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The inductively-heated plasma generator IPG6-B at Baylor University has been established and characterized in previous years for use as a flexible experimental research facility across multiple applications. The system uses a similar plasma generator design to its twin-facilities at the University of Stuttgart (IPG6-S) and the University of Kentucky (IPG6-UKY). The similarity between these three devices offers the advantage to reproduce results and provides comparability to achieve cross-referencing and verification. Sub- and supersonic flow conditions for Mach numbers between M a = 0 . 3 − 1 . 4 have been characterized for air, argon, helium and nitrogen using a pitot probe. Overall power coupling efficiency as well as specific bulk enthalpy of the flow have been determined by calorimeter measurements to be between η = 0 . 05 − 0 . 45 and h s = 5 − 35 MJ kg−1 respectively depending on gas type and pressure. Electron temperatures of T e = 1 − 2 eV and densities n e = 1 0 1 8 − 1 0 2 0 m−3 have been measured using an electrostatic probe system. At Baylor University, laboratory experiments in the areas of astrophysics, geophysics as well as fundamental research on complex (dusty) plasmas are planned. The study of fundamental processes in low-temperature plasmas connects directly to electric propulsion systems. Most recent experiments include the study of dusty plasmas and astrophysical phenomena and the interaction of charged dust with electric and magnetic fields. In this case, dust can be used as a diagnostic for such fields and can reveal essential information of the magneto-hydrodynamics in low-temperature plasmas. Although some of these goals require further advancement of the facility, it is proposed that several phenomena relevant to electric propulsion as well as to other fields of plasma physics can be studied using the existing facility.

A promising aspect of inductive plasma generators for space propulsion applications is their electrodeless nature that leads to a high propellant compatibility and opens the door to in-situ resource utilisation propellants. To further develop inductive plasma generators as a space propulsion technology it is important to understand the relationship between system inputs (e.g. input power and propellant material) and system outputs (e.g. power efficiencies) to aid design and control. In this work, methodologies are developed for non-intrusively assessing the inductive coupling from antenna current measurements in terms of the ‘instantaneous ignition power’, the ‘inductive duty cycle’ and the ‘inductive frequency shift’. Experimental results are presented for IPG7, a high-power (up to 180 kW) 5.5-turn helical antenna inductive plasma generator. Potential in-situ resource utilisation materials oxygen and carbon-dioxide, and their respective mixtures with argon, are assessed in terms of their inductive coupling characteristics and resulting power efficiencies. Oxygen is shown to be an attractive propellant material, especially when supplemented with argon. In particular, high thermal efficiency ( ∼ 84 % ) can be achieved at high input powers. A basis is provided for understanding the power efficiencies of an inductive plasma generator in terms of non-intrusive antenna current measurements, to aid power supply sizing and to develop monitoring and control techniques for thruster applications.

This study focused on the development of a vacuum cathode arc thruster (VAT), particularly on its design, manufacturing process, and demonstration. Characteristically, the proposed thruster does not require any additional propellant feeding system as the cathode electrode is simultaneously used as a propellant. In the ignition system, tiny spots are coated on the cathode surface to induce plasma flow. Such a setup has the advantages of simplicity, low price, small size, and low weight and is suitable for microsatellites. Moreover, a “trigger-less” method with an inductor storage power system was used for generating the pulsed plasma. This discharge method can significantly reduce input power. Thrust is mainly caused by high exhaust velocities of metal ions in the plasma flow, making ion density, ion velocity, and ion charge influential parameters. A battery was used instead of a power supply system to reduce the energy consumption of the entire VAT processing unit. The energy required for a single pulse was estimated to be 0.266 J, by measuring the change between the discharge current and the voltage. The ion current was measured using an ion detector and was 3.55 A, and the ion velocity was 23150 m/s. In the theoretical analysis, the VAT prototype proposed in this study achieved a single impulse of 4.31 μNs, a specific impulse of 1571 s, and a thrust efficiency of approximately 12.5%.

Characterization of IEC tight jet mode is achieved through Faraday probe measurements. Preliminary results indicate that the tight jet is a highly energetic electron beam with scattering of secondary ions and electrons, which result from electron beam impact ionization. A novel analytical model is proposed to evaluate this non-Maxwellian plasma beam, including a compensation of the secondary electron emission effect on the probe's surface. The results show prominent features on the extracted jet's kinetic energy and the respective (electron) current of IEC. In addition, they demonstrate a practical method to characterise non-Maxwellian plasma jets through Faraday probes.

