Stefan Loehle

Ablation radiation coupling investigation in earth re-entry using plasma wind tunnel experiments

An overview of a comprehensive experimental study to investigate ablative materials in a high enthalpy air plasma flows including radiation effects is given. Through the application of surface thermometry, in-situ recession measurements with photogrammetry, optical emission spectroscopy from vacuum-ultraviolet (120 nm) to near infrared wavelengths (950 nm), pyrometry and thermography, a complete set of data for the investigation of ablation radiation coupling has been acquired. The paper presents the background of this project, the suite of experimental setups and first results for a carbon preform sample and a cooled copper sample. All systems acquired data and first comparisons to chemical equilibrium calculations have been assessed.

Numerical Investigation of the Re-entry Flight of Hayabusa and Comparison to Flight Data

This paper presents a comparison of re-entry ight data measured by the NASA/JAXA Hayabusa Observation campaign with numerical calculations and ground testing in the NASA Ames Electric Arc Shock Tube (EAST) facility and the University of Stuttgart plasma wind tunnel, Plasmawindkanal 1 (PWK1). The numerical calculations were conducted using the University of Stuttgart URANUS code including thermal and chemical non-equilibrium as well as radiation transport modeled by the programs HERTA and PARADE. The ow eld calculations were extended by an ablation calculation using FABL (Fluid Gravity Engineering Ablation code). These simulations were performed for 17 trajectory points chosen from the nominal ight trajectory provided by NASA. The ground testing in the

Heat Flux Calibration Measurement Using the Noninteger System Identification Method for Cooled Surfaces

ACTIVE cooling mechanisms are used in many applications for thermal protection of surfaces in combustion chambers, rocket nozzles, gas turbine blades, and only recently structures of reentry vehicles [1–4]. The basic principle is to inject a cooling fluid close to the wall to reduce the wall heat load. In literature, the three different effects, i.e., film, transpiration, or effusion cooling, are named to occur in actively cooled structures. These cooling mechanisms, however, complicate the problem of the determination of a heat transfer coefficient to thewall and thus a calculation of the heatflux to the wall becomes challenging. Although transpiration cooling has been proven theoretically as an effective cooling mechanism by many investigators, its technological application became recently of particular interest with the invention of high-temperature porous material [2]. The progress inmaterial development allows a very new approach to high-temperature environments. From rocket engine combustion chambers and nozzles to turbine blades, from thermal protection systems for reentry vehicles to fusion chamberwalls, in all these modern high-temperature environments, transpiration cooling is currently in the focus of research to overcome the principle problem of exceeding the wall temperature limit [4–8]. Regenerative cooling is state of the art in rocket engines. Liquid fuel is used to cool the nozzle and the heated fuel is injected in the combustion chamber at higher temperature, which again increases the combustion efficiency. Results of numerical analysis show that the efficiency of regenerative cooling compared with transpiration cooling can lower the wall temperature by more than 30%. Transpiration cooling means that the fuel is injected through the wall into the combustion chamber or nozzle. This is advantageous because a layer of relatively cool fuel close to the wall protects the wall [5]. With the invention of high-temperature porous media for rocket engines, those numerically proved improvements could be verified by experimental tests [3]. For high-speed aircraft, material and cooling issues for both airframe and engine are the key elements which force the designer to limit the flight Mach number [9]. All these engineering problems face the fact that the heatflux at the surface is a design critical value yet difficult to measure. In the high heat flux regimes, a direct measurement is not applicable, since no sensor is available to measure directly at the surface. The only solution is therefore, a temperature measurement using in-depth temperature sensors. However, an inversemethod has to be applied to determine surface heat flux. In the case of a transpiration cooling environment, however, the often applied one-dimensional analytical calculation is prone to error: The thermocouple position is of particular importance in this approach and particularly in a porous material very difficult to know with sufficient precision. Moreover, the porosity has to be known, which complicates this analytical problem further because there is not only the dependency from heat capacity, heat conductivity (both values needed for the porous material), but also from thematerial specific parameter of porosity. A numerical approach lacks a precise information of porosity of the material which is impossible to know for a particular sensor system. Finally, for such complex materials, the thermophysical properties are not known and difficult to determine. A recent publication of Shi andWang presents a solution for this inverse problem [10]. Here, the accuracy of themethod again depends strongly on the accuracy of the information about the specific heat capacity, heat conductivity, and coolant temperature. In this paper, an approach based on the calibration of the real environment will be presented. For the analysis of the calibration measurement, the noninteger system identification (NISI) approach is applied to a transpiration cooled environment. Using basic calibration experiments, the transpiration cooling is inherently covered by the NISI calibration process. It is shown for the first time that the knowledge of the coolant mass flow is sufficient to determine the surface heat flux.

Plasma Wind Tunnel Investigation of European Ablators in Air Using Emission Spectroscopy

Plasma wind tunnel tests using the European ablative materials AQ61 and MONA were conducted in a nitrogen/ oxygen atmosphere at three relevant heat fluxes. Results of the emission spectroscopic investigation of the plasma radiation are presented, for both the freestream (without probe body) and the boundary layer 5 mm in front of the test samples. Emission spectroscopy was conducted at several wavelength ranges (320 nm < λ < 580 nm) and spectral resolutions. Results are compared with respect to the temperatures of the most common plasma species. Here, comparisons are drawn between the plasma emission at the different heat flux regimes and the test specimen material.

