Helmut Kurtz

PLASMA GENERATORS FOR RE-ENTRY SIMULATION

The qualification of thermal protection systems (TPS) and numerical design tools for re-entry vehicles and space probes requires the ability to understand and duplicate the prevailing complex physico-chemical phenomena, including thermal and chemical nonequilibrium near the surface of a body that enters the atmosphere of the Earth or another celestial body. At the Institut fur Raumfahrtsysteme of the University of Stuttgart, four plasma wind tunnels (PWK1-4) are in operation to simulate the thermal, aerodynamic, and chemical loads on the surface of a space vehicle. Three different plasma sources have been developed for this purpose: 1) a magnetoplasmadynamic generator for the simulation of the highenthalpy and low-pressure environment during the first phase of re-entry, 2) a thermal arcjet device for the follow-on flight path at moderate specific enthalpies and higher stagnation pressures, and 3) an inductively heated generator for basic materials experiments over a wide range of specific enthalpies and pressures. Special efforts were made to avoid electrode erosion to preclude impairing the erosion and catalytic behavior of TPS materials. A detailed description of these plasma generators and an overview of the simulation regions and operation areas of the plasma wind tunnels are presented.

High-Power Hydrogen Arcjet Thrusters

A radiation-cooled thermal arcjet thruster named HIPARC-R has been developed and investigated. It has been designed for the 100-kW power level and is operated with hydrogen as its propellant. A specie c impulse of 1970 s was obtained with a mass e ow rate of 150 mg/s at the 100-kW power level and at a thruster efe ciency of about 28%. This equals a specie c input power value of 670 MJ/kg. Parallel to the experiments a numerical code system was developed to further optimize the next generation of hydrogen arcjet thrusters. This code system consists of a e nite volume e ow code coupled with program modules for the calculation of thermal, chemical, and electronical properties. In addition, a program module for the calculation of the heat e ow inside the thruster, including heat exchange, has been applied to model the heat transfer processes during thruster operation. The thruster has been operated over a wide power range and has been intensively investigated for the qualie cation of the numerical code system. Within this paper the experimental setup and the code system are described, the performance data are presented, and experimental and numerical results are compared.

Numerical Modeling of the Flow Discharge in MPD Thrusters

Two theoretical models for calculating the current and flow distributions in self-field MPD thrusters have been developed and are applied to evaluate the effects of geometry, propellant type, scaling, and other parameters on the thruster performance. For continuous thrusters, a stationary code has been developed. The extended Ohms law is used to calculate the current contour lines, and a one-dimensional, two-component expansion flow model is applied to obtain the velocity, temperature, and pressure distributions for calculating the gas properties, which are again used in Ohms law. An integration over the volume and thermal forces equals the thrust. The differential equation is solved by means of a finite-difference method for the geometry of the nozzle-type plasma thruster DT2-IRS, which has been investigated experimentally in a steady-state as well as in a quasi-steady-state mode. The calculated current density distribution and the computed thrust are compared with these experimental results. For the starting phase of the steady-state MPD thrusters as well as for pulsed thrusters, a time-dependent, fully two-dimensional code has been developed. It uses a modified McCormack FD method in cylindrical coordinates to calculate the time-dependent flow, temperature, and pressure fields.

Arcjet Thruster Development

For several years an intensive program has been in progress at the University of Stuttgart to investigate and develop thermal arcjets for propellants including ammonia, nitrogen-hydrogen mixtures simulating hydrazine, and hydrogen. Since hydrogen yields the highest specific impulse /sp and best efficiencies TJ, special emphasis was placed on this propellant. Arcjet power levels between 0.7-150 kW have been studied, including water- and radiation-cooled laboratory models and flight hardware. Results yielded a maximal attainable 7sp as a function of the design and power level and showed that increasing power increased /sp. Radiation-cooled arcjets show better 17 and 7sp than water-cooled devices, but raise technical problems because of the high temperatures of the thrusters, which require the use of special refractory materials. Proper arcjet optimization was done with a thorough thermal analysis, including the propellant flow. A further improvement of these thrusters was reached by regenerative cooling and by optimizing the constrictor contour. The constrictor flow is modeled by a three-channel model, the results of which are compared with experimental data. A new two-dimensional computational fluid dynamics (CFD) approach for hydrogen arcjet thrusters is presented. In 1996 a 0.7-kW ammonia arcjet is scheduled for a flight on the P3-D AMSAT satellite.

