Sharp Hot Structures Project Current Status

IAC-06-C2.4.05 SHARP HOT STRUCTURES PROJECT CURRENT STATUS A. Del Vecchio, M. Di Clemente, M. Ferraiuolo, R. Gardi, G. Marino, G. Rufolo, L. Scatteia Downloaded by ITALIAN CENTER FOR on September 24, 2015 | http://arc.aiaa.org | DOI: 10.2514/6.IAC-06-C2.4.05

CIRA – Centro Italiano Ricerche Aerospaziali, Capua, ITALY

ABSTRACT This paper reports on the latest results obtained in the Sharp Hot Structures (SHS) project, that encompasses multidisciplinary activities aimed at developing the know-how needed for the manufacturing of advanced, high performance hot structures based on UHTC. The project is conducted within the frame of the Italian national Unmanned Space Vehicle (USV) program. At this regard, system and mission requirements coming from the second generation of the USV Flying Test Bed, denominated FTB-X, will be presented and discussed, together with aerothermodynamics calculations results. The thermomechanical design of a relevant technological demonstrator based upon the coupling of different classes of materials is presented, together with basic characterization results for the selected UHTC composition and an assessment of Italian national manufacturing capabilities for complex shapes and/or parts. Also scope of this paper is to discuss the issues related to the machining of UHTC and to the mechanical coupling of UHTC parts to conventional ceramic composites and or/ metal structures. Finally, the set up of a representative on-ground test for the technological demonstrator using CIRA Scirocco Plasma Wind Tunnel will be described.

1 - INTRODUCTION The Sharp Hot Structures (SHS) is a multidisciplinary project focused on the development, manufacturing, and on-ground testing of structurally efficient re-usable thermal protection systems for re-entry vehicles. The key features under study are: -

Innovative hot structures geometry, mainly based upon sharp profiles, intended to provide several advantages with respect to currently employed blunt shapes

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Mechanical coupling of different base materials, to exploit the strong points of each constituent material

The aforementioned steps are intended to lead to an extended operational range in terms of temperature, heat flux, and time of exposure with respect to the state of the art in the field of hot structures. The final objective of the SHS project is to provide technological demonstrators (identified as critical parts of re-entry vehicles such as nose cap and wing leading edges) to be qualified on-ground, and then tested and validated in flight conditions, in order to provide a definite assessment of the concepts under study.

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Downloaded by ITALIAN CENTER FOR on September 24, 2015 | http://arc.aiaa.org | DOI: 10.2514/6.IAC-06-C2.4.05

Within this frame, the development of a technology demonstrator of the nose cap started five years ago funded by the Italian National Aerospace Research Program PRO.R.A., while a feasibility study (Phase A) on the wing leading edges was launched two years ago with the support of the Italian Space Agency (ASI). The design of these components meets the high level requirements of the re-entry Flying Test Bed (FTB-X) and at the same time provides experimentation requirements to be met by the vehicle in flight conditions. The proposed technological concepts are mainly based upon the use of Ultra High temperature Ceramic materials and their coupling with conventional ceramic matrix composites (CMC) and metals: a dedicated research campaign has been conducted and is still ongoing on modified metal diboride compounds, widely recognized as the sole materials that can be conveniently employed at temperatures above 2200K [1]. SHS project activities are performed within a national research network, managed by CIRA, and involving the University of Naples “Federico II”, Centro Sviluppo Materiali S.p.A. (CSM), the 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 the University of Turin. This paper reports upon the current status of the development of one of the two aforementioned technological demonstrators, namely the Nose Cap.

