROCCO NozZle Flow FIGURE 5. Scaled Nose PWT Test-Article Iso-Contours Mach Numbers 134
along the axial direction are not uniform; 3) different wall temperature condition. Wall temperature of the cooled probe is constant in time and uniform in space, while that of the test article comes out from the balance of heat convected from the fluid towards the surface, heat radiated from the surface towards the fluid and heat conducted into the solid. By neglecting the latter contribution a radiative equilibrium assumption is made that allow to decouple the external flow field simulation from the thermal computation inside the solid.; 4) different catalytic behaviour of the copper probe (fully catalytic) and of the test article (finite rate catalysis). In figure 4 and figure 5 the results of the CFD simulations of the SCIROCCO nozzle flow and of the flow surrounding the test article are respectively shown. The above mentioned computations have been performed with the CIRA code H2NS that is capable of solving the Navier-Stokes equation in thermo-chemical non-equilibrium conditions. Results of this analysis in terms of heat flux distribution over the scaled nose geometry have constituted an input for the thermo-structural analysis that will be illustrated in a following paragraph. FIGURE 4. CFD Simulation of SCIROCCO Nozzle Flow. . FIGURE 5. Scaled Nose PWT Test-Article. Iso-Contours Mach Numbers. 134
THERMO-STRUCTURAL ANALYSIS The availability of the PWT facility, a unique high enthalpy hypersonic wind tunnel for testing materials and tructures for all possible configurations of space transportation systems, led to a specific interest in the validation of the thermo-structural analyses performed within the Italian Unmanned Space Vehicle (USV)-Sharp Hot Structure (SHS) project In fact, this facility provides the capability to simulate surface pressures, wall temperatures and flow duration over rge scale models in the most critical portion of the re-entry path The PWt facility will be used to validate the thermo-structural analysis methodology. The correct estimation of the temperature distribution over the external wall may be verified during the test. The case of study presented here is referred to the scaled nose cone. Moreover, a Safety Analysis is required before running any PWt test in order to guarantee that no catastrophic failure may occur. In particular, the maximum value of heat flux to run the tests is required to be determined. So, the Safety Analysis will be used as a case of study to investigate the correctness of the aero-thermo-structural modelling methodology by using PWt infrared cameras to acquire wall temperatures to compared with the nodal temperatures obtained herein. Thermostructural Analyses have been performed using two distinct phases 1. preliminary parametric thermo-structural anal yses(with no updating of the heat flux on the basis of the heating of the external wall) varying the heat flux within a specified feasible interval to determine the 2. detailed thermo-structural analysis with an updating subroutine to determine the correct amount of heat flux entering the body(referred to the maximum safe heat flux value determined before) In the first phase, CFD-calculated cold wall heat flux values have been applied as a transient boundary condition to the external wall, implementing the conventional built-in radiation boundary condition of the Fe code. This approach tends to be conservative because the heat flux will generally decrease while wall temperatures increase. On the other side, this approach is more computationally effective. So, this methodology was used varying the heat flux to quickly determine the maximum safe heat flux value, according to the selected Failure Criterion(see the rest of the paragraph for details). However, for the aim of comparing FE-estimated wall temperatures with experimental data that will be acquired the next step of the sHs project, an updating procedure was needed to correctly apply the heat flux load. So, once the maximum safe heat flux value was set, further CFD analyses were performed to get a more accurate estimation of cold wall heat flux. These analyses were based on the wall nodal temperatures previously calculated during the mentioned first phase. For the second step of this work, proportionality coefficients h needed to be determined at ach location x and at each time 't(Eq. 1): h(x,t)=Ho-C, Twal ( (1) where Ho is the total enthalpy and Cp is the specific heat, qceD is the cold wall heat flux calculated for the assumed Tall temperature The heat flux load at time 'Icould so be updated, at each time step, on the basis of the FE-calculated wal temperature T walLFEM at the previous time step on the same node, as described below dx, t)=hx, Ho(t-CpTmwoFEM(xt-1)l with the second approach, heat flux values had to be applied on wall nodes in order to allow the updating procedure expressed in Eq. 2. So, an interpolation procedure was necessary to apply re-calculated heat fluxes(Eq. 2)onto the
