IRS Ground-Testing Facilities: Thermal Protection System Development, Code Validation and Flight Experiment Development

IRS Ground-Testing Facilities: Thermal Protection System Development, Code Validation and Flight Experiment Development
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   1 24 th  Aerodynamic Measurement Technology and Ground Testing Conference, AIAA-2004-2596, Portland, June 2004 IRS Ground-Testing Facilities: Thermal Protection System Development, Code Validation and Flight Experiment Development G. Herdrich 1 , S. Löhle 2 , M. Auweter-Kurtz 3 , P. Endlich 2 , M. Fertig 1 , S. Pidan 2 , E. Schreiber 2   Institut für Raumfahrtsysteme (IRS), University of Stuttgart Pfaffenwaldring 31, D-70550 Stuttgart, GERMANY Tel. +49 711-685-2412, Fax. -7527 E-mail:  ABSTRACT Five plasma wind tunnels (PWK 1 - 5) are in operation for the investigation and qualification of TPS materials. Different plasma sources are developed for generating the high enthalpy plasma flows as they are expected for re- entry flights. Exemplary results of measurements for trajectory points of X-38 and the planned EXPERT vehicle are presented. The PWKs and auxiliary facilities such as the IRS black body source and the emissivity measurement facility as well as their different application possibilities are described. Results are shown exemplary such as in-flight measurement technique calibration, emissivity measure- ments of SiC, investigation of passive-active transition (German research program ASTRA) and heat flux profile simulation for X-38 (TETRA). Within the PYREX-KAT38 development, a miniaturized pyrometric system which was planned to measure the temperatures in the nose structure of the X-38 during re- entry, the stagnation point heat flux profile of X-38 was simulated by programming the computer controlled probe platform in the magnetoplasmadynamically driven PWK1. Efforts are being made at the inductively heated PWK 3, where the entry into Venus and Mars` atmospheres, also taking into account the Martian dust, are simulated. The accuracy of the simulation of entry conditions strongly depends on the ability to determine the flow conditions. Both, intrusive probe measurement techniques including mass spectrometry and non-intrusive, optical techniques such as laser induced fluorescence measurements (LIF) are used to investigate the flows. Surface temperatures are evaluated two-dimensionally using a calibrated charge injection device (CID). All measurements and the facility development are accompanied by numerical simulations. Comparisons of numerical and PWK simulations lead to a better understanding of the experimental plasma flows and together with newly developed numerical simulations of the gas surface interactions result in better development process possibilities. INTRODUCTION The potential of development for space capsules whose mission include a return to Earth or an entry in the atmosphere of neighbouring planets is mainly restricted by thermal and aerodynamic loads that occur during entry manoeuvre into the atmosphere. Concerning Earth, the space vehicle enters the atmosphere at an altitude of about 120 km with a velocity of 8 km/s is braked down by atmospheric friction only. The high kinetic energy is dissipated by a strong bow shock in front of the essentially blunt body. Hence, a partially dissociated and ionized hot plasma attains the vehicle's surface which has to be protected. Actually, thermal protection systems (TPS) that use the principle of radiation cooling to shield the capsule are favored because of the possibility to be reused [1], [2]. The materials currently under investigation are ceramics based on Silicon carbide (SiC) whose oxidation behavior restricts the maximum heat load [3], [4]. However, surface temperatures above 2000 K are expected [5]. Depending on the expected oxygen partial pressure and surface temperature a mission critical border for the re-entry trajectory can be estimated. The flight path is, therefore, designed according to this boarder. Thermal and aerodynamic heat loads are computed using computational fluid dynamic methods and the material's chemical and thermal boundary layer is simulated using plasma wind tunnels (PWK) [6], [7]. The on-stream condition is usually characterized by the local enthalpy and total pressure at stagnation point. Fig 1. Flight envelopes of X-38 and EXPERT together with operational envelopes of IRS-PWK Fig 1 shows calculated entry trajectories for X-38 and EXPERT together with the capabilities at IRS to rebuild such a trajectory experimentally. By simulating a 1 Research engineer, Member AIAA, IRS 2 Research engineer, IRS.  3 Professor, Associate Fellow AIAA, IRS Copyright   2004 by the authors. Published by the American Institute of Aeronautics and Astronautics, Inc. with permission.   