Optimization of Electric Thrusters for Primary Propulsion Based on the Rocket Equation

IEPC-01-181 During the last decade, electric propulsion systems have been established for orbit maintenance of satellites. More than 150 spacecrafts are now equipped with almost 400 thrusters for this purpose. The current decade will see the use of electric propulsion for primary propulsion. The purpose of this paper is to determine optimal mission parame ters for these tasks. Depending on the mission profile, ion thrusters, Hall thrusters, thermal arcjets or MPD thrusters are preferable. All electric propulsion systems have in common that they can be operated in a wide range of the specific impulse and that the thrust efficiency and therefore also the specific power of the propulsion system depend strongly on the specific impulse. The optimal specific impulse for a particular mission depends, therefore, on the kind of thruster and the chosen propellant. This paper shows that for MPD, ion, Hall ion and thermal arcjet thrusters the optimal specific impulse for a particular mission can be determined by an optimization which is based on the rocket equation. Using, in addition, a simple cost function, the influence of the cost factors is explained. Finally, the results for a few missions for which electric propulsion systems for primary propulsion have been selected are discussed.

Optimization of electric propulsion systems considering specific power as function of specific impulse

The deployment of large electric power supplies in space in the near future allows the use of electric propulsion for transfer missions. The purpose of this paper is to determine optimal mission parameters for these tasks. In this paper, contrary to previous investigations, the dependency of the specific power on the specific impulse has been taken into account. It shows that the optimal exhaust velocities are increased and the corresponding payload fractions decreased. The whole area of a possible optimization is diminished. To obtain a measure for the costs of missions, a transport rate is defined and included in the optimization. It further restricts the optimizable area and increases the specific impulse for higher velocity increments ( Av). For both optimizations, nomographs have been provided to map the whole parameter area of interest for transfer missions. The comparison of different thruster types shows that not only the maximal thrust efficiency but also the form of tjT(c) is highly important for the results.

Medium power arcjet thruster experiments

Experimental results on the electrothermal behavior and operating performance of water-cooled laboratory model arcjet thrusters using a variety of propellants in the range of 5-50 kW are reported. The highest voltage and specific impulse are attained with hydrogen propellant and the lowest with argon propellant; intermediate results are obtained with nitrogen and a mixture of hydrogen and nitrogen. The highest cathode erosion rate is measured with argon while the lowest is associated with hydrogen.

100 kW hydrogen arcjet thruster experiments

The High Power Arcjet thruster is investigated with respect to the effects of constrictor diameter, cathode gap, and cooling mechanism on specific impulse and general engine performance. The thruster is mounted on a thrust balance and integrated into a stainless steel vacuum chamber, and cathodes with 5-, 10-, and 14-mm diams and varying tip configurations are tested. The results are analyzed with specific attention given to cathode erosion, varied propellant-injection angles, the influence of tank pressure, and pressure data for the arc chamber. The thruster with a 4-mm constrictor is more critical than the 6-mm version, and the best results are obtained for the 4-mm cathode gap. The highest specific impulse for this configuration is 1300 s at 200 mg/s corresponding to an input power of 66 kW. The smaller throat diameter is shown to lead to better performance characteristics for the High Power Arcjet thruster. 8 refs.

Numerical and Experimental Investigations of Magnetoplasmadynamic Thrusters with Coaxial Applied Magnetic Field

In this paper an overview is given on a numerical simulation program for applied field magnetoplasmadynamic (AF–MPD) thrusters, which is currently under development at the Institute of Space Systems (IRS). The program allows the simulation of argon plasma flows under thermal and chemical non–equilibrium. The code is based on an axisymmetric finite volume method on unstructured, adaptive meshes. An externally applied magnetic field can be taken into account employing the vector potential formulation. Azimuthal velocity and magnetic field are handled by a quasi–three dimensional approach with vanishing azimuthal derivatives. Besides the numerical analysis, a radiation–cooled laboratory model of an AF– MPD thruster is under development at IRS. The thruster model is to be assembled into a thrust balance. The modular design allows adjustments of the electrode and magnetic field configuration to find an optimized thruster geometry based on numerical and experimental results.

Investigations of advanced medium power hydrogen arcjets

The performance of thermal arcjet thrusters has to be improved sign