This Orbital Re-entry Test mission shall be considered as the reference mission of the FTB_X vehicle. It shall consist of a complete re-entry flight from LEO orbit at 200 Km (to be confirmed by the launcher capability). An improved gliding re-entry and a high manoeuvring capability, as compared to the reference re-entry vehicle Space Shuttle, characterised by moderate angle of attacks (down to and below than 20o) and a longer flight duration (up to 1÷3 hours), shall be developed in order to allow for more extended in-flight testing capabilities in high energy hypersonic flight conditions. The LEO orbit inclination as well as the landing site and the re-entry trajectory footprint shall be suitably selected, in order to fulfil the safety requirements for ground population as per later requirement. The Sub-orbital Re-entry Test (SRT )type of mission of the FTB_X vehicle, may be envisaged in the mission plan as an intermediate step for both design validation and risk mitigation purposes. Indeed, the mission plan shall be conceived in order to gradually achieve the ORT reference mission capability, as above defined, according to a possible logical in-flight testing sequence, each mission representing a well demonstrated step toward the final mission. Mandatory requirements for the SRT mission to be eligible for such purposes are: a minimum acceptable total enthalpy greater than 5 MJ/Kg, in order to allow for aerothermodynamics non-equilibrium phenomena to occur; aero-thermo-mechanics characteristics fully representative of those of the reference mission, in order to progressively validate in flight critical design aspects.

2 – SYSTEM REQUIREMENTS The reference system requirements of the current iteration of the Nose Cap demonstrator are based upon the Unmanned Space Vehicle Flying Test Bed X, currently in the phase A of its development. Two types of missions have been identified within the re-entry flight mission envelope of the USV program: Sub-orbital Re-entry Test and Orbital Re-entry Test.

Table 1 SRT re-entry conditions.

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3 – PROTOTYPE CONCEPT:


Figure 1 depicts a schematic of the component dubbed Nose cap 1. 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) UHTC massive conical tip produced by hot pressing.

thermo-structural behaviour of the model with respect to thickness variations; The global model is useful to provide boundary conditions for the local model and is adopted to verify the optimum configuration obtained by the local model. Table 2 shows damaged area percentages for Nose final design:

Table 2 Damaged area percentages for final design Nose Figure 1. Schematic of the nose cap scaled demonstrator

4 – THERMOSTRUCTURAL DESIGN

Graphite tensile criticality is located at the interface with C/SiC.

Nose_2 Optimization

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Base model

•

Local model

The aim of Nose optimization is to eliminate damaged areas detected in the Final Design of the Nose. It has been possible to make use of solutions that are not reliable for the moment. However this study can be useful in a future design of the Nose. Nose_2 Final Design has been considered as the preliminary optimized Nose_2.

•

Global model

Two critical aspects have to be solved:

Three complementary approaches have been used to define the architectural configuration of the Nose cap:

The base model is a simplified twodimensional model able to take into account complex boundary conditions such as temperature and displacement distributions; The local model is an axial symmetric two dimensional piece of the global model; it is used to determine the quantitative changes in

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Compressive stresses in the coating

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Tensile stresses in Graphite

In order to solve these critical aspects, three different solutions have been identified: •

Employment of the optimal properties of ZrB2-SiC coating

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Increase of the length of the massive cone to reduce the tensile stresses in the graphite

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Concurrent adoption of solution 1 and 2 to remove all the critical aspects


Tables 3, 4 and 5 summarize damaged areas for the three solutions mentioned:

Table 5 Damaged area percentages adopting solution 1 and 2

Table 3 Damaged area percentages employing optimal coating properties

Critical compression behaviour of ZrB2-SiC coating has been solved by adopting enhanced material properties. Increasing of massive cone length has eliminated critical behaviour of Graphite. The best configuration achieved in the Nose_2 optimization is the one characterized by a massive cone length of 100 mm with optimum coating properties. 5 – UHTC MATERIALS CHARACTERIZATION, UHTC COMPONENT MANUFACTURING, MECHANICAL ASSEMBLY WITH CMC DOME The largest part of the laboratory research at material level within the SHS project is focused upon the massive UHTC, since the properties of these materials are highly composition and processing dependant, and literature data on them are still incomplete. The UHTC composition for the manufacturing of the nose tip is the following:

Table 4 Damaged area percentages increasing the length of the massive cone In conclusion, ZrB2-SiC coating compressive breakage cannot be solved since there is a great difference between C/SiC and ZrB2-SiC coating coefficients of thermal expansion. Then its breakage can be solved only by introducing optimum material properties.