2 hypersonic plasma flow around a material sample of the IRS-PWK dimensions a comparison of the computed with the experimental boundary layer becomes possible. For the understanding of the chemical behavior and further development of numerical codes the gas species in the boundary layer have to be quantified which becomes accessible through the usage of optical diagnostics. Furthermore, the plasma state in front of the probe has to be compared because its srcin is different. In numerical simulations the plasma state is a result of the strong bow shock. In PWK the plasma is generated using an appropriate generator and the probe is positioned in the streaming plasma [8]. There are three different generator concepts [9], [10]. The magnetoplasmadynamic plasma generator RD5 is used to simulate the re-entry path in higher altitudes where the maximum heat load is expected. With the thermal arc-jet generator RB3 a pressure and enthalpy regime expected at lower altitudes of a re-entry flight can be simulated. The third concept is an inductively heating plasma source whose advantages are the possibility to perform tests in chemically reactive flows, e.g. pure oxygen or carbon dioxide [11]. Pure oxygen is mainly used for the qualification of sensors for flight experiments, but also for basic investigations of thermal protection systems. The generators are also used for comparisons of numerical and experimental rebuilding of the boundary layer in front of material samples or copper heat flux sensors, whereas the latter is often used for reference heat flux measurements in PWK facilities. Experimentally generated boundary layers can be investigated using different measurement principles. Integral values as for example pressure, temperature and enthalpy are measured with appropriate gages or extracted from probe measurements. Locally detached values are measured mainly using optical diagnostic methods. The local surface temperature is measured using a calibrated charge injection device (CID). At IRS, temperatures and qualitative densities of major gas species are investigated using the laser-induced fluorescence measurement technique. Measurements of atomic oxygen and nitrogen as well as qualitative measurements of nitric oxide and silicon monoxide were successfully performed. Additional intrusive measurement techniques are approved at IRS to characterize the plasma flow, for example electrostatic probes, enthalpy probes or wedge type probes to extract Mach number [12]-[14]. All these experimental tools together with numerical simulations provide excellent opportunities to develop and qualify radiation-cooled materials for reuseable spacecraft and ablative material systems to be used for capsules and interplanetary probes. Additionally, they are much less expensive than experimental space flights. Atmospheric entry mission phases encounter problems, such as hyper-sonic aerothermodynamics and specific gas surface interactions due to the actually used ceramic heat shield materials. Goals include managing the guidance navi-gation, control, landing technology and inflatable technologies such as ballutes that aim to keep vehicles in the atmosphere without landing [15]. The requirement to save mass and energy for interplanetary missions such as the Mars Society Archimedes Balloon Mission, Mars Sample Return Mission, Mars Express or Venus Sample Return mission led to the need for new manoeuvres like aerocapture, aero-breaking and hyperbolic entries e.g. for sample return missions [16-18]. Concerning Mars missions the third well established generator principle at IRS is used. Its inductively heating principle is capable to generate CO 2  plasmas in a wide enthalpy and pressure regime. Actually, in this field no successful flight has been performed in Europe yet, but the importance of these manoeuvres and the need to increase the knowledge of the required TPS designs and its behavior during such mission phases point out the need of ground testing facili-ties, numerical codes and flight experiments. As a result of the experience within the plasma diagnostic tool develop-ment and the PWK data base acquired during the last 20 years, flight experiments like PYREX (PYrometric Entry Experiment, capsules EXPRESS, MIRKA) and HEATIN (HEATshield INstrumentation) on MIRKA have been de-veloped at IRS, qualified and successfully flown [4, 19, 20]. Flight experiments such as RESPECT (RE-entry SPECTrometer) [21] and PYREX for EXPERT are in the phase B. The rear-side temperature measurement system PYREX-KAT38 (PYrometric Entry EXperiment), measuring the temperature distribution in the X-38 nose structure, was contributed by the Space Transportation Division of IRS [22], [23]. The information provided by this system is pertinent to several fields of interest, i.e. the temperature histories at five positons in the nose structure of X-38 during entry, statements on the behavior of the TPS material and the heat flux distribution. 