The objective of this work is to study the behavior of surface layers of high-temperature metals in their interaction with a powerful pulsed plasma flow produced by a high-power ablative pulsed plasma thruster. This plasma generator produces plasma flows with a directed velocity of (7–9) × 106 cm/s, an initial diameter of 1.5–2 cm, and a maximum number density of about 1018 cm–3, as well as a maximum power of 5 GW. The main measured values are the residual temperature of the tungsten specimens and the evaporated mass. Also, metallographic analysis of the specimens was performed. The basis of the research method is to analyze the experimental data with the help of a numerical model describing the heating and evaporation of the material upon absorption of pulsed energy fluxes taking into account the evaporation kinetics based on the Hertz–Knudsen expression. Based on the developed numerical model and the obtained experimental data, the kinetics of evaporation of tungsten at high power fluxes to the surface (up to 1 GW/cm2) is investigated.

The inertial electrostatic confinement (IEC) has been commonly researched in the scope of application for space propulsion and fusions reactors since many years. The utilization as thruster is possible due to a free emitting stream in a so-called jet mode of the IEC source. Depending on the setup, this jet can be operated as high energetic electron beam (tight jet) with electron energies of several keV or as low energetic plasma jet close to quasineutrality (spray jet). These modes are of high interest for thin film applications and plasma treatment. However, the IEC source is still not used in this field of application. Relevant application scopes can be electron substrate heating, surface plasma pretreatment, ionized physical vapor deposition and plasma enhanced chemical vapor deposition. In this paper, a cylindrical IEC source is presented that emits the plasma in both known jet modes, spray jet and tight jet. The jet emission occurs in radial outlet direction along the entire source height (“curtain-like” emission). Furthermore, an ignition test and operational behavior test is introduced. A stable spray jet operation region was examined and dependencies on pressure and voltage are investigated.

Silicon Carbide (SiC) is a low Z material discussed as an alternative to graphite for fusion devices. The retention of hydrogenous species is an important plasma-surface interaction property. Deuterium was implanted into SiC, SiCf/SiC, Cf/C-SiC and SiC coated graphite under various particle energy and substrate temperature conditions. A TDS process was used to characterise the deuterium retention of the implanted specimens. While all SiC materials show elevated retention levels compared to graphite, the differences are limited to about a factor of two over the range of parameters investigated. Keywords: Deuterium Retention, Silicon Carbide, Ceramic Composite Materials, Low Z Materials

To extend the lifetime of commercial and scientific satellites in low Earth orbit (LEO) and below (100–250 km of altitude) recent years showed an increased activity in the field of air-breathing electric propulsion as well as beamed-energy propulsion systems. However, preliminary studies showed that the propellant flow necessary for electrostatic propulsion at these altitudes exceeds the mass intake possible within reasonable limits, and that electrode erosion due to oxygen flow might limit the lifetime of eventual thruster systems. The pulsed plasma thruster (PPT), however, can be successfully operated with smaller mass intake and at relatively low power. This makes it an interesting candidate for air-breathing application in LEO and its feasibility is investigated within this paper. An analysis of such an air-breathing PPT system shows that for altitudes between 150 and 250 km, drag compensation is at least partially feasible assuming a thrust-to-power ratio of 30 mN/kW and a specific impulse of 5000 s. Further, to avoid electrode erosion, inductively heated electrothermal plasma generator technology is discussed to derive a possible propulsion system that can handle gaseous propellant without unfavorable side effects. Current technology can be used to create an estimated 4.4 mN of thrust per 1 mg/s of mass flow rate, which is sufficient to compensate the drag for small satellites in altitudes between 150 and 250 km.

