IAC-14-D2.6.3 ESA INTERMEDIATE EXPERIMENTAL VEHICLE IN-FLIGHT EXPERIMENTATION. OBJECTIVES, EXPERIMENT IMPLEMENTATION, QUALIFICATION AND INTEGRATION

IAC-14-D2.6.3 ESA INTERMEDIATE EXPERIMENTAL VEHICLE IN-FLIGHT EXPERIMENTATION. OBJECTIVES, EXPERIMENT IMPLEMENTATION, QUALIFICATION AND INTEGRATION G. C. Rufolo*, F. Camarri, C. Pereira R. Romano, D. Vernani * CIRA/ASI Technical Assistance to European Space Agency, Paris, France, g.rufolo@cira.it Thales Alenia Space Italia, Italy, flavio.camarri@thalesaleniaspace.com RUAG Space, Zurich, Switzerland: carlos.pereira@ruag.com, romeo.romano@ruag.com, dervis.vernani@ruag.com The main objective of the ESA IXV Project, scheduled for launch in last quarter 2014, is the in-flight verification at system level of all the technologies deemed critical for the autonomous re-entry from Low Earth Orbit (LEO) with a particular focus on Aerodynamics, Aerothermodynamics and Thermal Protection System (TPS). IXV is fully instrumented with about 300 sensors with the aim at gathering valuable data to validate and improve the above mentioned technologies. The objective of the sensor definition and placement in the vehicle is to maximize the amount of data while minimizing number of sensors. The nose of the vehicle uses a total of twenty thermocouples and nine pressure sensors to capture angle of attack, stagnation pressure and thermal gradients of the C-SiC ceramic. Displacement sensors and high temperature strain gauges record mechanical loads and measure the differential expansion between the nose cap and the vehicle carbon composite structure. Similar combination of sensors is used in the windward TPS. A coated patch is applied on one of the tiles to quantify the catalytic level of the ceramic material in a re-entry environment. The leeward, lateral and base of the IXV vehicle are covered by ablative material. Its behaviour requires a dedicated design of the thermocouple measuring system. At the rear of the vehicle the aps and hinge are instrumented using thermocouples. An infrared camera with periscope optics based on a sapphire mirror generates thermal maps of the aps. Images are compressed for transmission and stored in the data handling unit which analyses them and sets gain and exposure of the camera. This work describes the implementation of the various experiment, the qualification of the adopted solution for the IXV environment, the integration on the vehicle and the final testing. I. INTRODUCTION In November 2014 the European Space Agency (ESA) Intermediate Experimental Vehicle (IXV) will be launched from European Space Port in Kourou (French Guyana) by means of a Vega launcher. After a ballistic arc IXV will reach the entry interface point at 120 km in conditions fully representative of a re-entry from Low Earth Orbit. The vehicle will autonomously guide through the Earth atmosphere by decreasing its speed from Mach 30 up to Mach 1.5 at the interface with the descent phase at about 24 km in which a three stage supersonic parachute system will provide the final deceleration up to a splashdown in the Pacific ocean and the subsequent recovery of the vehicle that will float thanks to a floatation system. At the moment of writing the paper the vehicle, fully integrated, has successfully completed the System environmental test campaign in ESTEC and it is ready for the shipment to the launch site [1]. The main IXV objective is the in-flight verification at system level of all the technologies deemed critical for the autonomous re-entry from Low Earth Orbit (LEO). Among such critical technologies a special attention is given to the Aerodynamics and Aerothermodynamics and the Thermal Protection System (TPS). A lack of in-flight validation for Europe in numerical tool for re-entry application implies the application of large uncertainties to the heat fluxes estimations that in turn impose to account for large margins in the design of TPS components with a consequent mass penalty. The same applies to the estimation of aerodynamic parameters for which the uncertainties are mainly driven by real gas effects, especially for what concern the control surface efficiency. In this scenario IXV will provide valuable data in an environment representative for most of the critical aerothermodynamics phenomena (velocity at interface point equal to 7.5 km/s) with the aim at obtaining a concrete and measurable reduction of uncertainties. For what concerns the TPS, the severity of the reentry environment induces several criticalities (e.g. thermal expansion at the interfaces, junctions, seals, gaps, steps and singularities) that make very complex IAC-14-D2.6.3 Page 1 of 14