1. DIAGNOSTIC TOOLS Black Body (BB) Source The Minipyrometer used for measurements of rear side temperatures of PWK material samples [14] and the PYREX-systems are calibrated with the IRS BB source working between 600°C and 2600°C. The source shown in Fig. 2 consists of an electrically heated graphite radiator with a cylindrical cavity. Fig 2. Diagram of the IRS black body To approximate the BB, the cavity has a large length-to-diameter ratio and the cavity walls should maintain an iso-thermal profile. This is achieved by a variable heat source distribution that is realized by varying the cavity wall thick-ness. The outline of the graphite rod was therefore calcu-lated using a numerical procedure [24]. To provide high accuracy, different geometries for different temperature ranges are applied achieving a maximum temperature departure ∆ T<20 K in the cavity. The second radiation outlet is used for the calibration. Temperature is varied using the control pyrometer; tem-perature and photo current are measured simultaneously.   3 1E-121E-111E-101E-091E-081E-071E-061E-05 700100013001600190022002500 T [°C]    I      [   A   ]   Fig 3. PYREX-KAT38 FM photo currents versus black body temperatures [23] Fig 3 shows the result of the PYREX-KAT38 FM cali-bration as an example. A photo current line versus the measured temperature is obtained. Here, the calibration curve is used to obtain temperature data from the photo current values achieved during flight. In a similar manner, the Charge Injection Device (CID) camera is calibrated. Its purpose is to measure two-dimensionally the temperature on the surface of a sample. This temperature distribution becomes of special interest when not only standard sample geometries are to be investigated but also structural parts, e.g. steps and gaps. Exemplary, Fig 4 shows the temperature distribution of a SSiC sample. The CID camera is calibrated using the BB source to obtain a temperature for each measured intensity value. Fig 4. Temperature distribution on the surface of a SSiC sample measured with CID camera Emissivity Measurement Facility In its working principle the device for emissivity measure-ment is similar to the IRS BB Source. However, it is more complex as a movable system is required. It can be divided into two parts. The first part is the emissivity measurement apparatus, composed of the devices that are outside the vacuum tank, i.e. vacuum pumps, pneu-matic cylinder etc. The second part is the movable BB assembly inside the vacuum tank. The apparatus including all sub-systems necessary for operation and measurement is shown in Fig 5. Fig 5. Scheme of the IRS Emissivity Facility 1. Vacuum tank. 7. Argon supply 2. BB assembly 8. Mechanical vacuum system 3. BB positioning system 9. Oil diffusion vacuum system 4. Front plate + optical glass 10. Power supply 5. ZrO 2  flange 11. Pyrometer 6. Pneumatic cylinder 12. Measurement control unit  For temperatures higher than 2200 K the increasing evaporation rate of graphite requires the use of an inert gas atmosphere. On the back side of the tank, a special flange has been attached. This flange provides the access for the ZrO 2  stick that moves the material sample inside the cavity radiator. Moreover, the flange is used to fix the assembly by inserting a Steatit isolator in it. On the front side of the tank (the side on the right in Fig 5), a steel plate with the optical glass functioning as optical access for the pyrometer is attached. Emissivity measurements consist of the following steps: 1. The sample material (26,5 mm in diameter) is placed in a sample holder and positioned at the end of the cavity radiator. The apparent emissivity of its surface is 1. 2. The cavity-radiator is electrically heated up to the temperature at which the emissivity has to be measured. The electrical current can be adjusted using a control unit, hence several temperatures can be obtained increasing or decreasing the current level. Due to heat radiation exchange, the probe and the cavity-radiator have the same temperature. 3. During the heating process, the temperature in the sample is measured by the pyrometer. 4. When the temperature is reached, the sample is moved into the vacuum chamber (pneumatic cylinder). 5. When the probe is in its end position, another measurement is taken, hence a second value of I (photo current) is obtained. The process of data evaluation depends on the pyrometer used for the measurment. With the linear pyrometer LP 20 (IKE, Univ. of Stuttgart) the emissivity can be obtained by the ratio of the photo currents measured with the device at the two positions: (1)   122  I  I  = ε  .   