Research of magnetoplasmadynamic Pulsed Plasma Thrusters (PPTs), or iMPDs, at the Stuttgart Institute of Space Systems (IRS), led to the thruster design ADD SIMP-LEX. For optimization of the thruster's discharge behavior and of the plasma acceleration, the magnetic self-field is measured. Inductive miniaturized probes are introduced and used to acquire data at points of the volume inside of the thruster's discharge channel. The results are used for analysis to identify the extend of the discharge zone along the discharge channel of the PPT. Measured magnetic field signals along the centerline of the discharge channel are presented. The underlying distribution of the dynamic discharge current is deduced from Amperes circuital law and interpreted in a two-dimensional plot against camera images for the current-plasma-interaction, current motion and validity of today's discharge models. The evaluation of the calculated integral discharge current in a plane of the discharge channel against the recorded thruster discharge current is presented and discussed.

The pulsed plasma thruster of the magnetoplasmadynamic type represents a promising and cheap solution to today’s propulsion challenges in space. The work presented in this paper covers new information on the electromagnetic conditions during thrust generation by revealing the temporal and spatial distributions of the current density vectors between the thruster electrodes. This is achieved by building and calibrating miniaturized induction probes. The local magnetic field components of all three spatial directions are combined to derive the current density vector. Quantitative determination of the temporal behavior of vectors at 27 points evenly distributed in three planes delivers a comprehensive picture of the position and development of the current in a pulsed plasma thruster. Maximum current densities of 80 kA/m2 and magnetic flux densities of up to 0.82 T have been observed close to the surface of the solid Teflon® propellant. It is concluded that no distinct movement of a current sheet was observed...

In the present work, two different concepts for fusion based space propulsion are compared. While the first concept is based solely on propulsion by hypothetic ejection of fusion products and hence may be called ash drive, the second one uses an additional coolant for thrust enhancement. Since this coolant was initially assumed to be gaseous and since it is doing most of the propulsion work, the name of “working gas drive” has been proposed. Propulsive characteristics for both types are evaluated for four fusion reactant couples (D–T; D– 3 He; 3 He– 3 He; 11 B–p). In working gas drives, only hydrogen is considered as coolant due to its exceptionally good caloric and propulsive properties. The results of comparative studies show that while ash drives excel working gas drives in terms of specific impulse the latter yield considerably more thrust than ash drives. Another major drawback of the ash drives is relatively small thrust efficiencies. The plasma power has to be disposed of nearly entirely as waste heat leading to prohibitive radiator masses.

In a pulsed plasma thruster discharge, the propellant utilization efficiency is commonly seen as the weak point for its thrust performance, and for its low thrust efficiency respectively. However, the value depends strongly on operational settings and the design of the pulsed plasma thruster itself. To access this efficiency to understand further the influence of parameters, a method combining experimental and modeling efforts is proposed within this paper. By using various optical methods, including high-speed camera imaging, emission spectroscopy, and Mach–Zehnder interferometry, the plasma bulk velocity for different energy configurations of the SPPT-68 ADD SIMP-LEX is derived. Together with the total ablated mass and an improved slug model, the ratio of mass accelerated in the discharge process is computed and, thus, the propellant utilization efficiency is derived. Velocities in the range of 25 to 55 km/s were measured, leading to propellant utilization efficiencies of about 40 to 60% depending on t...

We apply a particle code (Direct Simulation Monte Carlo, DSMC) to a nozzle flow expansion set-up which is typical for thermal plasma spray systems. Although those systems tend to have a pressure level which is too high to be treated with DSMC we obtain good results with respect to Pitot pressure measurement based Mach number estimations at different axial distances from the nozzle exit plane. Precisely, we compare the Mach numbers at 0, 30, 60, and 90 mm distance from the nozzle exit plane which was chosen as inflow boundary. The relative deviation of the simulated Mach numbers (Ma) from the measured is mostly about 10%. An influence of the flow field upstream due to a static boundary condition downstream is also observed. The vital code extensions which allowed this simulation with DSMC are briefly discussed. Further code extensions for future research activities are outlined.

The pulsed magnetoplasmadynamic thruster SIMP-LEX (Stuttgart impulsing magnetoplasmadynamic thruster for lunar exploration) is developed to serve as a main propulsion system for the lunar mission BW1. Optimizing the thruster's performance is necessary to increase the efficiency and the mean exhaust velocity without increasing the system's mass and energy requirements. Results of experiments varying the flare angle and the gap between the electrodes, the applied voltage, as well as using tongue-shaped or rectangular electrodes are shown and evaluated by means of a statistical approach using Taguchi matrices. Furthermore, an analytical approach toward the influence on the change in inductance along the x axis for these parameters was taken and the resulting performance characteristics are presented. Both tongue shape and flare angle, show a positive effect on the performance, that is, impulse bit, mean exhaust velocity, and thrust efficiency. A maximal I bit is given for a flare angle of ≈20 deg.

