Hannah Böhrk

Heat Balance of a Transpiration-Cooled Heat Shield

HEATS is a layout tool for the determination of transient wall heat flux to a transpiration-cooled parallel flat plate under laminar or turbulent flow conditions. The method is based on heat balances between wall material, transpired coolant, and surrounding hot gas. It is introduced here with particular focus on a reentry mission. The investigation of transpiration cooling necessitates considering the entire flight trajectory. Because of long calculation times, computational fluid dynamics tools cannot be used. The calculation of wall temperatures of an entire trajectory with HEATS takes less than five minutes, which offers the possibility of sensitivity analyses, e.g., of coolant mass flow rate. The results presented within this paper are compared with cooling experiments under laminar flow and a computational fluid dynamics solution of one trajectory point for the second Sharp-Edge Flight Experiment trajectory for turbulent flow. The values compare well with deviations of the wall temperature below 10% and deviations of heat flux below 18%. The investigation of coolant mass flow rates by means of HEATS show that for the entire second Sharp-Edge Flight Experiment trajectory the transpiration-cooling experiment should be run with 0,1 g/s < m < 1 g/s.

Thrust Measurement of the Hybrid Electric Thruster TIHTUS by a Baffle Plate

A novel, two-stage, electric thruster is under development at Institut fiir Raumfahrtsysteme at Universitat Stuttgart. The first stage is an arcjet, and the second stage uses inductive heating of the arcjet plume. Because of the complex setup of the thruster, it is impossible to mount the two-stage system on a thrust stand without causing unwanted momentum onto the stand by the numerous supply lines. Thrust is, therefore, inferred using a baffle plate by means of a variation of power- and mass-flow ratios between the two stages. To interpret the results, the present paper also provides radially resolved measurement of total pressure from which thrust can be determined. It is shown that the thrust obtained from the two measurement methods are in good agreement with each other and range between 1.7 and 2.5 N for the respective operating conditions of the thruster. At constant total input power of 50 kW to the thruster, thrust decreases as power is diverted from the first stage to the second (inductive) stage at a constant mass-flow rate. However, when power is applied to both stages, thrust increases as the mass-flow rate is diverted to the second stage at a constant total mass-flow rate of 300 mg/s.

Transpiration Cooling at Hypersonic Flight - AKTiV on SHEFEXII

The present paper presents the in-flight measurement of the transpiration cooling experiment AKTiV flown on the sub-orbital re-entry configuration SHEFEXII. Thermal response of the structure is measured just up- and downstream of a cooled sample with a non-cooled reference set-up on the opposite side of the vehicle. The measurement shows that the heat flux is reduced upon coolant exhaust. The maximum temperature reduction by transpiration-cooling is observed on the porous sample with 87 K, while downstream the sample, the temperature reduction by film cooling is 75 K. This corresponds to cooling effi�ciencies of 61% and 46% on the sample and downstream, respectively. The evaluation is supported by the semi-analytical tool HEATS, based on a transient heat balance at the surface. Comparison of the results with HEATS shows that the heat flux predicted with HEATS is in good concurrence with the measured temperatures on the cooled sample, as well as upstream and downstream with a maximum deviation by 14%. The analysis shows that the flow condition around the sharp-edged, faceted vehicle remains longer laminar than expected.

Thermal Testing of the Sharp Leading Edge of SHEFEXII

The octahedral sharp leading edge of SHEFEXII is made from C/C-SiC fiber reinforced ceramic. It is instrumented by three thermocouples and eight pressure ports. A duplicate was ground-tested in arc-heated wind tunnel L3K. Transient thermal response of the material to the load was also determined by the heat-balance method HEATS and compared to the measurement data. During ground-testing, the hermocouples measured 1179°C, whereas the maximum temperature recorded during flight was measured at 30km altitude with 848°C. The simulated data corresponds well to both thermocouple data and data from an external infrared camera. During ground-testing, the very tip of the leading edge reaches 1635°C. The nose was in good order after ground testing.