THERMO-STRUCTURAL ANALYSIS The availability of the PWT facility, a unique high enthalpy hypersonic wind tunnel for testing materials and structures for all possible configurations of space transportation systems, led to a specific interest in the validation of the thermo-structural analyses performed within the Italian Unmanned Space Vehicle (USV)-Sharp Hot Structure (SHS) project. In fact, this facility provides the capability to simulate surface pressures, wall temperatures and flow duration over large scale models in the most critical portion of the re-entry path. The PWT facility will be used to validate the thermo-structural analysis methodology. The correct estimation of the temperature distribution over the external wall may be verified during the test. The case of study presented here is referred to the scaled nose cone. Moreover, a Safety Analysis is required before running any PWT test in order to guarantee that no catastrophic failure may occur. In particular, the maximum value of heat flux to run the tests is required to be determined. So, the Safety Analysis will be used as a case of study to investigate the correctness of the aero-thermo-structural modelling methodology by using PWT infrared cameras to acquire wall temperatures to be compared with the nodal temperatures obtained herein. Thermostructural Analyses have been performed using two distinct phases: 1. preliminary parametric thermo-structural analyses (with no updating of the heat flux on the basis of the heating of the external wall) varying the heat flux within a specified feasible interval to determine the maximum safe heat flux value to run the test in safe conditions; 2. detailed thermo-structural analysis with an updating subroutine to determine the correct amount of heat flux entering the body (referred to the maximum safe heat flux value determined before). In the first phase, CFD-calculated cold wall heat flux values have been applied as a transient boundary condition to the external wall, implementing the conventional built-in radiation boundary condition of the FE code. This approach tends to be conservative because the heat flux will generally decrease while wall temperatures increase. On the other side, this approach is more computationally effective. So, this methodology was used varying the heat flux to quickly determine the maximum safe heat flux value, according to the selected Failure Criterion (see the rest of the paragraph for details). However, for the aim of comparing FE-estimated wall temperatures with experimental data that will be acquired in the next step of the SHS project, an updating procedure was needed to correctly apply the heat flux load. So, once the maximum safe heat flux value was set, further CFD analyses were performed to get a more accurate estimation of cold wall heat flux. These analyses were based on the wall nodal temperatures previously calculated during the mentioned first phase. For the second step of this work, proportionality coefficients h needed to be determined at each location x and at each time ‘t’ (Eq. 1): ( ) ( ) H ( )t C T ( ) x t q x t h x t p wall CFD , , , 0 − ⋅ = & (1) where H0 is the total enthalpy and Cp is the specific heat, qCFD & is the cold wall heat flux calculated for the assumed Twall temperature. The heat flux load at time ‘t’ could so be updated, at each time step, on the basis of the FE-calculated wall temperature Twall,FEM at the previous time step on the same node, as described below: ( , ) ( , )[ ( ) ( , 1)] q& tx = h tx H0 t −CPTwall,FEM tx − (2) With the second approach, heat flux values had to be applied on wall nodes in order to allow the updating procedure expressed in Eq. 2. So, an interpolation procedure was necessary to apply re-calculated heat fluxes (Eq. 2) onto the 135
FEM grid. A comparison he total energy implemented Ho h in the Fe analysis and the CFd simulation was needed to guarantee the less of the total energy supplied to the structure. The comparison showed a difference of about 0.3 % 0 the total amount of energy Ho h supplied onto the CFd grid(solid line) and the one interpolated to match the FE-nodes(circles)(Figure 6), validating the interpolation procedure 10×10 h *H. on fe nodes cony hony Ho on CFD grid 2 0.1 0.15 0.2 Curvilinear Abscissa(m) FIGURE 6. Ho hoony on the CFD Grid (Solid Line)and on FE-Nodes(Circles). A FE model of the structure was implemented in the ANSYSTM code. The same FE model was used for both the preliminary and detailed thermo-structural analyses. The heat flux updating procedure for the second step was developed using the built-in programming language of the named code About 25,000 ANSYSM PLANE 55 iso-parametric quadrilateral finite elements for 2D axisymmetric problems were used. All materials were considered isotropic with the exception of the C/SiC that was considered transversely isotropic. Some of the material properties were considered temperature-dependent. Previous runs with coarser meshes showed that mesh effects had been eliminated. Several full transient thermal analyses were performed to determine the opportune time-step for integration. The uniform initial temperature To was considered to be room temperature and radiation boundary condition was implemented using the conventional built-in procedure of th specific FE code. The