the design of solution compliant with the mission and system requirements. This aspect is made even more critical by the lack of knowledge of the surface behaviour (i.e. catalysis and oxidation) of Ceramic Matrix Composite (CMC) components in re-entry environment. The exploitation of IXV flight data will allow the verification, and characterization and, eventually, the improvement of the design strategy for TPS (e.g. mass reduction, extended flight envelope). II. IN FLIGHT EXPERIMENTATION PLAN During re-entry, the Reynolds number progressively increases as outside air density rises, while the velocity decrease of the air is much slower so that a laminar-toturbulent transition occurs leading to a significant increase of the convection heat flux on the TPS surface. In order to identify the areas where the turbulent transition appears and develops, 34 Type S thermocouples are placed in longitudinal and spanwise arrays to detect temperature jumps. Six displacement sensors will also indicate the size of the steps that could trigger a transition in critical TPS zones (e.g. nose/windward zones). To pursue the above mentioned objectives, two main categories of experiments have been selected for IXV: 1. Aerodynamic and aero-thermodynamic (AED/ATD) experiments whose objectives are related to the investigation of aero-thermal phenomena, in particular: Flap surface efficiency and aerothermodynamics Shock-Shock Interaction Shock-Wave Boundary Layer Interaction (SWBLI) Turbulent heating Laminar-to-Turbulent Transition (LTT) Base aerodynamics and aerothermodynamics Rarefied and continuous aerodynamics Reaction Control System (RCS) efficiency Materials catalytic behaviour Transitional separation Cavity heating Real Gas Effects 2. TPS experiments, whose objectives are the validation of insulating capabilities and thermomechanical performance of the TPS during re-entry. The TPS consists of the carbon reinforced silicon carbide (C-SiC) nose cap and windward tiles, a cork based ablative on the leeward and lateral surfaces as well as base of the vehicle and a C-SiC hinge panel and body flap assembly to manoeuvre the vehicle. Fig. 1. Sensors map for LTT experiment. Red circle: Temperature measurement. Blu circle: displacement measurement. Materials catalytic behaviour (CATE) In presence of a highly dissociated flow, the surface properties modify the dissociation-recombination process within the boundary layer. The use of a catalytic surface, in this case a mullite layer applied on a low catalytic surface, the C-SiC material of the tile, will allow the detection of the typical temperature jump and in turn it will allow the confirmation of the finite rate catalysis behavior of the ceramic material and the relaxation of the fully catalytic assumption commonly used for the design. The experiment sensors consist of multiple temperature sensors placed downstream from the coated surface. A total of 30 Type S thermocouple will allow the characterization of the temperature increase across the patches (see Fig. 2) In the following paragraph a description of the various experiment conceived to address and characterize the above mentioned phenomena will be given. II.1. Aero-dynamic and Aero-thermodynamic Experiments Laminar-to-Turbulent Transition (LTT) Fig. 2. Sensors map for CATE experiment. Red circle: Temperature measurement. Gray. rectangle: catalytic patch. IAC-14-D2.6.3 Page 2 of 14

Shock Wave/Boundary Layer Interaction (SWBLI) and Flap ATD Experiment The scope of these experiments is to investigate the complex aerothermodynamics phenomena occurring on the flap surface and surrounding volume after the generation of the shock waves induced by flap deflection (see Fig. 3). In particular if during the reentry the commanded flap deflection will be high enough, the shock forming upstream the flap hinge will induce a separation and a subsequent reattachment shock. The characterization of this phenomenon is extremely important since it may have significant effect on the flap efficiency. General heating Experiment The scope of the experiment is to investigate the global heating at the surface of the vehicle during the hypersonic phase of re-entry thus covering rarefied and continuous flow as well as both laminar and turbulent regimes. The measurement employs a large number of thermocouples (see Fig. 5). Synergy with temperature measurements of other experiments is exploited to a major extent. Fig. 5. Sensors map for General Heating experiment. Red circle: Temperature measurement. Flush air data system (FADS) Experiment Fig. 3. Shock Wave/Boundary Layer Interaction in flap area (Courtesy: CFS Engineering) As shown if Fig. 4, 21 Type S thermocouples will provide a full coverage of the temperature evolution on the hinge/flap zone. In addition an infrared camera placed on the base of the vehicle will monitor the temperature evolution of the leeward side of the flap. Finally 4 pressure sensors placed upstream the flap will also provide indication of the separation extent. The purpose of the experiment is to derive vehicle attitude (Angle of attack, angle of sideslip) and atmospheric quantities starting from pressure and temperature measurement taken on the vehicle nose. The sensors are disposed in a cross array in order to maximize gradient readings (see Fig. 6). The FADS will not operate in real time but the relevant quantities will be extracted in post flight analysis confirming or not the possibility of using this kind of system in a wide range of flow regimes for future re-entry application. Fig. 4. Sensors map for SWBLI experiment. Red circle: Temperature measurement. Blu circle: Pressure measurement. Fig. 6. Sensors map for pressure related experiments. Red circle: Temperature measurement. Black circle: pressure measurement. IAC-14-D2.6.3 Page 3 of 14