4 Other commercial pyrometers are either not linear or they do not provide a photo current information or both. Here, the information can be obtained by the consideration of a temperature change due to emissivity change [25]. For sintered SiC material samples spectral emissivities of 0,87 at 1000 °C were measured at 900 nm. 2. IRS PLASMA WIND TUNNELS The term plasma wind tunnel is to be understood as a steady state test facility in which a high enthalpy flow is produced with the help of a plasma generator. The purpose of PWK is to rebuild the thermal and chemical loads on a vehicles heat shield, rather than rebuilding the aerodynamics. For the air system the simulation regimes of the PWKs at IRS are shown in Fig 1. Here, pressure ranges and local specific enthalpies in the plasma jets are related with cor-responding velocities and altitudes of the space vehicle.   As examples for the facilities at IRS the scheme of the PWK 1 facility can be seen in Fig 6. Fig 7 shows a photograph of PWK 2. The vacuum tank used for PWK 1 and 2 is a 6 m-long steel tank with a diameter of 2 m which have a double-wall cooling. Fig 6. Scheme of plasma wind tunnel PWK 1 The tank is closed with a hemispherical part which is connected to the vacuum system and protected against heat by water-cooled copper shields. On the other side the vacuum chamber can be opened by moving the plane cover plate on a guide rail. The plasma source is not located in the vacuum tank, but flanged on a conical part of the plate. The cone-shaped element of the end plate enables the plasma source to be fixed at that point. The whole plasma jet range is accessible by optical methods. The tank is equipped with a 4-axis positioning system on which the different probes and the specimen support system can be mounted. This allows the simulation of parts of the re-entry trajectories by moving the specimen in different plasma flow regimes. A certain point on a re-entry trajectory is defined in the IRS-PWK by the ambient pressure, the axial distance to the generator, the arc current and the mass flow. Together, these values result in the proper enthalpy, total pressure and temperature or heat flux. Fig 7. Plasma wind tunnel PWK 2 Optical glass windows allow pyrometric temperature measurements on the front side of the specimen at distances from the plasma source between about 50 mm and 1 m. Moreover, optical measurements perpendicular to the plasma jet axis are possible through three movable flanges of three optical glasses each, which are located on both sides of the tank opposite each other and on the top (Fig 7). For the arc-jet driven PWK, the following naming is made. As long as the described vacuum tanks are equipped with MPGs, these plasma wind tunnels are called PWK 1 or PWK 2. If, instead, a TPG is used, they are called PWK 4. The whole experimental setup of PWK 3 (Fig 8), consists of the IPG and the vacuum chamber. The size of this vacuum chamber is about 2 m in length and 1.6 m in dia-meter. Material support systems and mechanical probes can be installed onto a moveable platform (2-axis) inside the tank. The plane lid of PWK 3 carries the plasma generator and the external resonant circuit, which contains the capacitors with the connection to the IPG. The measurement equipment of the arc-jet driven facilities can as well be used for PWK 3 Fig 8. Scheme of plasma wind tunnel PWK 3 It is well known that CO 2  requires a deactivation using a suitable gas such as N 2  because the produced CO may create an explosive mixture that especially could endanger the areas of higher pressure of the facility i.e. parts of the vacuum pump system. Therefore, all experiments for the simulation of CO 2 -atmospheres are carried out using an additional injection of N 2  at the right side flange of the chamber (Fig 8). Common studies [26-28] point out that the ignition limits for CO are 12.5 (lower limit of con-centration) and 74 (upper limit). These data are related as volumetric relative shares. Furthermore, the deactivating of CO requires a volumetric share of O 2  lower than 5 %. If the maximum production of CO (total dissociation of CO 2 ) is assumed, the required deactivating gas N 2  can be calculated. A calculation assuming total dissociation of CO 2  leads to the requirement that 5.4 g/s N 2  (safety) are   5 needed to deactivate plasma derived from 1g/s CO 2 . An advanced mass flow meter system is installed to control the three required mass flow rates i.e. the CO 2  mass flow rate for the IPG, the N 2  mass flow rate for the IPG (taking into account the 3 % volumetric