Holger Mai

Nonlinear effects in transonic flutter with emphasis on manifestations of limit cycle oscillations

This paper presents flutter and forced oscillation experiments in a transonic wind tunnel. For an aeroelastic supercritical 2-D airfoil configuration we studied typical transonic phenomena in as pure a form as possible. Various manifestations of small-amplitude limit cycle oscillations were observed for different flow conditions as well as coexisting limit cycles. We demonstrated how very small control forces were sufficient to excite or suppress flutter oscillations. Limit cycle oscillations occurred under free and forced turbulent boundary layer transition in a perforated wall test-section. Flutter calculations based on experimental aerodynamic forces yield stability limits which show good agreement with directly measured experimental flutter values. The results indicate that flow separation at the trailing edge, and the interactions between the shock and the marginal region of separated flow beneath it, may be responsible for limiting the amplitude of the observed limit cycle oscillations.

Experimental investigation of dynamic stall performance for the EDI-M109 and EDI-M112 airfoils

An experimental investigation of the dynamic performance of two new rotor blade airfoils was undertaken in a transonic wind tunnel. The EDI-M109 and EDI-M112 airfoils were tested at 0.3<M<0.5 for pitching motions with amplitude 0.5° to 8° and frequencies 3.3Hz<f<45Hz. The results show that both new airfoils have acceptable dynamic stall performance, and the effect of frequency, amplitude, and higher order pitching motion on these results is described. The pitching moment peak size was found to have an approximately linear correlation to the normalised mean angular velocity, and thus test cases with the same maximum angle of attack and oscillation frequency had similar dynamic stall qualities. The correlation between low aerodynamic damping for high frequency, low amplitude pitching motion and poor dynamic stall performance is shown to be low. The dynamic stall response of the EDI-M112 airfoil is shown to be better for M=0.3 and M=0.4, and the response of the EDI-M109 airfoil is better at M=0.5. The dynamic performance of the airfoils is compared to the OA209.

Application of Pressure-Sensitive Paint for Determination of Dynamic Surface Pressures on a Rotating 65° Delta Wing and an Oscillating 2D profile in Transonic Flow

Visualization and measurements of aerodynamic effects on a delta-wing model and a 2D-wing-profile model were conducted using an optical pressure measurement system, based on the pressure-sensitive paint (PSP) technique. The PSP technique can be used to obtain absolute pressure measurements on the surface of a model and in addition to evaluate quantitative aerodynamic flow phenomena by using scientific grade cameras and image processing techniques. The PSP technique has been used here for investigations of periodic and unsteady flows: first, a 65deg delta wing was tested in the transonic wind tunnel DNW-TWG in Gottingen. A specially designed roll apparatus enabled roll rates up to 10 Hz. The experiments were carried out at angles-of-attack up to alpha = 17deg at Ma = 0.8. Since the rotation of the delta wing is a periodic motion, the phase-locked unsteady PSP technique can be applied. In a second wind tunnel campaign in the DNW-TWG in collaboration with the DLR Institute of Aeroelasticity, a 2D-wing-profile model, which is pitch oscillating at up to 30 Hz, was investigated. The experiments were performed at angles-of-attack alpha = 1.12deg plusmn 0.6deg at Ma = 0.72. For these experiments pressure measurements were carried out in one wind tunnel entry by means of both phase-locked unsteady as well as unsteady PSP techniques.

Experimental Investigation of Unsteady Transition on a Pitching Rotor Blade Airfoil

The unsteady flow around the pitching helicopter main rotor blade airfoil EDI-M109 was experimentally investigated at conditions similar to those existing on a retreating rotor blade in forward flight. High speed pressure measurements and hot film anemometry were used to investigate the unsteady transition characteristics of the airfoil. Results are presented for dynamic test points with attached flow, light dynamic stall and deep dynamic stall at M = 0.3 and Re = 1.8 x 10^6. The results include the discussion of the periodicity of the hot film signals for different flow states. The transition process of the pitching airfoil is analysed and the significance of the intermittent region is described. A time delay between the transition and the model motion is discussed and a linear relationship between the transition position and the time is observed. The influences of the pitching amplitude on the transition characteristics are discussed and the flow separation initiating dynamic stall is analysed.

Experimental Investigation of Air Jets for the Control of Compressible Dynamic Stall

The experimental investigation of constant blowing air jets as fluidic control devices for helicopter dynamic stall control is described. A carbon fiber airfoil of constant OA209 cross section was fitted with a pneumatic system to deliver dry compressed air as jets for flow control at total pressures of up to 10 bar. The experiment used porthole jets of radius 1% chord, positioned at 10% chord and with spacing 6.7% chord. The positive dynamic stall control effects were demonstrated at Mach 0.3, 0.4, and 0.5 for deep dynamic stall test cases with the best test cases reducing the pitching moment peak after the main stall by 83% while increasing the mean lift over one pitching cycle by 30%. The conclusions from the experiments are supported by three-dimensional unsteady Reynolds-averaged Navier–Stokes (URANS) computations of the pitching airfoil with flow control using the DLR-TAU code.