Slip flow and Skin Friction Sensors (SFS) The scope of this challenging experiment is to assess the local aerodynamic characteristics at the wall both in the high altitude range corresponding to the transition regime between the free molecular flow and the continuum flow regimes and in the continuum flow. In particular the objective is to derive by means of differential pressure measurement an estimation of the slip flow velocity in transitional regime and the skin friction coefficient in continuum regime. The corresponding sensors are pressure sensors, one port with an aperture perpendicular to the surface and connected to both a differential and absolute pressure sensors inside the vehicle and the other port with inclined aperture connected to the second port of the differential sensor. A thermocouple is located in the vicinity of the ports. Areas of investigation are the front of the windward tiling and the middle of the vehicle (see Fig. 6). The experiment will require an extensive post flight calibration activities including also wind tunnel characterization. Jet-Flow Interaction The scope of this experiment is to address the interaction between the thruster-generated jet and the external flow field, in particular the phenomena of plume spreading and plume impingement on the wall. Thrusters efficiency is derived from the acceleration and angular rates measurements when the thrusters are activated. These measurements are performed in the vehicle control system. Moreover, the interaction of the plume with the vehicle surface in the base region is investigated by means of temperature measurements and pressure measurement perpendicular to the jets axis (see Fig. 7). Continuum flow Experiment The objective of this experiment is the investigation of Continuous Regime Aerodynamics, mainly the aerodynamic coefficients, the flap trimming deflection values and flow characterization through pressure mapping. The experiment uses all surface pressure measurements and in addition the aerodynamic coefficient derived from the vehicle model identification that will be performed as part of the GNC verification. High Altitude Aerodynamics The experiment deals with the estimation of aerodynamics coefficient in high altitude, rarefied/transitional regime. The estimation will be based on acceleration measurement performed with the IXV IMU. Even though the accelerations do not allow a direct derivation of aerodynamic coefficients, since the atmosphere density is not measured, an estimation is made possible deriving the density from indirect quantities like the ratio of normal vs axial force. Considering that the normal force coefficient changes are rather small over the transitional regime at the IXV re-entry AoA compared to changes in axial force coefficient, it allows a density estimation using the predicted normal force coefficient and normal acceleration. Consequently an estimation of axial force coefficient from axial acceleration measurement can be performed. Fig. 7. Sensors map for Base region experiment. Red circle: Temperature measurement. Blu circle: pressure measurement. Base Flow Field Experiment Numerical prediction of separated base flow both in laminar and turbulent regime is quite challenging. In particular the misprediction of pressure field in low supersonic regime may lead to significant error in the estimation of the drag coefficient. On the other hand in hypersonic regime, CFD tends to overestimate the heat flux in highly separated region. The scope of this experiment is to provide valuable flight data to validate computational tools. The sensors consist of thermocouples and pressure ports located in the base of the vehicle (see Fig. 7). Gaps and cavity heating The scope of this experiment is to investigate the local aerothermodynamics in cavities and the possible sneak IAC-14-D2.6.3 Page 4 of 14

flows that may be induced by pressure gradients under imperfect sealing. This experiment addresses in particular the heating of two cavities: the flap hinge cavity closed by a seal in depth and the open flap side cavity between the flap and the hinge TPS component. The experiment sensors consist in thermocouples. The pressure gradient information between windward and leeward sides can be obtained from the continuous flow and base flow field experiments. Infra-Red Camera Experiment This experiment aims to provide through an Infra- Red Camera an accurate mapping of real time temperature distribution in the flap area subjected to high gradient, rapid fluctuating aerothermodynamic effects such as, reattachment flows and trailing edge and flap side cavity heating. The thermocouples located on the left flap and on the base vicinity contribute to this experiment since they provide temperature references that can then be used for calibration of the IR measurements. A detailed description of the experiment setup is provided in par. IV.5. II.2. Thermal Protection System Experiments The thermal protection system experiments evaluate the performance of the various TPS systems used in the vehicle: ablative on leeward and lateral surface as well as the base and two different types of C-SiC used for the hotter parts of the vehicle: nose cap, rear hinge, flaps and vehicle windward tiling. particular positive steps result in local aerothermodynamic effects with increased thermal loading of the adjacent