Henning Rosemann

Numerical and Experimental Validation of Three-Dimensional Shock Control Bumps

Numerical and experimental studies have been performed to show the potential for drag reductions of an array of discrete three-dimensional shock control bumps. The bump contour investigated was specifically designed by means of CFD-based numerical optimization for wind tunnel testing on a modern transonic airfoil. The experimental investigations focused on turbulent flow at a Reynolds number of 5 million and were carried out at the

Experimental investigation of trailing-edge devices at transonic speeds

The influence of trailing-edge devices such as Gurney flaps and divergent trailing edges of different height on the aerodynamic performance of an airfoil at transonic speeds has been investigated experimentally. The investigation has been carried out in the Transonic Wind Tunnel Gottingen (TWG) using the two-dimensional airfoil model VC-Opt at freestream Mach numbers of M element of [0.755, 0.775, 0.790] and a Reynolds number of Re = 5.0 · 10 6. The results have shown that the trailing-edge devices increase the circulation of the airfoil leading to a lift enhancement and pitching-moment decrease as well as an increase in minimum drag compared to the baseline configuration. The maximum lift-to-drag ratio is considerably improved and the onset of trailing-edge flow separation is shifted to higher lift. Besides the increased rear-loading, a downstream displacement of the shock provides the main lift enhancement in transonic flow. The simple Gurney flap provides the largest additional circulation of all geometries tested. The smoother turning of the flow due to the additional ramp of the divergent trailing edge leads to a smaller increase of circulation. Slightly less lift but considerably less viscous (pressure) drag is generated enhancing the maximum lift-to-drag ratio compared to the Gurney flap. The negative affect of the Gurney flap on the pitching moment is also reduced. For the high divergent trailing edges, different ramp slopes have a significant influence on the aerodynamic performance whereas at low device heights the influence is considerably diminished. The results show that the divergent trailing edge proves to be the better trailing-edge device at transonic speeds. The application as an element for an adaptive wing is generally possible.

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.

Simulation of Oscillating Airfoils and Moving Flaps Employing the DLR-TAU Unsteady Grid Adaptation

The use of unsteady adaptation with the DLR-TAU Navier-Stokes solver is presented as a method of improving the modelling of flows where the aerodynamic performance of a body is determined by the action of moving localised regions of high-gradient flow. Examples are presented of transonic limit cycle oscillation and dynamic stall. First results indicate that good grid convergence can be achieved without necessarily requiring that the flow around the specific airfoil is well understood in advance.

Prediction of the Wind Tunnel Sidewall Effect for the iGREEN Wing-Tailplane Interference Experiment

Computations using the DLR TAU code are presented to predict the effect of the wind tunnel on models which are directly attached to the wind tunnel side-wall. Results show that the effect of the sidewalls is similar to an offset in the angle of attack (but can only partially be corrected by a change in α), and that this effect is proportional to the lift and Mach number. These results were used in the design phase for the iGREEN wing-tailplane interference experiment.

Numerical Simulation of Wind Tunnel Wall Effects on the Transonic Flow around an Airfoil Model

For wind tunnel measurements in closed-wall test sections, possible interference effects of the wind tunnel walls play an important role. Three-dimensional TAU simulations were performed for the transonic flow around an airfoil model in the adaptive-wall test section of the Transonic Wind Tunnel Gottingen (DNW-TWG) to investigate the existence of wind tunnel wall effects. The results revealed a global side wall interference that is affecting the entire flow around the model and a local side wall interference disturbing the flow in the supersonic regime.

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.

On Transition Detection at High Subsonic Freestream Mach Numbers Using Thermoresistive Surface Sensors

This paper looks at the results of characterizing laminar-turbulent boundary layer transition on a 2D transonic wing profile. The experiments were conducted in the Transonic Wind Tunnel Gottingen (DNW-TWG) with a surface hot-wire and a hot-film sensor array. The aim of the experiments was to detect the Tollmien-Schlichting waves at high subsonic freestream Mach numbers. Therefore, thermoresistive sensors were used, as they offer a high temporal and spatial resolution. Varying the total pressure in the DNW-TWG makes it possible to achieve a constant Reynolds number at different freestream Mach numbers and thus to examine the influence of the freestream Mach number on the transition independently of the Reynolds number effect. The Mach number was increased up to M∞ = 0.72 with the Reynolds number related to the chord length of Rec = 7.7 · 10 6 . Furthermore, the angle of attack as well as the Reynolds number were varied to investigate their influences on transition.

Gurney Flaps in Transonic Flows

The application of Gurney flaps to airfoils, high aspect ratio wings and delta wings in transonic flow is being investigated at the Institute of Aerodynamics and Flow Technology of DLR. The present paper gives an overview over these studies explaining the basic working principles and advantages of these trailing edge devices at transonic Mach numbers. Numerical studies confirmed by wind tunnel experiments show similar global effects as conventional trailing edge flaps, with the additional advantage of a reduction of the adverse pressure gradient on the upper side of the airfoil towards the trailing edge.