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ZrB2 + HfB2 + SiC + s.a

It has been selected after a trade-off performed on various compounds based mainly on ZrB2 and HfB2 with different reinforcing phases and sintering aids. The material has been characterized in its microstructure (using SEM coupled with EDS

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and XRD on fracture surfaces and polished cross sections), mechanical properties (from R.T. up to 1600°C), while characterization of its thermal properties are ongoing. As an example, figure 2 show the fracture surface as inferred by SEM: the picture shows an uniform microstructure, diboride grains with a regular shape, mean size < 5 µm and a controlled size dispersion, uniform distribution of the SiC particles, a prevailing intragranular fracture path.

mgs (µm)

HV1.0 (GPa)

E (GPa)

ν

5

18.2±0.5

506±4

0.128

Table 6 Basic mechanical properties of composites A and B KIc (MPa⋅m1/2)

KIc (MPa⋅m1/2)

RT

1500°C

4.08±0.75

3.43±0.02

σRT (MPa)

σ1500°C (MPa)

763±73

240±20

Table 7 Basic mechanical properties of composites A and B (continued) Figure 2 Fracture surface of the hot pressed UHTC Detailed microstructure analyses on polished surfaces (Figure 3) provided evidence also of the formation of (Zr,Hf)B2 solid solutions (Fig. 3). Beside SiC (darker features in Fig. 4), a core-shell configuration of the diboride matrix clearly appears. The core consists of ZrB2, whilst the shell is constituted by (Zr,Hf)B2 solid solution. Rarely HfB2 was found.

Figure 3 SEM image of a polished surface of UHTC The values of the mechanical properties are summarized in Table 6 and Table 7.

UHTC components production Another focus of the material research is represented by the machining of the UHTC sintered pieces into final components. The use of Electrical Discharge Machining (EDM), and its effects upon the surface properties of the components are under investigation. Ceramic prototypes of the nose-tip were manufactured from hot-pressed cylinders. The final shape (figure 4) was obtained by electrical discharge machining. The EDM process introduced surface damage on the treated surfaces. Typical signatures of some local melting are shown in figure 8. The SEMEDS analyses upon the surfaces processed by EDM confirmed the formation of a partially oxidized thin layer, which in turn includes also contaminants like Cu and Zn from the electrically active tools used for EDM. In addition, the application of the EDM technique leads to a decrease of the material flexural strength, due to the micro-cracked surface layer induced by the action of the electrical discharges. In spite of those minor hindrances, whose effects can be taken into account and controlled in the design of a thermal structure, the employment of the EDM technique has proven to be effective for shaping a mediumscaled and complex component from a sintered ceramic billet.

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Figure 4 Nose tip prototype

Mechanical assembly of UHTC nose tip with CMC dome With regard to the connection between the massive cone and the dome, the use of a coupling pin is necessary since adhesive ceramics present several attaching difficulties. The design variables considered in the thermal-structural design of the coupling pin are: pin length, hole length, pin and hole shapes and diameters. The coupling pin is made of the same material of the massive cone to avoid the coming up of thermal stresses caused by the mismatch between coefficients of thermal expansion; however, metallic coupling pin will be analysed because of the reduced manufacturing costs. The results of this work will be useful to prepare a prototype of the Nose cap concept that will be on-ground tested into the arc-jet Plasma Wind Tunnel, available at CIRA. Ansys commercial code, employing finite element modelling, has been adopted in order to perform non-linear contact analyses. The concept of the coupling pin is described in figure 5. Simulation performed have demonstrated that contact pressures between the pin and the hole are not critical for the structure; the only one criticality is given by the compressive stresses. The configuration chosen is a mix between the third and the fourth simulation in order to couple small compressive stress values together with small hole dimensions.