tiles. An objective of the in-flight instrumentation of IXV is to document the aerodynamic and thermal loads on the TPS including deflection and the evolution of steps along the vehicle windward tiling. To this end a combination of high temperature strain gauges and thermocouples will be placed at the metallic stand-offs behind the highest loaded tiles and on the nose cap attachments (see Fig. 9). The deflection at the edges of the tiles and the steps will be measured using linear variable differential sensors (LVDT). Fig. 9. IXV Nose-Cap with instrumented metallic standoff. A separate objective consists in verifying the performance of the C-SiC flaps in flight. This verification in achieved by use of a thermocouple grid on the flaps and an infrared camera with periscope optics at the rear of the spacecraft allowing thermal mapping of the leeward side of the flaps. Fig. 10. Temperature measurement locations for Flap and Hinge region. Fig. 8. Heating of IXV during Re-entry. (Courtesy: ESA Artist Rendition) The windward surfaces of the IXV and its nose cap are exposed to large thermal gradients and pressure loadings which result in changes to the surface topology and high transient loading of fixation elements. In The use of thermocouples placed at different depths in the ablative will allow determination of heat fluxes and ablative performance (see Fig. 11). The interface areas between different TPS types such as monolithic nose to tile, windward tile to lateral ablator, nose cap to ablator and between tiles will be IAC-14-D2.6.3 Page 5 of 14

investigated by placing sensors on both sides of the transition areas. The sensors will allow an investigation on the effect of surface mismatches and the resulting overheating. complete experimental datastream to two flight water proof solid state recorder; to send the same datastream to the experimental layer transmitter that by means of two dedicated antennas will send all the data on ground via telemetry. In addition to the flight recorder additional recording devices are placed directly into each of the four DAUs. In order to cope with black-out during re-entry phase, recorder data are send in playback mode to telemetry during the descent phase under main parachute thus to ensure full data retrieval before splashdown. Fig. 11. Temperature and pressure measurement points in ablative TPS area TPS TC Thermocouple Bracket Ceramic elongation Displacement Sensor Pressure Port Inconel tube Pressure Sensor Manifold Conventional Sensors Insulation Strain Gauge Srain gauge Bracket III. IFE SYSTEM ARCHITECTURE IXV Data Aquisition Unit In Fig. 12 a schematic of the IFE architecture is depicted showing that the acquisition of all the sensors is centralized at IXV avionics level as it will be described in the following. The sensors of IXV are divided into two categories according to the level of complexity: o Conventional sensors including thermocouples, pressure sensors, strain gauges and displacement sensors which deliver raw data directly to the data acquisition unit which in turn transmits it via telemetry to ground stations. o Advanced sensors consisting of an infrared camera which require processing of signals in a dedicated data handling unit prior to telemetry. This data handling unit communicates with the vehicle unit via an RS 422 interface The architecture also distinguishes three types of sensor components based on their location: those components such as thermocouples and pressure ports placed on the thermal protection surfaces, bracket located sensors such as strain gauges and sensors placed on the vehicle interior (cold structure). In Fig. 13 the schematic of the IXV Experiment Data Acquisition System is reported. The system has been designed and procured by Alenia Aermacchi and ACRA. Basically all the 300 sensors of IXV plus the IRC data coming from the DHU are gathered by four Data Acquisition Units (DAUs). Data coming from DAUs are passed through a dedicated Ethernet switch to a DAU Master that has a twofold objective: to send the Cold Structure TPS RS 422 IR Camera DHU IR Camera Fibre Optic Cable FOA + Periscope Fig. 12. IXV IFE Architecture. Fig. 13. EDAR Schematic Advanced Sensors (Infrared Camera Experiment) Insulation IAC-14-D2.6.3 Page 6 of 14

IV. SENSORS IMPLEMENTATION, QUALIFICATION AND INTEGRATION The realization of all the experiments described in the previous sections is made possible through the implementation of a large number of sensors. On board of IXV are currently integrated: 105 Type S platinum thermocouples, 89 Type K thermocouples, 37 absolute pressure sensors, 2 differential pressure sensors, 12 displacement sensors, 48 strain gauges and advanced Infra-Red Camera system. In the following a description of the adopted measurement systems will be provided. For all the items of the in-flight experimentation subsystem the compatibility with re-entry environment has been verified and dedicated qualification test campaigns have been performed. The integration of the 300 sensors in the IXV structure including the routing of all the harness from the measurement points up to the data acquisition system has been one of the most challenging activities of the entire IXV project. IV.1. Temperature Sensing System Different solutions for