Stefan Hein

Surface-Pressure Measurements of Second-Mode Instability in Quiet Hypersonic Flow

Surface pressure-sensors have been used to measure the second-mode boundary layer instability on a 7 o half-angle sharp cone at zero angle of attack in the quiet Mach-6 wind tunnel at Purdue and in the conventional Mach-6 wind tunnel in Braunschweig. The measurements were made using a stream-wise array of high-frequency sensors. They show the second-mode waves in quiet and noisy flow. The quiet flow data is compared to results in noisy flow. The signal quality allows for the calculation of amplification rates, which are compared to the results of linear stability computations.

Complementary Numerical and Experimental Data Analysis of the ETW Telfona Pathfinder Wing Transition Tests

Within the European Project Telfona the Pathfinder Model was designed, analyzed numerically, constructed and tested with the aim of obtaining a laminar flow testing capability in the European Transonic Wind Tunnel (ETW). The model was designed for natural laminar flow (NLF) for transonic flow conditions with high Reynolds number. Results of pre-test numerical analysis demonstrated that the Pathfinder wing pressure distribution was adequate for providing calibration test points. The ETW tests provided pressure distribution data while transition positions were determined from images using the Cryogenic Temperature Sensitive Paint Method (cryoTSP). The evaluation of this data with several transition prediction tools was used to establish the transition N-factor values for ETW. In this work, after-test CFD solutions are obtained using numerical Navier-Stokes solutions. In the first part of this work, numerical results are given which verify the requirements of the Pathfinder wing as a calibration model. In the second part, it is shown that for selected flow conditions a good agreement is obtained between stability analysis based on experimental and numerical data. In the third part the correlation of experimental transition locations to critical N-factors is summarized for ETW Test Phases I and II. In the fourth part numerical analysis and experimental data are used complementarily.

Numerical and experimental investigation of laminar-turbulent boundary layer transition on a blunt generic re-entry capsule

Numerical and experimental results on laminar-turbulent transition in the boundary layer of a blunt Apollo-like capsule at 0° and 24° angle of attack are presented. Local stability analyses have been performed and a measurement campaign in the Hypersonic Ludwieg tube Braunschweig at a Mach number of 5.9 was carried out. Infrared thermography showed laminar and transitional surface heating in the unit Reynolds number range of Re_infinity = 6x10^6 /m to Re_infinity = 20x10^6 /m at a surface mean roughness of Ra = 10 micrometer, whereas for a mean roughness of Ra = 0.5 My micrometer no indications for a transitional boundary layer was noted. PCB and Kulite sensors used to measure pressure fluctuations inside the boundary layer do not show any peaks in the frequency spectra which might be related to boundary layer disturbances. The only relevant peak in the spectra does not change with unit Reynolds number and is currently attributed to a bow shock oscillation. Consistent with the experimental findings, the modal instability analysis does not provide any modal boundary layer disturbance growth at windtunnel conditions. Therefore, a scaling ansatz for the laminar boundary layer is introduced and evaluated in order to estimate the unit Reynolds numbers required for the onset of modal disturbance growth on Apollo-like capsules. Results for both first-mode and cross-flow instability are presented.

Pressure Gradient and Nonadiabatic Surface Effects on Boundary Layer Transition

The influence of the streamwise pressure gradient and a nonadiabatic surface on boundary layer transition was experimentally investigated at the Cryogenic Ludwieg-Tube Gottingen, Germany. Boundary layer transition was detected nonintrusively by means of the temperature-sensitive paint technique. The wind-tunnel model was designed to achieve a quasi-uniform streamwise pressure gradient over a large portion of the model chord length. This allowed the effects on boundary layer transition of the streamwise pressure gradient and wall temperature ratio to be decoupled. The model was tested at high Reynolds numbers and at a high subsonic Mach number. Favorable, almost-zero, and adverse streamwise pressure gradients were considered; and various temperature differences between the flow and the model surface were implemented. Stronger flow acceleration and lower wall temperature ratios led to an increase of the transition Reynolds number. Larger increases in the transition Reynolds number were obtained at more pron...

Wake flow instability studies behind discrete roughness elements on a generic re-entry capsule

Numerical and experimental results on disturbance growth in the wake flow downstream of a roughness element submerged in the boundary layer of a typical re-entry capsule at an angle of attack are presented. Laminar basic flow computations were conducted at Mach 5.9 freestream conditions of the Hypersonic Ludwieg Tube Braunschweig (HLB) for different roughness heights, widths and planform shapes. The modal instability characteristics of the wake flow were studied by spatial two-dimensional eigenvalue analysis and three-dimensional parabolized stability equations. For all cases considered, the varicose wake modes are most amplified in terms of maximum N-factors. A comparison of the disturbance kinetic energy production terms reveals that the wall-normal shear is mainly responsible for the dominant instability modes. A flush-mounted high frequency pressure transducer (PCB) is used to measure the wall-pressure fluctuation of the instability modes behind a cylindrical element at various Reynolds numbers. A comparison of the spectral distributions of pressure fluctuations with and without roughness clearly shows the appearance of distinct spectral humps.