Application of Pressure-Sensitive Paint for Determination of Dynamic Surface Pressures on a 30 Hz Oscillating 2D Profile in Transonic Flow

Visualization and measurements of aerodynamic effects on a 2D-wing-profile model were conducted using an optical pressure measurement system based on the pressure-sensitive paint (PSP) technique. The PSP technique can be used to obtain absolute pressure distributions on the surface of a model and in addition to evaluate quantitative aerodynamic flow phenomena e.g., shock location, shock-shock interaction, and shock boundary layer interaction, by using scientific grade cameras and image processing techniques. The PSP technique has been used here for investigations of periodic and unsteady flows. In a wind tunnel campaign in the DNW-TWG, a 2D-wing-profile model, which is pitch oscillating at up to 30 Hz, was investigated. The experiment presented here was performed at angles-of-attack α = 1.12° ± 0.6° at Ma = 0.72. With this work the area of application of PSP to dynamic systems where oscillating pressure changes of the order of 1000 Pa have to be measured at rates of up to 100 Hz is demonstrated.

Experimental investigation of air jets to control shock-induced dynamic stall

This paper shows experimental results for dynamic stall control on a dynamically pitching OA209 airfoil at Mach 0.5, with Reynolds numbers 1.9 × 106 and 0.85 × 106 . The control was by supersonic constant blowing on the suction side of the airfoil. Dry compressed air was blown normal to the airfoil chord, from portholes at 10% chord, with diameter 1% chord. At both Reynolds numbers, the OA209 without blowing experienced shock-induced stall with a hysteresis in lift and pitching moment around the static values, rather than the overshoot in forces typically associated with a dynamic stall vortex. The forces and the stall control were primarily functions of the maximum angle of attack, with full control of stall possible for maximum angles of attack of 14◦ and less. The higher Reynolds number required relatively more blowing (higher Cq , Cμ ) to control the dynamic stall. Drag was reduced for separated flow, but the energy required in compressed air to achieve this was more than the savings in drag, and no cases were found in which flow control resulted in a reduction in total power used. Increasing the jet spacing resulted in equivalent flow control with less air use. Jets spaced at 20% chord and mass flux ratio Cq = 0.004 (momentum ratio Cμ = 0.016) resulted in a reduction of the pitching moment peak by 60%. The flow control with air jets was uncritical regarding the aerodynamic damping.

Investigation of the unsteady flow development over a pitching airfoil by means of TR-PIV

The flow over an OA209 airfoil subjected to a sinusoidal pitching motion under dynamic stall conditions is investigated experimentally by means of time resolved particle image velocimetry (TR-PIV) and surface pressure measurements. Dynamic stall is distinguished by the formation and convection of large scale coherent structures and a delay in massive flow separation. A vortex detection scheme based on an identification function derived directly from the velocity fields is adopted to identify vortex cores. The combination of global time resolved imaging and an automated vortex identification algorithm allows for the investigation of the spatial and temporal evolution of vortical structures within a single oscillation. Furthermore, the mechanisms associated with the dynamic stall delay are considered.

Helicopter Aerodynamics with Emphasis Placed on Dynamic Stall

Dynamic Stall is a flow phenomenon which occurs on helicopter rotor blades during forward flight mainly on the retreating side of the rotor disc. This phenomenon limits the speed of the helicopter and its manoeuvrability. Strong excursions in drag and pitching moment are typical unfavourable characteristics of the Dynamic Stall process. However compared to the static polar the lift is considerably increased. Looking more into the flow details it is obvious that a strong concentrated vortex, the Dynamic Stall Vortex, is created during the up-stroke motion of the rotor blade starting very close to the blade leading edge. This vortex is growing very fast, is set into motion along the blade upper surface until it lifts off the surface to be shed into the wake. The process of vortex lift off from the surface leads to the excursions in forces and moment mentioned above. The Dynamic Stall phenomenon does also occur on blades of stall regulated wind turbines under yawing conditions as well as during gust loads. Time scales occurring during this process are comparable on both helicopter and wind turbine blades. In the present paper the different aspects of unsteady flows during the Dynamic Stall process are discussed in some detail. Some possibilities are also pointed out to favourably influence dynamic stall by either static or dynamic flow control devices.