Figure 5 Pin concept 6 – COMPUTATIONAL FLUID-DYNAMIC ANALYSIS 6.1 – Flight conditions simulation The starting point for the design of a suitable test to be realized in a ground facility as the PWT “Scirocco” has to be the characterization of the aero-thermal environment that the vehicle will experience during the reference re-entry mission. 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 Sub-Orbital Reentry Test to be performed with the CIRA USV vehicle should take place at an altitude of about 20Km and a Mach number of about 7.5. At this low altitude and relatively low speed the high heat flux value over the sharp nose (blunted cone with 1 cm 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 simulations have been performed with fully turbulent assumptions in order to provide a conservative heat flux distribution over the model that has been used for the design of the Nose_1 demonstrator. In figure 6Errore. L'origine riferimento non è stata trovata. the flow field surrounding the first part of the vehicle is shown in terms of Mach number iso-contours. It is evident how, due to the low radius of curvature of the nose,

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Figure 7 Heat flux over the spherical nose in laminar and turbulent regimes. Cold Wall Eq. Rad - Emiss=0.85 100000

Figure 6

90000 80000

Mach number iso-contours around the nose

70000 60000

In figure 7Errore. L'origine riferimento non è stata trovata. a comparison between cold wall heat flux distribution over the sphere region with turbulent or laminar boundary layer assumption is reported. The ratio between the maximum value and the stagnation point heat flux is 1.85. As already said, in order to be conservative for the structural design the fully turbulent profile has been adopted. However it has to be remarked that it is strongly improbable that transition occurs before the sphere-cone junction so that the rise shown in Errore. L'origine riferimento non è stata trovata. can result to be extremely conservative. Moreover, being the Reynolds number obtainable in PWT much lower than the flight one it is clear that this kind of heat flux distribution over the sphere cannot be reproduced during the test in Scirocco. In order to have a time history of heat flux along the trajectory, the above reported axissymmetric distribution has been normalized with respect to the stagnation point value. Then, the temporal history of heat flux at the stagnation point along the trajectory has been derived by means of simplified Fay-Riddell formula (see figure 8Errore. L'origine riferimento non è stata trovata.).

Altitude (m)


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 capabilities.

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Figure 8 Stagnation Point heat flux time history for SRT trajectory

7 – ARC-JET TESTING The SHS Nose 2 assembly will be tested in the most powerful arcjet plasma facility, the Scirocco Plasma Wind Tunnel. The SCIROCCO arc-jet facility presents some peculiarities that make it unique in the world. In particular, the arc heater length and diameter are respectively of 5500mm and 110mm and its maximum power is 70MW. The details of the facility design, the characteristics of the various components and the expected performances of the Scirocco PWT are described in (1,2,3). Briefly, the SCIROCCO facility is a typical segmented-constrictor Arc-Jet Wind Tunnel (AWT). The gas used for the tests is dry compressed air with mass flow variable from 0.1 to 3.5kg/sec. It accelerates through a convergent-divergent conical nozzle with

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The thermocamera used is the Agema Thermovision THV900LW. It is equipped with a Cd-Te-Hg detector scanning and its field of view holds in the IR waveband 8÷12µm with a frame of 136 not-interlaced lines in 1/15s. The thermal image is digitized with a resolution of 272x136 pixels that, in combination to a 10°x20° IR lens placed at the distance of about 4m from the test sample, results in a spatial resolution of about 1cm/pixel. The thermocamera is mounted in a protected box on a robotized remotely controlled gear, in order to avoid any disturbance from ambient radiation. In front of the thermo camera a ZnSe IR windows with a special coating to reduce the band pass wavelength region to the proper values of about 8-10 m is installed.

Stagnation Pressure Ps, mbar

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Figure 9 Stagnation heat flux transversal distribution 1200 2

The experimental set of measurements available is divided in the standard aerothermodynamic parameters measured by intrusive sensors such as heat flux and pressure transducers for the stagnation point over a copper cooled hemispherical probe and pressure levels along the test leg (nozzle, test chamber and diffuser wall static pressure) and optical not-intrusive infrared temperature measurement by using pyrometers (one and dual color) and thermocameras. An extended comparison between experimental aerothermodynamic parameters and CFD computed values is reported in the following paragraph. In Fig. 9 and 10 the stagnation heat flux and pressure transversal distributions are reported in order to demonstrate the high uniformity and extension of the core region of the plasma flow impinging the test article. For what concern the pyrometers, the dual color versions are not only capable to measure the temperature without the knowledge of the emissivity but are switchable between one and two colors making possible a correct and direct estimation of the emissivity itself in a range temperature from 300° C to 2500° C.