share of N 2  in the Martian atmosphere) and the N 2  safety mass flow rate injected at the end of the chamber. A powder feeder is available for the simulation of dust particles during entry, e.g. the dust in the Martian atmosphere [29]. The plasma wind tunnels PWK 1 – 4 are connected to the central vacuum, central power and central gas supply sys-tem. The plasma wind tunnel PWK 5, also equipped with a thermal plasma generator, is not connected to a vacuum system. Thus, this test facility is suitable for testing TPS materials at pressure levels higher in the range 10 5  Pa as it is for example expected for the EXPERT vehicle (Fig 1). Fig 9. Plasma wind tunnel PWK 3 Power Supplies The electric power for the arc plasma generators is supplied by a current-regulated thyristor rectifier consisting of six identical units supplying 1 MW each. These may be connected in series or parallel, thus varying the desired output level of current, voltage, and power. The current ripple is less than 0.5%. The maximum current is 48 kA supplied at 125 V and the maximum voltage is 6000 V at a current of 1000 A. For the inductively heated plasma wind tunnel PWK 3 a radio frequency generator with a primary power of 375 kW is used. This device allows the operation of an induction coupled plasma generator with a plate power of 180 kW at nominal frequencies between 0,5 and 1,5 MHz. An external resonance circuit is designed for an optimal coupling into the plasma over a wide pressure range with different gases (Fig 9). The resonant circuit is built in Meissner type switching using a metal-ceramic triode with an oscillator efficiency of about 75% [11]. Its nominal frequency can be changed by switching the order or number of capacitors as well as by the use of coils with different inductivities. A maximum of seven capacitors can be connected. The external resonant circuit is cooled by a water cooling circuit. Vacuum System A vacuum pump system is used to simulate pressures at altitudes up to 90 km (Earth). This pumping system con-sists of four stages: the first two stages consist of roots blowers, the third stage is a multiple slide valve type pump, and the last stage is a rotary vane type pump. The total suction power of the pumps amounts to 6 000 m 3  /h at atmospheric pressure and reaches about 250 000 m 3  /h at 10 Pa measured at the intake pipe of the system, which has a diameter of 1 m. The base pressure of the system is 0.5 Pa. The desired tank pressure can be adjusted bet-ween the best achievable vacuum and 100 kPa by re-moving one or more pumps from the circuit and/or mixing additional air into the system close to the pumps. 3. PLASMA GENERATORS Two plasma generator concepts, the thermal and the magnetoplasmadynamic generator (TPG and MPG), are in use in the arc heated plasma wind tunnels; they mainly differ in the acceleration concept [10]. In TPGs the test gas is heated by means of an electric arc and accelerated through a nozzle. For MPGs, additional electromagnetic forces are used to accelerate the plasma. The inductively heated plasma generators (IPG) also belong to the group of TPGs. With the electrodeless design the plasma is produced by inductive heating using a triode-driven radio frequency power supply. Arc-driven Thermal Plasma Generator (TPG) In order to simulate high enthalpy air flows at pressure levels above 5 kPa, which is the limit of MPGs and in heat flux ranges between 0,1 MW/m 2  and about 3 MW/m 2  in stagnation point configuration, a coaxial thermal plasma generator (TPG) called RB3 (Fig. 13) has been developed for PWK 4. Fig 10. Plasma Source RB3 The test gas is heated in the discharge chamber by an electric arc and accelerated in a nozzle. A 2% thoriated tungsten cathode is used. The anode is a water-cooled copper cylinder, whereas the nozzle is electrically insulated. Since contact between the oxygenic part of the test gas and the cathode has to be avoided, the air used for re-entry simulation is divided into two parts. As main part the nitrogen is passed along the cathode into the plenum chamber. The oxygen is injected at the downstream end of the anode towards the nozzle throat. For other atmospheric entries, as for example Mars, the RB3 plasma source can also be driven with CO 2  as long as the reactive part of the mass flow is injected downstream from the cathode [30]. To ensure a good mixing of the nitrogen and the oxygen, the injection point is positioned in the subsonic part of the TPG such that a backflow of oxygen into the cathode region cannot be ruled out completely. However, tests have shown that the cathode erosion rate due to the bi-throat design for RB3 is as low as observed in the MPGs operated in PWK 1 and
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