The magnetic field between the electrodes of an 80 J ablative pulsed plasma thruster was both measured and determined analytically. This was done to better understand the acceleration process and assess the accuracy of the analytical method. The measurements at different positions in both breech-fed and side-fed propellant configurations were performed using a shielded induction probe of 1 mm in diameter. To calibrate the probe, a Helmholtz coil was built. The magnetic field was calculated using the law of Biot-Savart and assuming a current sheet thickness of 3 mm. The measured magnetic field showed an overall peak at 0.7 T. It was possible to confirm the induced current loop formed between the electrodes at the time when the overall current passes through zero. The comparison between the magnetic field model and the measurements at the propellant surface of a breech-fed thruster showed reasonably good accordance.

The development of the inductively driven plasma wind tunnel PWK3, which enables the electrodeless generation of high-enthalpy plasmas for the development of heat shield materials required for space vehicles performing entry manoeuvres in the atmospheres of Venus, Earth and Mars, is described. The facility with its modular inductive plasma generators allows operation with gases such as carbon dioxide, air, oxygen and nitrogen and was qualified for thermal plasma powers up to 60 kW. Previously developed models for determining plasma properties and plasma source related characteristics enable a maximum plasma power in combination with long operational periods using different operational gases and gas mixtures. This is achieved by an optimization using the optimum operational frequency, a minimization of field losses using very thin plasma tube wall thicknesses and the successful application of MHD effects. Based on the solved cylinder problem for ICPs, a one-dimensional model for radial Lorentz forces and magnetic pressure has been developed. Here, a synthesis of previously published data and works is made where the new algebraic model for the calculation of Lorentz forces and magnetic pressures in an ICP was used and applied to experimental data. In addition, results from the model using the experimental data are shown to be consistent and, in addition, a comparison with a simpler model based on the well-known exponential approach for ICPs showed that the simpler model is covered without fail by the new model. The new model also states that there is a maximum of the Lorentz forces over the damping parameter d/δ (plasma diameter divided by skin depth) which almost corresponds with the position of the maximum plasma power of the cylindric model for ICPs. For the magnetic pressure the position of the maximum pressure is identical to the value for d/δ for the maximum plasma power.

Cp = specific heat at constant pressure, J=kgK hchem = chemical potential, MJ=kg h0 = specific total enthalpy, MJ=kg h0 = time-averaged specific total enthalpy, MJ=kg I = probe laser intensity, mW=mm I0 = incident laser intensity, mW=mm 2 M = Mach number _ m = mass flow rate, g=s P = input power, kW pamb = ambient pressure, Pa ptube = pressure in the discharge tube, Pa R = gas constant, J=kgK r = radial position, mm T = translational temperature, K T0 = total temperature, K t = elapsed time, ms u = flow velocity, m=s = specific heat ratio = wavelength, nm = laser frequency, Hz 0 = center absorption frequency, Hz = density, kg=m = fluctuation cycle, ms

The method for determining the recombination coefficients for ceramic materials in air plasma and pure oxygen plasma is described. Additionally, a new method lor the in situ determination of the spectral emissivity of high-temperature ceramic materials is presented. Values of the recombination coefficients for considered materials, resulting from the described methodology, are given at temperatures between 1483 and 1851 K and pressures between 356 and 950 Pa and compared with literature. Values of the spectral emissivities of the investigated materials are presented in a temperature range from about 1200 to 1820 K and compared with values obtained by other authors. Obtained results will be used for the development of catalytic sensors in the frame of the European reentry program EXPERT.

Laser absorption spectroscopy was applied for diagnostics of inductively heated plasma generator flows. Temporal variation of translational temperature was deduced from measured absorption line broadening because the flow properties fluctuated at 300 Hz in synchronization with the induction current. The specific total enthalpy and mole fraction of oxygen were estimated from the deduced temperature assuming thermochemical equilibrium. Consequently, the averaged degree of dissociation of oxygen is 0.92. The specific total enthalpy was estimated at 33.7 ± 2.9 MJ/kg; 39% of it was in the form of chemical potential. The results show good agreement with intrusive measurements.

