Erich Schülein

Experimental and Numerical Modeling of the Bow Shock Interaction with Pulse-Heated Air Bubbles

The influence of various parameters characterizing the single- and double-pulse energy deposition (ED) upstream of a blunt body (distance between the deposition point and the body, amount of the energy and time delay between the pulses) on the topology of the supersonic flow at Mach 2 is studied experimentally and numerically. The obtained pressure-time diagram at the upstream stagnation point of the body as well as some significant topological properties of the bow-shock / heated-bubble interaction, like shock deformation and reflection, as well as the evolution of the heated bubble into the vortex ring downstream, are analyzed as resulting from the shock-decelerated spherical light-gas inhomogeneity. The evolution and topology of the interfering heated bubbles by double-pulse ED show the significance of vorticity generation initiated by blast-waves coming from the neighboring bubbles.

Steady longitudinal vortices in supersonic turbulent separated flows

Large-scale longitudinal vortices in high-speed turbulent separated flows caused by relatively small irregularities at the model leading edges or at the model surfaces are investigated in this paper. Oil-flow visualization and infrared thermography techniques were applied in the wind tunnel tests at Mach numbers 3 and 5 to investigate the nominally 2-D ramp flow at deflection angles of 20◦, 25◦ and 30◦. The surface contour anomalies have been artificially simulated by very thin strips (vortex generators) of different shapes and thicknesses attached to the model surface. It is shown that the introduced streamwise vortical disturbances survive over very large downstream distances of the order of 10000 vortex-generator heights in turbulent supersonic flows without pressure gradients. It is demonstrated that each vortex pair induced in the reattachment region of the ramp is definitely a child of a vortex pair, which was generated originally, for instance, by the small roughness element near the leading edge. The dependence of the spacing and intensity of the observed longitudinal vortices on the introduced disturbances (thickness and spanwise size of vortex generators) and on the flow parameters (Reynolds numbers, boundary-layer thickness, compression corner angles, etc.) has been shown experimentally.

Comparison of experimental and numerical investigation on a jet in a supersonic cross-flow

Flow interaction of three different jet configurations ejecting air from a flat plate into a supersonic cross flow were investigated experimentally and numerically. The test conditions encompassed a jet pressure ratio of P 0j , P ∞ = 100 at a Mach number of M ∞ = 5 and Reynolds number of about Re ∞ = 25 x 10 6 based on the length of the flat plate. The investigated test cases are: a) single jet; b) four jets positioned in-line in main flow direction; c) four jets positioned side-by-side in spanwise direction. The prediction of the overall flow phenomena as occurring within the interaction area was in fair agreement with the experiments, although quantitatively differences occur that will be discussed in the paper. The results of the comparison are presented and the experimental data are used to validate the applied code.

Wave Drag Reduction Approach for Blunt Bodies at High Angles of Attack: Proof-of-Concept Experiments

Proof-of-concept experiments for a spike aligned into the oncoming flow at the nose of a supersonic blunt body were performed to investigate its aerodynamic effectiveness over a wide range of bodys angles of attack. The investigations have been conducted in the Ludwieg Tube Facility at DLR Gottingen at Mach 2, 3 and 5 for angles of attack from 0 to 30 degrees. The model tested is a cylindrical body with a hemispherical nose. Additionally to the body equipped with an aligned spike, a reference body without a spike and a body with a conventional fixed spike were investigated. Furthermore, a test model with a prototype of a pivoting self-aligning spike-device has also been investigated. The results containing shadowgraph visualizations, direct force measurements and infrared heat flux measurements show the clear advantages of the spikes aligned into the flow.

Effects of Laminar-Turbulent Transition on the Shock-Wave/Boundary-Layer Interaction

This paper presents results of experimental investigations of the nominally 2-D impinging shock-wave interaction with the transitional boundary layer developing on a flat plate at Mach 6 flow conditions. Experiments were conducted in the Ludwieg-Tube Facility at DLR Gottingen using quantitative infrared-thermography, high-speed shadowgraphy and free-stream Pitot-pressure fluctuation measurements. Natural-transition and shockimpingement locations on the flat plate were varied independently in order to establish different initial boundary-layer conditions at the interaction region. The results demonstrate significant impact of these mutual positions on the induced flow topology and on the heat-flux distribution. A range of CFD-validation test cases was created representing transition phenomena in some details. The distribution of the normalized Stanton number, generalized in the present work for a wide range of unit Reynolds numbers, indicate a clear narrow maximum inside the transition region.