Aerothermodynamic investigation of inductively heated CO2 plasma flows for Mars entry testing

This paper presents the characterization of an inductively coupled CO2 plasma relevant for Mars entry. Key plasma parameters are measured using Two Photon Absorption LaserInduced Fluorescence (TALIF), Optical Emission Spectroscopy (OES) and a High-Speed Camera (HSC) as non-intrusive diagnostic methods. In addition heat flux, enthalpy and total pressure are intrusively probed to supplement the characterization. TALIF provides data about translational temperature, velocity and the density of ground state atomic oxygen. From OES rotational and vibrational temperatures for identified molecular species and excitation temperatures for atomic species are derived. Preliminary HSC results show the pulsing behaviour of the generator and, in combination with bandwidth filters, atomic emission distributions. The diversity of the applied measurement techniques offers an extensive characterization of the flow, hence enabling the identification of suitable test conditions for the early stages of Martian entries.

Comparison of Calorimetric Sensors - NASA Ames and IRS

A space act agreement between NASA Ames Research Center and the Institute of Space Systems at the University of Stuttgart was put in place in October 2011. The focus of this collaboration is the comparison and refinement of measurement techniques for plasma characterization. Within this paper, first test results from this collaboration will be presented, based on tests completed at NASA Ames arc jet facilities. A water-cooled energy balance calorimeter and a null-point calorimeter from the Institute of Space Systems were compared to a slug calorimeter and null-point calorimeter from NASA Ames. The results are expected to serve the plasma and entry community by providing comparative data between the two facilities and the three types of sensors.

Three-Dimensional Thermal Analysis of the HIFiRE-5 Ceramic Fin

The German Aerospace Center (DLR) has developed a ceramic fin experiment (FinEx) for the HIFiRE-5 flight in order to test the performance of new structures with sharp leading edges during flight. The analysis of the thermal performance of the fin has been foreseen to be conducted by in-depth mounted thermocouples protocolling temperatures during flight. In a cooperative attempt between DLR and the Institut fur Raumfahrtsysteme (IRS) of the University of Stuttgart, the thermal behavior is studied. In this paper, an approach is presented to determine the heat flux distribution onto the surface of the three-dimensional geometry of the fin. Using the Non-Integer System Identification (NISI) method, heat flux onto different surface elements is derived from only three thermocouples mounted inside the fin. Moreover, to increase accuracy, an area-weighted inverse method is used. The principal proof of concept is given based on a finite element thermal analysis of a cube. Measurements with a backup fin are used to show the approach experimentally in a plasma wind tunnel. Finally, the in-flight data has been evaluated and heat flux onto the fins during flight is presented. It can be concluded that the method is strongly dependent on surface discretization. However, the qualitative distribution is accurately determined. This approach allows to improve heat flux measurements for sparsely equipped flight hardware.

Laser Induced Fluorescence Spectroscopy on Neutral Xenon: Two-Photon Cross Sections and Measurements in an Ion Thruster Plume

Two-photon absorption laser induced fluorescence (TALIF) measurements in neutral xenon are presented both in a cold gas cell and in the plume of the radiofrequency ion thruster RIT-10. Measurements in the cold gas cell concentrate on the characterization of TALIF schemes involving levels of the 6p and 6p multiplets with respect to natural lifetimes, collisional deactivation coefficients and two-photon absorption cross sections. The feasibility of two TALIF schemes including the 6p[1/2]0 and 6p [3/2]2 levels has been demonstrated in the plume of the RIT-10. Relative particle densities have been measured at a distance of 12 cm downstream the acceleration grid in the center of the plume axis.

Analysis of Wavelength Modulation Spectroscopy for Water Vapour Measurements in Supersonic Combustion

In this paper measurements of temperature and mole fraction of water vapour by wavelength modulation spectroscopy (WMS) are presented. This work is focused on hypersonic combustion diagnostics using tunable diode laser absorption spectroscopy (TDLAS) at 1995 nm. The experiments are performed with a vertical cavity surface emitting laser diode (VCSEL) which is tunable by current and temperature. Water vapour temperature and mole fraction are determined by matching the second harmonic (2f) peak height normalized by its first harmonic (1f) value to a database of simulated signals based on the characteristics of the laser. The measurements have been performed on flat flames for proof of concept. Two different flame conditions are investigated and the results are compared with direct absorption measurements using the same equipment and the application of a calibration free WMS method to the acquired data as presented by Rieker et al. The investigated flames have been extensively analysed using coherent anti-stokes Raman scattering (CARS) measurements and equilibrium calculations by Prucker et al. The results using direct absorption agree fairly well with the published CARS data. The applied WMS methods agree well with each other, however, there is a significant discrepancy to the direct absorption measurements at the high temperature condition. Since the resulting values for the two different flame conditions with a nominally large temperature difference are close to each other, independent from the applied WMS method, it is concluded that further research of the WMS technology is required with respect to the applied transitions and its sensitivity to the investigated flows.

Comparison of Heat Flux Gages for High Enthalpy Flows - NASA Ames and IRS

This article is a companion to a paper on heat flux measurements as initiated under a Space Act Agreement in 2011. The current focus of this collaboration between the Institute of Space Systems (IRS) of the University of Stuttgart and NASA Ames Research Center is the comparison and refinement of diagnostic measurements. A first experimental campaign to test different heat flux gages in the NASA Interaction Heating Facility (IHF) and the Plasmawindkanäle (PWK) at IRS was established. This paper focuses on the results of the measurements conducted at IRS. The tested gages included a flat face and hemispherical probe head, a 4” hemispherical slug calorimeter, a nullpoint calorimeter from Ames and a nullpoint calorimeter developed for this purpose at IRS. The Ames nullpoint calorimeter was unfortunately defective upon arrival. The measured heat fluxes agree fairly well with each other. The reason for discrepancies can be attributed to signal-to-noise levels and the probe geometry.