Laser absorption spectroscopy was applied to pure oxygen flows generated by an inductively heated plasma generator IPG3. Temporal variation of translational temperature was deduced from measured absorption line broadening of OI 777.19nm because the flow properties fluctuated at 300Hz in synchronization with the induction current. The specific total enthalpy and mole fraction of oxygen were estimated from the deduced temperature assuming thermo-chemical equilibrium. Consequently, the averaged degree of dissociation of oxygen is 92%. The specific total enthalpy on the centerline was estimated at 33.7±2.9MJ/kg; 43% of it was in the form of chemical potential. Results show good agreement with intrusive measurements.

The thruster ADD SIMP-LEX (ADvanceD Stuttgart Impulsing Magneto-Plasmadynamic Thruster for Lunar EXploration) has been developed at the Institute of Space Systems (IRS) in Stuttgart, to provide secondary and primary propulsion capabilities to small satellites. Extensive optimization led to a mature design of world leading thrust efficiency around 30 % with russian APPT devices. The thruster was chosen for development into a stand-alone system in 2010. The system is capable of operating under satellite-like vacuum conditions in a test chamber, foregoing the usual outside laboratory equipment. The only connections required for system operation are a serial command interface, thermal housekeeping and a 28 V power line. Telemetry of the system for thruster health monitoring includes measurement of the pulsed current signal at up to four thruster capacitors. Requirements of the thruster and operation conditions ruled out any available COTS solutions for this task. A custom designed current measurement system was developed at the ETH Zurich and successfully integrated into the ADD SIMP-LEX design at IRS. The measurements revealed the current signal for every capacitor terminal and allowed for reconstruction of the combined flow. This data showed a current ringing effect between capacitors, which is undesirable for efficient thruster operation. The capacitor setup was changed and impact on the ringing was analyzed. The results showed reduction of the effect. Although complete elimination was not achieved, the data suggest careful consideration of the power train configuration can achieve satisfying mitigation levels. The sensor data quality was significantly improved by overcoming electromagnetic compatibility issues through introduction of a fiber-optical (LWL) transfer interface.

The examination of a permeable flow element for PPT propellant feeding is presented. The intention is to identify the control parameters for a self-controlled propellant flow through the element. As the element, a permeable \(\hbox{WHIPOX}^{\circledR}\)-fibre ceramic is used, feeding water propellant. It is found that the propellant transport is both driven by pressure gradients within the structure and evaporation into the discharge region. For functional testing, a mass flow rate of 340 μg/s is chosen, the PPT being operated against 5 Pa ambient pressure and showing reproducible ignition behaviour. Finally, the system is compared to a PTFE propellant system by spectrometric and photographic analysis of the plasma and by comparison of the electrodes indicating less erosion for water operation.

The pulsed plasma thruster Add Simp-lex, developed at the IRS, was investigated at the University of Tokyo to further characterize the thruster’s performance and discharge behavior. This was done by experimental investigation of configurations with a different amount of capacitors and discrete applied voltages. To do so, a measurement system for the discharge current and the optical properties was built up and successfully applied. Further, the numerical model for the prediction of the current-normalized magnetic flux density was improved and the convergence properties of the integration towards the change in inductance studied. From the discharge current waveforms and the pictures taken from the propagating plasma, information about the amount of plasma creations, their propagation velocity and the oscillation behavior was deducted. For further characterization, the energy transfer efficiency and the electrical efficiency was derived from these data, leading to a tool to compare different configurations. It was found, that a middle voltage yields higher electrical efficiencies of about 40% whereas the energy transfer efficiency is higher the lower the applied voltage.

The Institut fur Raumfahrtsysteme has developed the combined sensor system COMPARE for SHEFEX II. The goal of this system is to simultaneously measure the temperature of the thermal protection system, the radiative heat flux and the absolute pressure. The gained data will allow the reconstruction of important trajectory parameters. Furthermore, due to the fast response capabilities of the detectors used in the sensor system it is possible to determine the dynamic behavior of the vehicle during flight. With regard to upcoming re-entry missions with relative high entry velocities the application of COMPARE on SHEFEX II will allow to increase the Technology Readiness Level of this sensor system. In this paper, the design and layout of the sensor system are presented.