Wave drag reduction by means of aerospikes on transonic wings

Wave drag and shock induced boundary layer separation are important issues of transonic wings. The negative effect of the transonic flow regime can be mitigated by controlling the shock terminating the supersonic region above the wing. In the past many different concepts based, for example, on passive ventilation, active suction, contour bumps or on adaptive walls have been pursued (see references). These approaches have in common that measures for controlling the shock are applied directly at the surface of the wing. However, a control of the shock wave is also possible by external devices placed above the surface of the wing in the supersonic flow regime. Experiments related to the latter concept that is related to the one of aerospikes on blunt bodies, will be presented in the present contribution. In a test series the effectiveness of a variety of different spike-shaped bodies placed above a transonic wing was tested in the transonic wind tunnel DNW-TWG, Gottingen. In addition to pressure measurements a colour schlieren system was set up for providing information about the influence of spikes on the flow field. The drag reduction mechanism of spike-shaped bodies that are placed in the supersonic flow above a transonic wing is based on the generation of wake flows and oblique shock waves interfering with the normal shock terminating the supersonic region. In this manner the pressure increases more gradually thus limiting losses. Since the spike is located above the surface of the wing the boundary layer on the wing is not directly disturbed. In the streamwise direction the exact position of the spike is of less importance than in the case of measures applied at the surface of the wing. The height at which the spike is arranged above the surface is chosen so that the shock is especially weakened in its lower part where the shock strength is greatest. Typical dimensions depend on the chord length and the Mach angle. Similarly, in order to weaken the shock over the whole span width several spikes are placed next to each other in spanwise direction. Bodies of various geometries that are acting as wake and shock inducing spikes, have been studied. Results are reported that were obtained with a cylindrical body having a needle-like tip, a punctured pipe that was open at its leading edge and a cone. A 400mm-chord model of the transonic airfoil VC-opt was mounted in the 1m x 1m adaptive walls test section of the TWG. Initially, a single spike was placed on the suction side of the VC-Opt model, the tip of the spike being located about in the middle of the chord. A colour schlieren system was set up for observing the flow field. A comparison of colour schlieren pictures of the flow about a clean airfoil and the flow about an airfoil with a conical spike shows only little differences. This is due to the span width (1m) being much greater than the size of the spike (diameter about 12 mm). However, a shock wave and Mach lines originating from the tip and the surface of the spike, respectively, can well be seen. Lift and drag were determined by pressure measurements. On the wing the static pressure was obtained via pressure taps that were arranged in a slightly diagonal manner thus avoiding interferences between the taps. The drag was calculated from total pressure data obtained by wake-rake measurements one and a half chord lengths behind the trailing edge. Initially, the rake was laterally displaced with respect to the spike by ten percent of the chord length. Tests were performed at a Reynolds number of Re ≈ 5⋅106 and at two Mach numbers, M = 0.775 and M = 0.795, respectively. Lift and drag polars were obtained for different configurations such as a clean airfoil and for an airfoil with shock inducing bodies. At certain angles of attack a gain in lift is observed and the drag is clearly reduced. This effect is most pronounced for a conical spike, i.e., for a body which produces a notable displacement in the flow. Hence, the concept of using aerospikes on transonic wings clearly shows a potential for reducing wave drag.

Experimental Investigations on the Phantom Yaw Effect on a maneuvering slender body

Phantom yaw can be induced by, for instance, small perturbations or deviations in geometry that cause an asymmetric separation. Experiments were conducted investigating this phenomenon under both static (no model motion) and dynamic (pitching motions at almost realistic pitch rates) conditions. The tests were performed at M=0.8 and ReD=4.2×105 using a generic missile configuration. The effect of small geometric deviations on the phantom yaw was studied by changing the model’s roll angle. Additionally, the possibility of triggering asymmetric vortices by a fixed, asymmetric transition at the model’s nose and shoulder was investigated. Indeed, the asymmetry can be triggered but the mechanisms are complex and the effects not always distinct. Furthermore, two phantom-yaw-reduction devices were tested. These were “conventional strakes” and “aerostrakes” (two symmetrical side jets at the model’s shoulder). In combination with a fixed asymmetric transition, both devices revealed deficiencies and were not able to...

Effects of Localized Flow Heating by DC-Arc Discharge Ahead of Non-Slender Bodies

The influence of steady energy addition into the flow by low-voltage DC-arc discharge located upstream of conically-nosed and spherically-blunted bodies was investigated experimentally in the Ludwieg-Tube Facility at Mach 5. The results include drag force measurements and shadowgraph flow visualizations. The flow-field structure arising due to the bow-shock/heated-wake interaction, as well as the effects of bow-shock intensity and heating power variation on the drag reduction of different non-slender bodies are analyzed in this paper. The results demonstrate the existence of an optimum heating rate providing a maximum effectiveness of energy deposition and show distinct drag reductions up to 70% dependent on the test conditions and model geometry.

Experimental Investigation of Laminar Flow Control on a Supersonic Swept Wing by Suction

The experimental results of the investigations of the transition-delay on a supersonic swept wing by suction in the framework of the European research project SUPERTRAC are presented in this work. The experiments were performed for a variety of parameter settings at Mach 2 in the Ludwieg Tube Facility at DLR Gottingen. Three test-cases have been investigated: two for the 30°-sweep at unit Reynolds numbers of 17 mio/m and 25 mio/m, and one for the 20°-sweep at unit Reynolds numbers of 30 mio/m. In all these cases the suction velocity was varied as a free parameter in the range from zero up to approximately 1.0m/s. Modern measurement techniques, which are essential for the detection and investigation of transition, such as the global skin-friction interferometry (GISF) and the infrared (IR) thermography, are used for the present investigations. The results obtained include wall pressure and heat flux distributions, as well as transition locations and limiting streamlines pattern.

Numerical investigation on the development of the Phantom Yaw Effect on a maneuvering missile

The aerodynamics of a pitching slender body at