temperature measurement on IXV have been adopted depending on the location on the vehicle (i.e.: ceramic nose and windward, cold structure, ablative TPS) and on the measurement range (i.e. type S for high temperature, type K for medium/low temperature). For what concern the Type S thermocouples for the ceramic TPS, the highest challenge in IXV is the long duration of the re-entry phase. Prolonged exposure of ceramics to high temperatures allows solid state reactions between the C-SiC ceramic and the platinum sheath of the thermocouples. It also forces use of refractory metals and ceramic insulations for the pressure pipes and displacement sensors. The damage to the thermocouples is avoided by use of a non-oxide ceramic coating sputtered on the platinum sheath. The coated thermocouples are flexible enough to be used in the flap area in which they are subject to repeated deflection spanning a 40 angle between end positions. For the length compensation of the thermocouples placed on the flap a comb system made out of Inconel is used inside the outer flap support structure (see Fig. 14). The thermocouples are clamped before and after the comb system to avoid tension on the bonded sensor while routing it on the flap. Fig. 14. Moving Flap comb system. Each sensor is guided between two Inconel plates to avoid buckling. The comb system is fixed with a clamping system on the shear pin and the whole unit is covered with Nextel insulation to close the flap support and avoid any plasma infiltration. The comb system with the sensors has been placed in the support before integration into the vehicle. Integration and routing of the sensors on the body flaps themselves has been done directly on the vehicle, as the body flaps are installed at last. A lifetime bending test has been performed with a type S thermocouple after heating. The sensor showed no degradation during the test. In addition Type S thermocouples have been integrated in the Body Flap Qualification model and subject to the mechanical test campaign of this latter: vibration, shock and flap chain movement testing. In Fig. 15 the routing and the fixation of the Type S thermocouple on the IXV ceramic flap, manufactured by MT-Aerospace is shown. Thermocouples head is kept in place through a groove made in the CMC that is then filled with ceramic glue. Fig. 15. Type S thermocouple integrated on body flap leeward surface. IAC-14-D2.6.3 Page 7 of 14

The thermocouples are routed along the ceramic TPS surface and the difference in their thermal expansion with respect to that of the ceramic substrate is absorbed by allowing for certain waviness as they transverse the surface. In Fig. 16 the internal part of a HERAKLES ceramic shingle is shown with several thermocouples routed on it. In particular the shingle depicted in Fig. 16 is the one hosting the CATE experiment. Differently from the body flap, for the fixation of the thermocouples on the shingle a coinfiltrated wedge concept was used in addition with zirconia based glue. Capability of the thermocouple fixation to withstand the launch and re-entry environment has been verified during the qualification test campaign of nose and shingles hardware: vibration, shock, thermal tests. The original design envisioned introducing pairs of sheathed thermocouples into the ablative so that the outermost thermocouple would be at the surface and a second at the interface to the vehicle structure. Stagnation tests showed that the heat transport through the thermocouple sheaths is significant and falsifies the measurement. The design was modified to include three exposed measuring elements with the first one embedded just underneath the surface and a second one in the mid plane of the ablative. The third one remained at the interface between ablative and cold structure. The resulting sensor can be seen in. Fig. 16. Type S thermocouple fixed and routed on HERAKLES CMC shingle. The leeward, lateral and base of the IXV vehicle are covered by ablative material. Measurement of ablative behaviour requires a dedicated design of the thermocouple measuring system adapted to the behaviour of the ablative material. Fig. 18. Ablative Thermocouple plug. The routing of the harness of 194 thermocouples inside the IXV vehicle has been a quite challenging task; in Fig. 19 one of the bracket mounted on the cold structure and used to collect thermocouples cables interfaced with harness compensation cable trough LEMO connector is shown. Fig. 17. Ablative Thermocouple Concept. Fig. 19. Thermocouple Harness bracket installed on IXV Cold Structure. IAC-14-D2.6.3 Page 8 of 14