same mesh was converted into the corresponding structural PLANE 42 ANSYSTM elements run a sequentially-coupled thermo-structural analysis, considering previous FE-calculated temperatures as body loads. Reference temperature for thermal stress calculation was considered to be room temperature The Maximum Normal Stress Criterion was adopted as a Failure Criterion. This criterion is often used to predict the failure of brittle materials as Ultra High Temperature Ceramics (UHTC). According to the selected criterion, failure occurs when the maximum(normal) principal stress reaches either the uniaxial strength or the uniaxial compression strength. The Maximum Normal Stress Criterion has been preferred to the von Mises Stress Criterion because it distinguishes between Tensile and Compressive behaviour of the materials. Besides, the Von Mises Stress Criterion considers an equivalent stress which is an average of the stresses and does not take account of the difference between Compressive Strength and Tensile Strength. This distinction has been considered to be fundamental to characterize brittle materials like UhTCs Thermal results are shown in Figure 7 for the not updated(solid line) and the updated(dashed line) approaches Nodal temperatures on the external side of the structure vs. the cone axis are represented. Temperature distribution in this case of study revealed that the effects of the updating approach is more significant on the frontal side of the structure than on the lateral one, as shown in Fig. 8. The wall temperature distribution obtained with the updating procedure will be used in the next step of the shs project to be compared with experimental data acquired during 136
FEM grid. A comparison between the total energy implemented H0 h in the FE analysis and the CFD simulation was needed to guarantee the correctness of the total energy supplied to the structure. The comparison showed a difference of about 0.3 % between the total amount of energy H0 h supplied onto the CFD grid (solid line) and the one interpolated to match the FE-nodes (circles) (Figure 6), validating the interpolation procedure. FIGURE 6. H0 hconv on the CFD Grid (Solid Line) and on FE-Nodes (Circles). A FE model of the structure was implemented in the ANSYS™ code. The same FE model was used for both the preliminary and detailed thermo-structural analyses. The heat flux updating procedure for the second step was developed using the built-in programming language of the named code. About 25,000 ANSYS™ PLANE 55 iso-parametric quadrilateral finite elements for 2D axisymmetric problems were used. All materials were considered isotropic with the exception of the C/SiC that was considered transversely isotropic. Some of the material properties were considered temperature-dependent. Previous runs with coarser meshes showed that mesh effects had been eliminated. Several full transient thermal analyses were performed to determine the opportune time-step for integration. The uniform initial temperature T0 was considered to be room temperature and radiation boundary condition was implemented using the conventional built-in procedure of the specific FE code. The same mesh was converted into the corresponding structural PLANE 42 ANSYS™ elements to run a sequentially-coupled thermo-structural analysis, considering previous FE-calculated temperatures as body loads. Reference temperature for thermal stress calculation was considered to be room temperature. The Maximum Normal Stress Criterion was adopted as a Failure Criterion. This criterion is often used to predict the failure of brittle materials as Ultra High Temperature Ceramics (UHTC). According to the selected criterion, failure occurs when the maximum (normal) principal stress reaches either the uniaxial strength or the uniaxial compression strength. The Maximum Normal Stress Criterion has been preferred to the Von Mises Stress Criterion because it distinguishes between Tensile and Compressive behaviour of the materials. Besides, the Von Mises Stress Criterion considers an equivalent stress which is an average of the stresses and does not take account of the difference between Compressive Strength and Tensile Strength. This distinction has been considered to be fundamental to characterize brittle materials like UHTCs. Thermal results are shown in Figure 7 for the not updated (solid line) and the updated (dashed line) approaches. Nodal temperatures on the external side of the structure vs. the cone axis are represented. Temperature distribution in this case of study revealed that the effects of the updating approach is more significant on the frontal side of the structure than on the lateral one, as shown in Fig. 8. The wall temperature distribution obtained with the updating procedure will be used in the next step of the SHS project to be compared with experimental data acquired during 136