Efficiency Analysis of a Two-Stage Hybrid Electric Thruster

A RCJET thrusters have been developed throughout the past few decades. They provide relatively high thrust at moderate exhaust velocities, and their plasma flows are characterized by high specific enthalpy and high flow velocity. These characteristics are, however, combined with the presence of steep radial property gradients as in a hot, energy-rich core with a relatively cold gas layer at its edge. When the limits to the nozzle exit velocity are imposed bywall and electrode materials, another possibility to transfer more power into an arc heated plasma flow is to reheat the relatively cold plume edges by means of another heating mechanism, as done in an afterburner. TIHTUS has been developed over the past four years and is a novel two-stage plasma thruster where reheating of the arc heated plume is realized bymeans of inductive heating, as sketched in Fig. 1 [1–3]. It is therefore considered the predecessor of a future propulsion system for the transport of heavy payload on interplanetary trajectories [1]. The principal question within the development of this thruster is whether it is possible to specifically heat the outer edges of an arcjet plume so that higher exhaust velocity can be attained. A dependency of the power staging between the two thruster stages and of the gas mass flow rate staging is expected. For the presented research, the thruster was investigated with a cavity calorimeter which provides data of plasma power so that a total efficiency can be derived. Moreover, a method for deducing the two-stage system’s efficiency from operational data is presented and its data are compared with the measured calorimetric data.

Flowfield and Pressure Decay Analysis of Porous Cones

A transpiration-cooled cone is investigated with respect to varying cooling mass flow along the surface caused by the orthotropic permeability of a ceramic composite material and varying wall thickness. The measurements comprise the dynamic pressure distribution at the cone surface, the reservoir pressure, and the overall mass flow rate. Material permeability depending on fiber orientation is measured on cylindrical samples. Furthermore, a numerical approach with the two-dimensional Darcy equation is presented to simulate the pressure loss across the orthotropic wall structure and the mass flow distribution along the surface. Finally, the results from the experiment and the simulation are compared and assessed with respect to the fiber orientation of the material. The resulting mass flow distributions demonstrate the flow dependence on the fiber orientation. Furthermore, the outflow behavior at the surface is investigated, focusing on the exhaust angle, and the results of the measurements coincide with th...

Experimental and Numerical Analysis of SHEFEX-II

The SH arp E dge F light EX periment (SHEFEX) of the German Aerospace Center (DLR) has been established to demonstrate the feasibility of space vehicles with facetted Thermal Protection System (TPS). In this study, first results to the aerothermodynamic behavior of SHEFEX − II in the High Enthalpy Shock Tunnel Gottingen (HEG) are presented.The main focus is on the experimental and numerical investigation of the active cooling experiment AKTiV. The measured and calculated results confirm that a cooling effect is feasible.

Kinetic Parameters and Thermal Properties of a Cork-Based Material

In order to model the thermal response of recessing or degrading material, the rate of thermal decomposition and various physical properties must be known. A cork-based ablative material is investigated for the rate of thermal decomposition from room temperature to 1823 K. Thermogravimetric analysis (TGA) measurements are obtained at different heating rates and several isoconversional methods are used to determine the activation energy, and a parallel distributed activation energy model was used to describe the overall rate of decomposition. The parameters of the model were calculated using a pattern search optimization method, and predicted behaviors of the ablator at additional heating rates are discussed. Results from differential thermal analysis (DTA), laser flash analysis (LFA) and scanning electron microscopy (SEM) are also presented here.

Experimental Investigation of the Plasma Jet Generated by ATTILA

In the inductive stage of the DC-HF hybrid thruster Attila, power characteristics are measured and interpreted. This is done with two different induction coils of 2.5 and 4.5 turnsunder variation of ambient pressure. Test gases used are Nitrogen and Oxygen. At the present operation frequencies, generator power is coupled into the plasma with higher efficiency for Nitrogen than for Oxygen. Efficiencies determined for the inductive plasma generator are 44% and 43% in total.