IV.2. Pressure Sensing System IXV Pressure sensing system is constituted by three main elements: pressure transducers, pneumatic line, pressure ports (both in ceramic TPS and in ablative/cold structure). Customized Kulite Pressure transducers have been selected for IXV application. They represent an optimal solution balancing good accuracy with suitability to harsh environment. The sensitivity of the pressure system to input voltage variations and sensor temperature are a concern as the measurement occurs at very low pressures (below 10 hpa). These error sources were quantified and reduced through appropriate sensor selection and a dedicated calibration procedure. Pressure transducer are grouped in manifold bracket attached to the IXV cold structure (see Fig. 20). The temperature of the pressure manifolds is monitored to allow post flight compensation. Capability of pressure transducer installed on the manifold to withstand the mission environment has been verified through a dedicated qualification test campaign. In Fig. 21 the test jig used for the mechanical qualification of various IXV sensors is shown. On the right of the picture the pressure manifold is visible with a pressure transducer installed plus additional dummies. In addition to the mechanical testing, pressure sensors have been also submitted to thermal vacuum testing. The measurement of pressure on the nose and windward surfaces requires a port located on the outer ceramic skin of the vehicle connected via piping to the pressure sensors inside the vehicle. The pressure port was developed in cooperation with the TPS manufacturer allowing co-infiltration during the manufacturing of the C-SiC nose cap and tiles. Varying the inclination of the entry holes in adjacent ports allows measuring skin friction effects. These ports are direct links between the hot plasma and the vehicle interior and are therefore subject to very large thermal gradients and thermo-mechanical loads. Failure or leakage of the ports would be catastrophic so they have been designed to withstand re-entry heat loads without stressing the surrounding C-SiC ceramic. The design uses a stabilized zirconia ceramic bush as insulator to reduce heat transfer to the vehicle interior. This bush is decoupled from the port by means of compressed graphite washers which also dampen movement of the pipes and any transmission of vibroacoustic loads during launch (see Fig. 22). Fig. 20. Pressure transducer grouped in manifold bracket attached to IXV cold structure. Inconel Pipe Zirconia Bush Graphite washers C-SiC Port Fig. 21. Test Jig for mechanical qualification of IXV sensors. Fig. 22. Pressure Port Design. IAC-14-D2.6.3 Page 9 of 14

The ceramic port interfaces to a 4 mm diameter Inconel pipe which transverses the ceramic fiber insulation and is then fixed to the vehicle structure. The piping changes to Teflon and is brought into the inside of the vehicle were it interfaces with the pressure manifold in which the sensors are fixed. The bending of the pipe (Fig. 22 and Fig. 22) absorbs the differential thermal expansion between its vehicle fixation point at 140 C and the interface to the ceramic bush at 1000 C. A primary concern during design of the pressure sensing system is the ability to capture slight pressure changes in real time. The hydraulic delay of the piping and the entry effects were minimized through design of two alternative bush geometries and measurement of the delays after implementing the proposed geometries in a laboratory bench. It has been demonstrated that the delays are much lower than the sampling rate of 100 ms (10 Hz) and are therefore considered acceptable. Flight representative pressure ports have been included in both nose and shingle qualification model and therefore subjected to a full mechanical qualification test. In addition, in order to verify the capability of the pressure ports to withstand the strong thermal gradient generated during the re-entry, dedicated hot test have been performed both by using an oven capable of reaching 1100degC at the pressure port surface and a plasma torch by which it was possible to reach temperature higher than 1600 degc. In Fig. 23 pressure ports with Inconel and Teflon tubing during the integration phase are shown both for the nose (on the left) and for the shingle (on the right). As for the thermocouples also for the pressure ports the integration was a quite challenging task considering the presence of the Inconel tubing that had to pass through the various insulation layers. Fig. 24. Ablative Pressure Port Design In Fig. 25 it is visible the zirconia tube of the pressure port sticking out from the IXV cold structure surface before the installation of the ablative tiles. Fig. 25. Pressure port ceramic tube sticking out from IXV cold structure surface before ablative tile integration. IV.3. Displacement Sensors Fig. 23. (Left) Integration of pressure piping in ceramic nose. (Right) Pressure port in ceramic shingle with Inconel piping. The pressure ports on the ablative have a zirconia tip that crosses the ablative and interfaces to a metal insert. The other end of the insert holds a coupling interface to a Teflon pipe (see Fig. 24). Displacement sensors are used to measure steps evolution between the adjacent TPS units in flight. Of particular interest is the step between the nose and the windward TPS due to the high temperature evolution in this region. The displacement sensors were designed to be placed on the outer surface of the vehicle and are capable of penetrating the plasma seals (see Fig. 26). The displacement sensor integrates an LVDT (Linear Variable Differential Transformer) sensor in an aluminium housing. The stroking element is a ceramic point with poor thermal conductivity and very low thermal expansion. The sensor can detect movements of 0.2 mm over a stroking distance of ± 5 mm. IAC-14-D2.6.3 Page 10 of 14