the PWT test in order to validate the updating methodology presented in this work 1100 1000 电E二 Updated Not Updated Cone Axis (m) FIGURE 7. Wall Temperatures on the Frontal Side of the Structure. To be Noticed the Difference in Wall Temperatures with (Dashed Line)and without (Solid Line)the Heat Flux Updating Procedure. ON GROUND ARC JET TESTING The SHS nose cap concept will be tested into the arc-jet Plasma Wind Tunnel that is available in CIRA, named test facility in the world(70 MW) and produces a very uniform and large test jet (up to 2 m diameter). In Figure 8 a schematic of Scirocco is reported. The process air is thermally energised into the segmented constricted Arc Heater reaching temperature values between 2000 and 10000 K. This energy is transformed in kinetic by the air through a convergent-divergent Conical Nozzle and an hypersonic test jet is generated with velocity between 2000 to 6000 m/s and Mach number between 6 to 12 depending on the exit nozzle size. Different with different length and exit diameter are available. After the air flow impact with the model in Test Chamber the jet is collected by a tube of 50 m length, Diffuser, goes through a Heat Exchanger, is sucked by a Vacuum System and is emitted into the atmosphere. The performances of the scirocco facility are represented in Figure 9 in terms of stagnation pressure and stagnation heat flux on a standard model (fully catalytic 300 mm radius semi-sphere) The test requirement is the achieving on the SHs model of a stagnation pressure of 1009 Pascal and a stagnation heat flux of 833 kW/m2. These data have been transformed in the equivalent ones to be obtained on the calibration probe that is used in Scirocco to verify that the requested flow conditions are reached. This probe is a cooled copper emi-sphere, considered fully catalytic, able to measure the stagnation pressure by a tap and the thermal load by a Gardon gauge sensor; it has a diameter of 100 mm, while the shs test model has a curvature radius at stagnation point of 13 mm. With the CFD support the flow requested conditions, specifically on the probe, have been calculated: 1) Stagnation Heat Flux of 277 kW/m2: 2) Stagnation Pressure of 947 Pascal. A deviation from these requirements of 10 per cent on the heat flux and of 5 per cent on the pressure will be admitted. The test setting of the facility is foreseen produce the experimental conditions reported in table 1
the PWT test in order to validate the updating methodology presented in this work. FIGURE 7. Wall Temperatures on the Frontal Side of the Structure. To be Noticed the Difference in Wall Temperatures with (Dashed Line) and Without (Solid Line) the Heat Flux Updating Procedure. ON GROUND ARC JET TESTING The SHS nose cap concept will be tested into the arc-jet Plasma Wind Tunnel that is available in CIRA, named Scirocco (Caristia, 2001; Russo, 2002; Caristia, 2003). This plant represents one of the most powerful hypersonic test facility in the world (70 MW) and produces a very uniform and large test jet (up to 2 m diameter). In Figure 8 a schematic of Scirocco is reported. The process air is thermally energised into the segmented constricted Arc Heater reaching temperature values between 2000 and 10000 K. This energy is transformed in kinetic by the air passage through a convergent-divergent Conical Nozzle and an hypersonic test jet is generated with velocity ranging between 2000 to 6000 m/s and Mach number between 6 to 12 depending on the exit nozzle size. Different nozzles with different length and exit diameter are available. After the air flow impact with the model in Test Chamber the jet is collected by a tube of 50 m length, Diffuser, goes through a Heat Exchanger, is sucked by a Vacuum System and is emitted into the atmosphere. The performances of the Scirocco facility are represented in Figure 9 in terms of stagnation pressure and stagnation heat flux on a standard model (fully catalytic 300 mm radius semi-sphere). The test requirement is the achieving on the SHS model of a stagnation pressure of 1009 Pascal and a stagnation heat flux of 833 kW/m2. These data have been transformed in the equivalent ones to be obtained on the calibration probe that is used in Scirocco to verify that the requested flow conditions are reached. This probe is a cooled copper semi-sphere, considered fully catalytic, able to measure the stagnation pressure by a tap and the thermal load by a Gardon gauge sensor; it has a diameter of 100 mm, while the SHS test model has a curvature radius at stagnation point of 13 mm. With the CFD support the flow requested conditions, specifically on the probe, have been calculated: 1) Stagnation Heat Flux of 277 kW/m2; 2) Stagnation Pressure of 947 Pascal. A deviation from these requirements of 10 per cent on the heat flux and of 5 per cent on the pressure will be admitted. The test setting of the facility is foreseen produce the experimental conditions reported in table 1. 137
CONTROL SYSTEM POWER SUPPLY EST CHAMBER HEAT AIR SUPPLY EXCHANGER TEST ARTICLE SUPPORT NOZZLES DIFFUSER SEGMEN ARC HEATER AUTOMATIC DATA ACQUISITIO WATER PUMPS STEAM EJECTOR FIGURE 8. Scirocco PWT Facility Schema Stagnation Pressure(mBar) 9. Scirocco PWT Performance Map 138
FIGURE 8. Scirocco PWT Facility Schema. 1 10 100 1000 0 200 400 600 800 1000 1200 1400 0 200 400 600 800 1000 1200 1400 Stagnation Heat Flux (kW/m 2 ) Stagnation Pressure (mBar) FIGURE 9. Scirocco PWT Performance Map. 138