A preliminary test by using a Nose mock up in order to check the spatial temperature resolution and the applicability of all chosen diagnostic techniques has been realized. In Fig. 11 a typical image from the thermograph system is reported confirming the optimal configuration in terms of filed of view, window coating, accuracy and geometrical set up.

Stagnation Heat Flux qs, kW/m


interchangeable exit diameters up to 1950mm. The flow velocity at the nozzle exit can reach a value of 7000 m/s. During a test, after the stable flow condition is, in the test chamber, attained and confirmed by probe sensor, the testing model is put into the flow by the support system with characteristic time of 2÷5sec.

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Figure 10 Pressure transversal distribution

Figure 11 Typical thermograpy picture of the nose cap dummy model

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7.2 – Extrapolation from flight and PWT operating conditions The on-ground high enthalpy facilities do not allow the simultaneous experimental reproduction of all the thermo-fluid-dynamics conditions that characterize above all the lowearth orbit part of a typical space vehicle reentry trajectory. Even in the case of the CIRA Plasma Wind Tunnel “Scirocco”, it may be difficult to contemporary reproduce both stagnation point heat flux and pressure (remember that Scirocco class facilities are specifically designed to duplicate contemporary both the stagnation point heat flux q and Ps typical of a lifted re-entry trajectory). Moreover, due to the strong dissociation of the flow occurring trough the arc-heater and the expansion through the nozzle, the environment reproduced within the arc-jet facilities can be quite different from the real one. Even if the energetic level of the flow within the arc jet is the same of the flight one, in the former case a large amount of energy is frozen within the fluid as formation enthalpy of dissociated atomic species. For this reason, 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. 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 real flight conditions because of differences of the Reynolds and Mach number; 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 rarefaction effects. For the above reasons an extensive use of CFD is required both for the extrapolation from simulated flight condition to suitable operating condition of the plasma wind tunnel (extrapolation-from-flight) and for the extrapolation of the test results to flight conditions (extrapolation-to-flight). Starting from the value of heat flux to be realized during the experimental test at the model stagnation point, numerically estimated in a certain point of the re-entry trajectory,

theoretical-numerical activities have to be conducted in order to aid the set up of the facility driving conditions for on-ground qualification of the demonstrator. In order to properly execute the test, the values of heat flux and pressure to be realized on the hemispherical copper calibration probe of the facility (10cm diameter.) cooled at a constant temperature of about 50° C, have to be determined to be used as setting point of the facility itself. The aim of the numerical activities in the first phase is the derivation of heat flux and pressure values to be realized on the calibration probe of the facility starting from the requirements given on the model. The process of “translation” of the requirements given on the model to those on the calibration probe, that is necessary for the facility operating conditions definition, is influenced by several factors that cause differences between the probe and the test article that has to be taken under consideration: different shape different positioning within the test chamber: due to the effects of the nozzle expansion, the conditions along the axial direction are not uniform different wall temperature condition: wall temperature of the cooled probe is constant in time and uniform in space, whereas the temperature 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 different catalytic behaviour between the copper probe (fully catalytic) and the test article (finite rate catalysis) Starting from the requirement of the test given in terms of wall heat flux to be realized at the model stagnation point, PWT operating conditions that assured this requirement have been determined through a detailed CFD activity that has taken into account several aspects of the problem (position of the model into the test chamber, catalysis, wall temperature). For the same conditions, the values of heat flux and pressure on the

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• H0=13.7 MJ/kg that is the value estimated before the test • P0= 6.7 bar that is the value measured during the test

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Figure 12 Calibration probe heat flux