Estimation of the local dissociation degree and the local mass-specific enthalpy of a pure oxygen plasma flow determined mainly from laser-induced fluorescence measurements are reported. Measurements have been conducted for several generator parameters in an inductively heated plasma wind tunnel. Additional probe measurements of total pressure together with the deduced translational temperature are used to estimate the local mass-specific enthalpy. For a reference condition, full dissociation has been measured. The measured translational temperature of atomic oxygen for this condition is T = 3500 K. Subsequently, the local mass-specific enthalpy has been derived using these local density and temperature measurements. For the reference condition the estimated value of h = 27 MJ/kg is in good agreement with the probe measurements and results from diode laser absorption spectroscopy.

Laser absorption spectroscopy was applied for diagnostics of inductively heated plasma generator IPG3 air and nitrogen flows. Although no absorption signal could be detected in the nitrogen flow, temporal variation of absorption profiles at OI 777.19 nm were measured in the air flow. As a result, the flow properties were found to fluctuate at 300 Hz with the plasma emitted duration of 46%. The specific total enthalpy and degree of dissociation in oxygen and nitrogen were estimated from the deduced total temperature assuming thermo- chemical equilibrium in the discharge tube. Consequently, the averaged total enthalpy, degree of dissociation in oxygen and nitrogen in the emitted duration time were estimated as 23.5±10.1 MJ/kg, 98% and 44 %, respectively.

The paper describes the development of the inductively driven plasma wind tunnel PWK3 enabling the electrodeless generation of high enthalpy plasmas for the development, investigation and qualification of heat shield materials required for space vehicles performing entry manouevres in the atmospheres of Venus, Earth and Mars. The facility, equipped with modular inductive plasma generators enabling the operation with gases such as carbon dioxide, air, oxygen and nitrogen has been qualified for thermal plasma powers up to 60 kW. The measurement of these powers, coil currents, operational frequencies, electromagnetic fields, heat fluxes on probes, Pitot pressures and Mach numbers allows the extensive characterization of both the plasma generators’ operational behavior and the plasma flow conditions. Newly developed models and numerical simulations, that correspond pretty well with the experimental data, enable a maximum plasma power in combination with long operational periods using different operational gases and gas mixtures by optimizing the plasma container wall thickness, the number of coil turns and the number of capacitors in the resonant circuit.

Inductive Plasma Generator developed at the Institut fur Raumfahrtsysteme (IRS) was used to produce high enthalpy pure oxygen plumes. Translational temperature was obtained by laser absorption spectroscopy using an absorption line from meta -stable atomic oxygen at 777.19nm. As a result, emission and absorption signals were found to fluctuate at 300Hz. Maximum translational temperature w as 6,700K on the centerline and 4000K at the edge of the plume. Then, the spatially averaged static enthalpy and kinetic enthalpy estimated by measured temperature and corresponding frozen loss calculated under thermo -chemical equilibrium were 6.5MJ/kg 5.5 MJ/kg and 15.5MJ/kg, respectively . Then, the total enthalpy was estimated to be 27.5MJ/kg. This value agrees with input enthalpy of 30MJ/kg.

The operating parameters and resulting plasma properties of a blind hollow cathode (BHC) discharge have been investigated. The hollow cathode was driven capacitively with a pulsed dc signal of 200 kHz in a power range between 50 and 100 W at an ambient pressure of about 10 Pa. The working gas was argon, which was introduced with a ceramic capillary at different positions of the longitudinal axis of the hollow cathode with flow rates of between 30 and 1000 sccm. The current–voltage characteristics were recorded.The pressure at the end of the BHC was measured with a miniaturized pressure transducer with varying volumetric flow rate and axial position of the capillary in the hollow cathode. To characterize the ignition behaviour of the system, the measured breakdown voltages were compared with phenomenological Paschen curves calculated from the pressure data.Optical emission spectroscopy was used to examine the origins of the light emission, comparing the glow mode and hollow cathode mode in particular.A high-speed camera recorded some plasma processes. A mounting with an indium tin oxide coated glass was used to observe the inner volume of the BHC along the longitudinal axis, while the plasma was operated with different parameters. The optical observations revealed an inhomogeneous plasma condition along the axis.

