In the Transonic Wind Tunnel Göttingen, pitch oscillations were performed with a Lambda wing to study unsteady pressure distributions of vortex dominated flow including transonic effects. The free stream Mach number was varied between 0.3 and 0.7. Small pitching amplitudes of \(0.08^{\circ }\)–\(0.4^{\circ }\) at excitation frequencies up to 40 Hz were used. In this paper, particularly the data of unsteady Pressure Sensitive Paint measurements and unsteady pressure sensors are analyzed. With increasing angle of attack, a suction peak and a shock occur near the leading edge. Then a shock-induced separation triggers the development of a vortex at the main wing. The unsteady pressures show: for lower angles of attack, the transonic influences are dominant. For higher angles of attack, the influence of the vortex becomes of similar magnitude and dominates the behavior of the pressure variations. The shock exhibits, with increasing angle of attack, an inverse motion. For angles of attack beyond the maximum lift, the unsteady pressure distributions and the lift show a significant phase lag, already at very low oscillation frequencies. Compared to subsonic cases, the supersonic region shifts the vortex induced pressures downstream.

The development of vortical flow strongly changes the properties of motion-induced aerodynamic loads. Tests were performed in the Transonic Wind Tunnel Gottingen with a half-wing model of a lambda wing. A hydraulic actuator imposed pitch oscillations with amplitudes and frequencies relevant for aeroelastic analyses. Unsteady surface-pressure and flowfield measurements were performed. With increasing angle of attack, a leading-edge vortex emerges and moves inboard. The unsteady loads show a strong dependency on the effects of the vortical flow. For the pitching cases, the effects are lagging the motion. In the rear outboard part of the wing, the unsteady pressures are small, but the amplitude of the motion is high. Hence, a significant amount of energy-force integrated over motion-is transferred from the fluid to the motion. For a free motion or deformation of the wing, this would cause a destabilization. A significant effect of the vortical flow is the discrepancy between the position where it emerges and is affected, and the position (e.g., further outboard) where the flow may generate unsteady, potentially destabilizing loads.

In the Transonic Wind tunnel Gottingen, different test campaigns were performed with an oscillating lambda wing to study the unsteady pressure distributions at vortical flow conditions including transonic influences. The free stream Mach number was varied between 0.3 and 0.7. Small pitching amplitudes of 0.08° - 0.4° at excitation frequencies up to 40 Hz were used. In this paper, particularly the data of unsteady Pressure Sensitive Paint measurements and unsteady pressure sensors are analyzed. With increasing angle of attack, at first a suction peak and a shock occur near the leading edge, then a shock induced separation triggers the development of a vortex at the main wing. The unsteady pressures show the following: For lower angles of attack, the shock is dominant. For higher angles of attack, the effect of the vortex is of similar magnitude and becomes dominant. The shock exhibits an inverse motion with increasing angle of attack. For high angles of attack, beyond the maximum lift, the unsteady pressure distributions and the lift exhibit a significant phase lag, already at very low oscillation frequencies.

Aim: Some years ago, in the framework of the Motar Cooperation Agreement between ONERA and DLR, it had been agreed to carry out comparison testing of the two instationary PSP methods on a common test configuration with an unsteady flow produced in a suitable wind tunnel. The results from this comparison would form the basis for a benchmarking of the two different approaches, leading to an aid in decision-making as to which method is better suited for a particular application. / Test configuration: For many years there has been interest at ONERA on the (quasi) two-dimensional unsteady transonic flow in a channel with strong interactions between an oscillating shock and a separated boundary layer: various diagnostic methods were used in measurements on their S8Ch transonic wind tunnel at ONERA Meudon (Schlieren, pressure sensors, PIV, LDV), including also PSP with Ru/AA paint (uPSP). Flow separation of the boundary layer is produced over a convex contour profile (a bump) on the floor of the tunnel, with the subsequent flow acceleration leading to formation of a downstream shock wave. The sectional area of a second throat, situated downstream of the bump, is varied periodically (ca. 15 Hz or higher) by rotation of an elliptical rod (cam) placed downstream of the throat position, leading to forced upstream pressure fluctuations and hence an oscillation (movement in flow direction) in the shock position. The resulting pressure fluctuations on the bump can then be measured with the help of pressure sensors (Kulite) fixed in the floor, and also with PSP coated on the floor. / Procedure/Approach: The floor can also be coated with the DLR iPSP paint. As with ONERA, there exists sufficient optical access for placement of the DLR excitation LEDs and cameras adjacent to or above the test section. PSP measurements were then be carried out with a new set of measurements using both ONERA and DLR paints, consecutively. / Results: Pressure measurements with Kulite and iPSP/uPSP are compared. Phase correlation is good, but there remain some differences in amplitude for both paints. Static and instationary calibration were carried out on samples of both paints; the latter showed no phase loss up to 120 Hz, although the amplitude with iPSP dropped due to surface heating by the LED over time. Fourier analyses of Kulite and PSP showed good agreement. A table was compiled with a summary of characteristics for both paints: there is no “better” paint, since the choice is governed mainly by desired application.

Dynamic stall on a pitching OA209 airfoil in a wind tunnel is investigated at Mach 0.3 and 0.5 using high-speed pressure-sensitive paint (PSP) and pressure measurements. At Mach 0.3, the dynamic stall vortex was observed to propagate faster at the airfoil midline than at the wind-tunnel wall, resulting in a “bowed” vortex shape. At Mach 0.5, shock-induced stall was observed, with initial separation under the shock foot and subsequent expansion of the separated region upstream, downstream and along the breadth of the airfoil. No dynamic stall vortex could be observed at Mach 0.5. The investigation of flow control by blowing showed the potential advantages of PSP over pressure transducers for a complex three-dimensional flow.

For aerodynamic profile tests on aircraft or wind mill models, transition detection is generally of great interest. Under ambient flow conditions the infrared technique (IR) is a well-established image-based method to carry this out. In high Reynolds number tests which are conducted at low temperatures or at high pressures the IR technique does not work very well whereas the Temperature-Sensitive Paint (TSP) technique is in principle well-suited to operate under these conditions. However, for the use of TSP at high pressures the feasibility has to be checked and some technical aspects have to be solved. Boundary layer transition detection by means of TSP generally requires an artificial temperature step to enhance the temperature difference between the laminar and turbulent boundary layers. This paper presents combination of TSP and Carbon Nanotubes (CNT) to generate a well-defined temperature step to detect laminar-turbulent transition. After pre-testing CNT and TSP in the laboratory, a wind tunnel experiment was conducted in the high pressure wind tunnel in Gottingen (DNW-HDG) for laminar-turbulent transition detection on a two-dimensional model for Re ≤ 8 Mio.