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Pressure [Pa]


hemispherical calibration probe have been determined to be used as driving conditions for the test. For the requirement corresponding to a model stagnation point heat flux equal to 2440 kW/m2, PWT operating conditions (H0=14.2 MJ/kg, P0=5.1 bar) have been numerically determined to perform an experimental test on a preliminary model. For the above determined operating conditions, a stagnation heat flux of 1090 kW/m2 and a stagnation pressure of 18 mbar on the hemispherical probe have been estimated: these two value have been used as target point for the execution of the test. Obviously the effectively measured heat flux and pressure over the calibration probe could be slightly different with respect to the predicted ones, then it is necessary to determine the reservoir conditions of enthalpy and pressure that allow to realize, from a numerical point of view, the measured values of the parameters over the probe. At the end of this retuning process, the following operating conditions has been considered:

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The results in terms of heat flux and pressure over the hemispherical probe obtained with these reservoir conditions are reported in figure 12 and 13. In the following Errore. L'origine riferimento non è stata trovata. it is reported the comparison between experimental and numerical values at the hemispherical probe stagnation point.

Heat Flux [kW/m2] Pressure [Pa]

Exp.

CFD

∆ (%)

1130

1144

-1.2

1850

1810

2.1

Table 5 Comparison between experimental and numerical values at probe stagnation point

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Figure 13 Calibration probe wall pressure

8 – CONCLUSIONS The analyses of the system requirements related to the USV-X orbital re-entry mission served as input for the design of a nose-cap technological demonstrator based upon the use of innovative UHTC materials and their coupling with conventional CMC structures. The base UHTC materials were selected at the end of an extensive trade-off campaign performed on different compositions. A prototype of sharp-shaped nose cap based upon the coupling of a C/SiC dome (produced by Fabbricazioni Nucleari) coated with a UHTC protective layer (produced by Centro Sviluppo Materiali), and mechanically

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assembled to a UHTC tip (National Research Council Institute for Ceramic Materials) is currently under production.

the 3th International Symposium on Atmospheric re-entry vehicles and systems, Arcachon, France, 24-27 March, 2003.

Computational fluid-dynamic analyses have been conducted in order to se-up an onground test for the final assembled prototype into the SCIROCCO Plasma Wind Tunnel.

[9] Del Vecchio A., Cardone G., IR thermographic measurements of temperatures in hypersonic large-scale plasma flow, 12th AIAA International Space Planes and Hypersonic Systems and Technologies, AIAA paper 2003-6926


REFERENCES [1] Beherens, B., “Technologies for Thermal Protection Systems Applied to Reusable Launchers” in Proceedings of the 54th International Astronautical Congress, Bremen, 2003. [2] McKenzie P., “Lockheed Martin Orbital Spaceplane Program” in Proceedings of the 54th International Astronautical Congress, Bremen, 2003. [3] Monteverde F., Bellosi, A., “Advanced Diboride Ceramics”, Scripta Materalia, 46, 223-228, (2002). [4] Levine S. R., Opila E. J., Halbig M. C., Kiser J. D., Singh M., Salem J. A., “Evaluation of Ultra-High Temperature Ceramics for Aeropropulsion Use”, Journal of the European Ceramic Society, 22, 27572767, (2002). [5] Valente T., Bartuli G., Visconti G., Tului M., “Plasma Spray Deposition” in Thermal Spray Surface Engineering via Applied Research, edited by Berndt C., ASM International Material Park, OH, 2000, pp. 837-841. [6] De Filippis, F., Caristia, S., and Del Vecchio, A. ,The Scirocco facility: Qualification phase, 2nd International Simposium on Atmospheric Reentry Vehicles and Systems, Arcachon, France, 26-29 March 2001. [7] De Filippis, F., Caristia, S., Del Vecchio, A., and Purpura, C., Scirocco final tests measured data: comparison between theory and experiments, Proceedings of the Fourth European Symposium on Aerothermodynamics for Space Vehicles, Capua, Italy, 15-18 October, 2001. [8] De Filippis, F., Caristia, S., Del Vecchio, A., and Purpura, C., The Scirocco PWT Facility Calibration Activities, Proceedings of

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