The effect of pressure gradient and a non-adiabatic surface on boundary layer transition investigated by means of TSP

The influence of streamwise pressure gradient and a non-adiabatic wing surface on boundary layer transition was experimentally investigated at the DNW-KRG blow-down wind tunnel facility in Gottingen, Germany. Boundary layer transition was detected non-intrusively by means of the Temperature-Sensitive Paint technique. A new wind tunnel model was designed with the aim of systematically investigating the influence of streamwise pressure gradient and a non-adiabatic wing surface, including Reynolds number effects and Mach number effects, on boundary layer transition. The model was tested at high Reynolds numbers and at a high subsonic Mach number. Favorable, neutral, and unfavorable streamwise pressure gradients were considered and various temperature differences between flow and model surface were implemented. More pronounced negative streamwise pressure gradients and surface temperatures closer to the adiabatic wall temperature were shown to stabilize the boundary layer and allowed larger transition Reynolds numbers to be achieved. The resulting effect of the coupling of streamwise pressure gradient and a non-adiabatic wing surface was found to be strongly dependent on the considered stability situation. The favorable effect on boundary layer transition of surface temperatures closer to the adiabatic wall temperature was shown to be more pronounced for stability situations characterized by a markedly negative pressure gradient.

Redesigned Swept Flat-Plate Experiment for Crossflow-Induced Transition Studies

In this publication the redesigned DLR swept flat-plate experiment is introduced. It is based on the well-known crossflow reference experiment by Bippes et al. and consists essentially of a swept flat plate and a displacement body arranged above it. The differences between the old and the new setup and the reasons for the redesign will be described. The experiment is used for investigations on the mechanisms leading up to crossflow-induced laminar-turbulent transition in the flat-plate boundary layer and on different actuation methods for influencing them. The main features of the 3D boundary layer flow field such as the development of crossflow instabilities and the onset of high-frequency secondary instabilities during the laminar breakdown are presented.

High Reynolds Number Transition Experiments in ETW (TELFONA project)

A wind–tunnel experiment on laminar-turbulent transition has been performed in ETW (the European Transonic Wind Tunnel in Koln) at high Reynolds number and cryogenic conditions. The studied geometry is a sting mounted full model in swept–wing configuration. The transition location was determined by means of Temperature Sensitive Paint (CryoTSP). The experimental observations were further analysed using different transition prediction tools, based on linear stability theory.

Experimental investigation of Mach number and pressure gradient effects on boundary layer transition in two-dimensional flow

The influence of Mach number (\(M = 0.35\)–0.65), chord Reynolds number (\(Re_c=6\times 10^6\) to \(10\times 10^6\)) and pressure gradient (\(dc_p/dx = -0.6\)–0.07 m\(^{-1}\)) on laminar-turbulent boundary layer transition was experimentally investigated in the Cryogenic Ludwieg-Tube Gottingen (DNW-KRG). For this investigation the existing two-dimensional wind tunnel model, PaLASTra, which offers a quasi-uniform streamwise pressure gradient, was modified in order to reduce the size of the flow separation at its trailing edge. The streamwise temperature distribution and the location of laminar-turbulent transition were measured by means of temperature-sensitive paint (TSP). It was found that the transition Reynolds number exhibits a linear dependence on the pressure gradient, characterized by the Hartree parameter, and that an increasing Mach number leads to a linear decrease of the transition Reynolds number. The latter effect is likely due to an increase of the total pressure turbulence level with Mach number in DNW-KRG. The measured pressure and temperature distributions served as input for boundary layer calculations and linear-stability analysis. N-factors were calculated according to compressible and incompressible stability theory. At zero pressure gradient a critical N-factor of approximately 9.5 and 9.0 was found for incompressible and compressible calculations, respectively.

Experimental and Numerical Investigation of Instabilities in Conical Boundary Layers at Mach 6

Studies on the stability of three-dimensional hypersonic boundary layers for 7 and 15 deg. half-angle cones at 0 and 6 deg. angles of attack are reported. Measurements were carried out in the Hypersonic Ludwieg tube Braunschweig at Mach number 6. Surface mounted pressure and heat flux sensors were used to determine the spatial extension, frequency contents, and wave structure of second-mode and cross-flow instabilities. Infrared thermography was used to capture location, direction and growth of stationary cross flow instabilities. The experimental data are analyzed and compared to the results of corresponding stability computations.