Pulsed plasma thrusters are electric space propulsion devices which create a highly transient plasma bulk in a short-time arc discharge that is expelled to create thrust. The transitional character and the dependency on the discharge properties are yet to be elucidated. In this study, optical emission spectroscopy and Mach-Zehnder interferometry are applied to investigate the plasma properties in variation of time, space, and discharge energy. Electron temperature, electron density, and Knudsen numbers are derived for the plasma bulk and discussed. Temperatures were found to be in the order of 1.7 to 3.1 eV, whereas electron densities showed maximum values of more than 1017 cm−3. Both values showed strong dependency on the discharge voltage and were typically higher closer to the electrodes. Capacitance and time showed less influence. Knudsen numbers were derived to be in the order of 10−3−10−2, thus, indicating a continuum flow behavior in the main plasma bulk.

An overview of the IRS payload development for ESA's EXPERT mission is given. The final design and performance parameters of the payloads PYREX, PHLUX and RESPECT are described. PYREX is a sensor system measuring the thermal protection system (TPS) rear side temperature. PHLUX is a catalysis based experiment to determine the dissociation degree of the boundary layer. RESPECT applies optical emission spectroscopy to measure spectrally resolved the radiation onto a TPS surface.

The research and development process of the ablative pulsed plasma thruster SIMP-LEX is summarized. By experiment, the magnetic field in between the electrodes was measured with an induction probe, and a high-speed camera was applied to estimate the propagation of the plasma plume. Further, the optimization process towards higher thrust efficiency by means of geometric and electrical modifications is described. By analysis, the calculation of magnetic field and change in inductance is summarized and the numerical models for the performance simulation of SIMP-LEX introduced. A peak value of 0.7 T was found for the magnetic field created during an 80 J discharge whereas a propagation velocity of 28 km/s was estimated for the first plasma plume. The optimization process towards a new engineering model yielded an increase of 40% in mean exhaust velocity to 26 km/s, and a maximum thrust efficiency of 32% was evaluated respectively. Calculation of magnetic field and change in inductance based on the Biot-Savart law yield input for the slug model, which can then predict the general performance parameters with a good accordance.

1 . Among experiments conducted by American and Japanese researchers, a contribution of Germany was made by the SLIT experiment. Here, a conventional spectrometer in a Cerny Turner configuration (f=300mm, grating 600 l/mm) was fed by fiber optics with light collected by a small mirror telescope. The telescope was able to collect light from a viewing angle of only 0.45 degrees. Manual tracking was performed using a tracking camera with a viewing angle of 20 degrees. A wavelength range from 324 nm to 456 nm was covered with a pixel resolution of 0.08 nm. Data was taken during 30 s around the point of maximum heating until the capsule left the observed region with a measurement frequency of 5 Hz. Due to the difficulties in tracking not every spectrum contained data. Nevertheless, the emission of CN as a major erosion product as well as N2 + and different atoms could be monitored successfully during that time. Due to the nature of the set up, no spatial resolution of the radiation data was achieved so all measured values represent an integration both over the visible part of the glowing heat shield and over the plasma in the post shock region. The measured spectra could be split up into continuum spectra which represent a superposition of the heat shield radiation and the continuum radiation of potential dust particles in the plasma and into line spectra from the plasma in the post shock layer. Planck temperatures as well as rotational and vibrational temperatures of CN and N2 +