Fig. 28. Displacement sensors integration phase. Fig. 26. Displacement sensors concept. Nose-Shingle paired measurement. Testing in the early phases of the programme indicated that the surface of the C-SiC was too rough to allow an error free measurement and the resulting high coefficient of friction was a concern as the tiles move with respect to the vehicle structure. It was decided to smoothen the contact surface using an adhesive layer placed on the tile (see Fig. 27). Displacement sensors have been fully qualified with respect to mechanical (see Fig. 21) and thermal environment. IV.4. Strain Gauges A second type of mechanical sensing element is used to measure the forces on the metallic stand-offs of the nose and the windward TPS. To this end high temperature strain gauges are placed on the stand-offs. 10 mm long temperature compensated encapsulated gauges have been selected for the strain measurement with an operational limit of 750 C and an Inconel sheath which allows welding to the brackets. The gauge integrates a half bridge sensor with a 120 Ω resistance and a strain limit of ± 10000 μm/m and a dummy element for thermal compensation. The use of four strain gauges per bracket (Errore. L'origine riferimento non è stata trovata.) allows a clear distinction between bending, torsion or purely unidirectional loading. Although the curved areas exhibit the higher strains the sensors must be placed in flat regions. Fig. 27. Displacement sensor during testing. Integration of displacement sensors has been made quite complex by the fact that the ceramic rod has to pass through the shingle insulation layers. First the displacement sensor assembly have been mounted on the cold structure insert and afterwards the shingles with the insulation stack have been installed (see Fig. 28). During this operation any excessive movement of the shingle would have caused the breaking of the displacement sensors. Fig. 29. Loaded Areas of Stand-Offs and Measurement Points. IAC-14-D2.6.3 Page 11 of 14

In order to allow modular integration the strain gauges are cabled to coupling connectors and the cable integrates a Wheatstone bridge which allows temperature compensation of the measured strain. Fig. 30. Instrumented nose metallic stand off during straing gauges calibration campaign. The temperature monitoring of these stand-offs is done by spot welded Type K thermocouples. The instrumented stand-offs and brackets have been tested under load and temperature to establish their performance baseline. Furthermore a dedicated calibration campaign on the flight stand-off equipped with strain gauges has been performed (see Fig. 30) The strain gauges will not be able to record launch strains but will record all re-entry loads including the initial loading during vehicle splashdown. In Fig. 31 the fully integrated nose (ceramic shield plus metallic cold structure) immediately before the installation on the vehicle is shown; the bundles of sensors cable (thermocouples, pressure sensors, strain gauges) to be routed inside the IXV structure are visible in the center of the picture. In Fig. 32 a close-up of the metallic stand-off with the strain gauges and thermocouples is shown. IV.5. Infra-Red Camera Assembly The goal of the IR camera experiment is the continuous (25 Hz) high resolution backside temperature mapping of one body flap for the whole deflection angle range during the re-entry phase. The near infrared camera uses periscope optics (a sapphire mirror with narrow band coating) to observe the flap area. In order to minimize thermal load on the infrared camera a fiber optic waveguide is used to allow placement inside the vehicle (see Fig. 33). Fig. 31. Ceramic nose ready for integration on IXV with bundles of sensors routed (Thermocouples, Pressure tubes, Strain Gauges, Bridge adapter box) Fig. 33.Infra Red camera hot periscope on vehicle base surface with flap field of view. Fig. 32. Nose Instrumented metallic stand off integrated on IXV. The infrared system integrates a three wavelength filter wheel (see Fig. 34 Right) allowing independent IAC-14-D2.6.3 Page 12 of 14

measure of temperature and emissivity as discussed in [2]. This allows correction for the angular position of the flaps as well as hydrazine and ablative soot contamination. The derived thermal maps are then used to determine the windward temperatures using inverse methods. Fig. 34. (Left) IR camera periscope with sapphire mirror and Inconel housing. (Right) IR camera with filter wheel system The infra-red camera communicates to a central data handling unit, an electronic box integrating power regulation and distribution, analog to digital signal conversion, a central processing unit, redundant solid state memories and an input /output board in a four card cage frame (see Fig. 35). Shock damping elements and water tight seals ensure its survival during landing and vehicle retrieval operations. The unit communicates to an outer connector allowing access to data after vehicle integration. profile. In order to derive an estimate for the appropriate exposure times an algorithm is employed which analyses the grey level content of the recorded images. Based on this analysis the data handling unit changes the settings for the next image. As the camera is faster than the algorithm the change is implemented after the