The appliedeld magnetoplasmadynamic thruster AF-MPD ZT1 was successfully put in operation at IRS. The thruster operated in steady-state mode with Argon as propellant at low electric arc power of 6 kW. Discharge voltage and power increased as expected almost linear with applied magnetic ux density. The variation of mass ow rate ratio between anode and cathode gas towards higher cathode gas fraction showed increasing thrust and thrust e ciency at applied magnetic ux density of 0:1 T and are comparable with DLR’s X13 thruster. Additionally a steady-state 100 kW AF-MPD thruster SX3 was developed, set in operation and preliminary characterized at IRS. The SX3 thruster was operated at relatively low arc powers up to 30 kW and applied magnetic ux density of 0:1 T generated by the modi ed ZT coil. Due to low arc current and magnetic ux level, the AF-MPD ZT1 thruster achieved thrust of 70 mN and exhaust velocity up to 10 km/s at 6 kW arc power and up to 6 % thrust e ciency. The SX3 thruster reached thrust of 362 mN and thrust e ciency more than 12 % in steady state-operation at 25 kW arc power and at applied magnetic ux density of 0:1 T. However both thrusters have been operated at limited magnetic uxes only. For SX3 a total operation time of more than 3600 s together with 30 ignitions could be accumulated. The electrodes, however, do not show signi cant erosion nor a respective degradation. The performances of thrusters provide an outlook for future investigations on AF-MPD thrusters at IRS and give a hint to improvement in thrust e ciency of presented devices with the new appliedeld coil, which will be manufactured in the future to produce magnetic ux densities up to 0:6 T allowing further increase in thrust e ciency up to 30 %.

The production of hi gh enthalpy flows in a stationary manner requires plasma wind tunnels using plasma ge nerators. Magneto plasmadynamic generators (MPG), thermal arcjet devices (TPG) and inductively hea ted plasma generators (IPG) have been de ve loped at the IRS for differ ent atmospheres, different atmos pheric entry regimes and/or different basic in vesti gations. The absence of electrodes enables the use of IPGs for basic TPS material tests (e.g. catalycity) and the simulation of re -active atmospheres such as the Martia n or Venerian atmosphere. With IPG3 an rf -source is provided where the coil is closer to the plasma than it was with previous de signs. Therefore, the electromagnetic field loss is re duced. The water cooling system surrounds both the induction coil and th e plasma container. This paper describes the design of IPG3 and the experimental results using O 2. Three test cases with pure O 2 plasma were investigated using the absorption spectroscopy system of Department of Advanced Energy , the University of Tokyo, in combination with a Pitot pressure probe and a material sample support system of IRS, Universitat Stuttgart. Number density and translational temperature distributions of atomic oxygen were obtained from the measured absorption line profiles including abso rption lines of meta -stable atomic oxygen at 777.19 and excited atomic oxygen at 844. 64 nm. Mach number was deduced by the ratio between pitot pressure and ambient pressure. Surface temperature of TPS materials was measured by thermocouplers . 1. Introduc tion A variety of stringent requirements upon technology are imposed by both space travel to a rather close Earth orbit and to more distant celestial bodies such as the planets of our solar system. Among those requirements there are the high temperature ef fects occurring during atmospheric entries. Here, spacecrafts encounter gases at velocities of more than ten km/s, thereby being subjected to great heat loads. This makes it necessary to use high temperature thermal protection systems (TPS) to prevent the destruction of the space vehicle. Both TPS and environment (plasma) during the entry have to be investigated by means of computational and ground facility based simulations. Such ground facilities are the five IRS plasma wind tunnels PWK1 -5 that reproduce the thermal, aerodynamic and chemical load on the surface of a space vehicle entering a celestial body’s atmosphere. MPGs enable gas flows on high enthalpy levels. They are mainly used to investigate the erosion of radiation cooled heat shield materials li ke C -C or C -SiC as well as the behavior of ablative ma terials under thermal and chemical loads. Two IRS plasma wind tunnels are operated with MPGs. Furthermore, a TPG producing moderate enthalpies and higher stagnation pressures to simulate the follow on flight path is in use [1, 2]. As an IPG is contactlessly heated by induction currents, exceptionally pure plasma flows of any gases can be produced. Therefore, the application of an IPG delivers two major advantages. First, researchers are aware that the catalytic behavior of TPS materials is one of the main topics to be investigated in future. A main part of the heat flux seen by a TPS material derives from the recombination of atomic plasma species. Using the IPG the related heat flux measurements are not falsified by impurities as with the electrode plasma generators. Furthermore, the catalytic mechanisms of single gases such as of O 2 and N 2 [3, 4] in the case of air of an atmosphere can be investigated. Fig. 1 shows a typical pure oxygen plasma plume of IPG3 at 500Pa ambient pressure, an oxygen mass flow rate 3g/s and at an anode voltage U A=6250V. The resonant circuit was adjusted to nominal 650kHz using a 5 -turn coil with 7 capacitors each of 6nF (see description below).