second image has been written to an intermediate buffer. Gain and exposure speed adaptations allow measurement of the entire temperature range from 300 C to 2000 C with a 10 degree error. Infra-Red camera images are both recorded into the solid state recorded of the DHU and transmitted on ground to telemetry. In order to cope with telemetry budget several data reduction techniques are combined in order to reach the target bandwidth reduction. The Infra-Red Experiment assembly has passed a full series of qualification tests. All the components (i.e. DHU, IR camera, filter wheel, front optic assembly, hot mirror) have been subjected to mechanical and thermal testing according to the IXV environment. In addition for the front optic assembly, and in particular for the sapphire mirror that will be directly exposed to the reentry environment, an hot test has been successfully performed. In Fig. 36 an infra-red picture of the IXV body flap taken during one of the numerous functional testing performed is shown. With the aim at deriving the actual flap deflection in post flight analysis, still picture of the flap in different angular position, spanning the whole deflection range, have been taken. An algorithm has been developed which can determine the flap orientation angle with a root mean error on the order of 0.1 degrees by using multiple image points in the target scene (notably the flap edges) and deriving measures (e.g. edge orientation, corner angles, visible flap area) which are directly and uniquely linked to the flap orientation angle. The use of a large number of feature points enables precise measurements despite the comparatively small resolution of the camera (320 x 256 pixel sensor). Fig. 35. IXV IR Camera Data Handling Unit. A change in the exposure time of the infrared detector is necessary to record the images at the various temperatures encountered throughout the mission Fig. 36. IXV Body Flap view from IR camera. IAC-14-D2.6.3 Page 13 of 14

V. CONCLUSION The definition of the IXV In Flight Experimentation plan has been described as well as the implementation of the various experiment, the qualification of the adopted solution for the IXV environment and the integration on the vehicle. More than 300 sensors have been successfully integrated and tested in the IXV vehicle. All the adopted solutions for the realization of the various experiment have successfully passed a comprehensive qualification campaign. As it was expected the integration of the sensors (e.g. thermocouples, pressure sensors, displacement sensors) in the ceramic TPS has been one of the most critical and time consuming activity of the entire project. The Infra-Red camera experiment with the realization of the outboard periscope with a sapphire mirror capable of withstanding temperature higher than 1000 degc will provide a unprecedented opportunity to monitor the thermal behaviour of a critical ceramic TPS component like the IXV flap all over the re-entry mission. The huge quantity of data that will be gathered during the IXV mission will require an extensive Post Flight Analysis (PFA). The IXV PFA will be articulated on three different level. Level-0 and Level-1 PFA will be performed in the frame of the IXV industrial consortium and will be mainly devoted to the verification and reconstruction of the acquired data, the identification and resolution of anomalies, the reconstruction of the flown trajectory and the preliminary identification of the system performance. Afterward the Level-2 PFA will be performed also with the involvement of additional entities (e.g. research institutions, national agencies representative, ESA- ESTEC researchers) of contributor member states and it will be aimed at the scientific investigation of the various experiment and at the identification of improvement in the prediction tools and technologies to be utilized for the future development of space transportation system. VI. ACKWNOLEDGMENTS The authors wish to thanks all the members of the IXV industrial team that together with the European Space Agency, the prime contractor Thales Alenia Space (Italy) and RUAG SPACE (Switzerland) worked for the successful implementation of the In Flight Experiment subsystem: Lambda-X and the institute of Fluid Dynamic at the ETH in Zurich for their respective contribution in the development of the infrared camera experiment; the Von Karman Institute in Belgium for the design and implementation of the CATE experiment; the TPS manufacturers HERAKLES, MT- AEROSPACE and AVIO for their support to the development of solution for sensors integration; DASSAULT Aviation for the scientific contribution to the design of the experiment with the support of ONERA and CIRA. REFERENCES [1] G. Tumino et. al. The Ixv Programme: Ready for Flight, 65 th International Astronautical Congress, Toronto, Canada (2014). [2] C. Pereira et. al. Thermal Measurement Techniques for (Movable) Control Surfaces in Re-entry Vehicles, 6 th European Workshop on Thermal Protection Systems and Hot Structures, Stuttgart (2009) [3] C. Pereira, et al. In Flight Experimentation for the IXV re-entry vehicle: objectives, experiment design and implementation 62 th International Astronautical Congress, Naples, Italy (2012). IAC-14-D2.6.3 Page 14 of 14