Christian C. Wolf

Experimental Investigation of Dynamic Stall on a Pitching Rotor Blade Tip

Measurements on a pitching finite rotor blade tip are performed. Sectional forces obtained from surface pressure measurements show a significant difference in maximum loads and the extent of hystereses between a section near the parabolic blade tip and two sections further inboard. Different characteristics appear also in the flow topology over the suction side of the airfoil investigated by means of Particle Image Velocimetry (PIV) at these locations.

Rotating Blade Stall Maps Measured by Differential Infrared Thermography

RAFFEL and Merz [1] demonstrated the feasibility of unsteady boundary-layer transition detection by differential infrared thermography (DIT), a method whereby the difference of two sequentially recorded infrared images is analyzed to extract the point of greatest difference between the flows, which, for attached flows, is equivalent to the transition position. For flows with dynamic flow separation, the strongest feature in the infrared difference images is no longer the transition, but the separation of the flow. Gardner and Richter [2] showed that the standard deviation of pressure measurements at the transition point has a small peak, but that separated flow results in an even larger peak. Similarly, the infrared images can be analyzed to extract the presence and extent of separated flowover an airfoil. DIT offers a great advantage over using pressure sensors as indicators for separated flow because DIT does not require any special equipment to be attached to or built into the model observed. The major aim of the investigations presented here is the generation of “stall maps” in which stalled areas in the rotor plane are geometrically described. Currently, these maps are often produced through the investigation of pressure sensor data from sensors integrated into the rotor blades [3]. Alternatives are the use of tuft visualization [4] or pressure-sensitive paint [5]. However, these techniques require installations or coatings on the rotor blades, which can be avoided by the technique described in the following

Detection of Unsteady Boundary Layer Transition Using Three Experimental Methods

Transition measurements were simultaneously performed using hot lm anemometry, pressure analysis, and differential infrared thermography (DIT) in a pitching airfoil wind tunnel test. The principles of the three detection methods will be described and results will be presented for the unsteady transition locations detected on the DSA-9A rotor blade airfoil at M = 0.3 and Re = 1.8 x 10^6. Test cases with periodic ramping motion will be compared to cases with sinusoidal oscillations. The influence of surface heating (necessary for the DIT measurements) will be analyzed and the effects of varying pitching amplitudes and frequencies on the transition detection will be discussed. Furthermore, different post-processing strategies of the DIT infrared images will be investigated and compared to the results obtained with hot film anemometry and pressure analysis.

Tip-Vortex Dynamics of a Pitching Rotor Blade-Tip Model

The tip vortex of a finite rotor blade-tip model undergoing pitch oscillations is investigated by means of a wind-tunnel study. The motion and flow parameters as well as the model geometry were adapted to rotorcraft applications, including test cases with fully attached flow, light dynamic stall, and deep dynamic stall. High-speed particle image velocimetry was used to study the spatiotemporal behavior of the tip wake at different streamwise positions. Combined with surface pressure measurements, the data establish a connection between the sectional lift of the model and the vortex parameters (for example, swirl velocity and circulation), including dynamic phenomena like hysteresis and vortex breakdown during stalled flow conditions. For pitch oscillations below static stall limit, only small hysteresis effects were observed, and the vortex structure is similar to static reference measurements. Beyond static stall, the tip vortex retains its general structure in a time-averaged or phase-averaged frame, bu...

Starting Process of a Pulsed Jet as Seen by Schlieren Measurements

Pulsed jets are amongst the most common measures of active flow control. For example applied to airfoils, stall effects can be dampened or even suppressed, and, hence, the aerodynamic performance can be improved. Although the implementation is more complex, pulsed jets often prove to be a superior alternative to their continuous counterparts. The current jets were designed with respect to the dynamic stall behavior of helicopter rotor blades. Dynamic stall is a well-known limit for forward flight speeds or maneuver loads that can be addressed using flow control. The current work investigates the starting process of a pulsed jet by means of the Background Oriented Schlieren Technique (BOS), using both ‘standard’ and ‘laser-speckle’ background patterns. In agreement with PIV measurements, the results reveal that the initial structure of the start-up process consists of a vortex ring with increasing diameter but constant convection velocity. The vortex is followed by a turbulent wake and a steady but unadapted supersonic jet with corresponding cell structure. The significance of the results is increased by a volumetric reconstruction of the density data, offering further details of the spatial structures. The experimental setup of BOS is not only more convenient in comparison to a standard PIV setup, but also reveals very different phenomena in a single exposure, without additional considerations regarding the follow-up behavior of tracer particles, double-frame time shift, etc.

Spanwise Differences in Static and Dynamic Stall on a Pitching Rotor Blade Tip Model

An experimental investigation of static and dynamic stall on a rotor blade tip model with a parabolic tip geometry and aspect ratio 6.2 at a chord Reynolds number of 900,000 and a Mach number of 0.16 is presented. The resulting flow is analyzed based on unsteady surface pressure measurements and quantitative flow visualizations by high-speed particle image velocimetry. The flow separation is found to be delayed near the parabolic blade tip for static angles of attack as well as for sinusoidal angle of attack motions. The maximum effective angle of attack prior to stall is shifted to approximately two-thirds of the span outboard from the root because of a positive twist of the model with an increasing geometric angle of attack towards the tip. The stall onset is observed near the section with the maximum effective angle of attack, with a subsequent spanwise spreading of the flow separation. Different stages of flow separation for static angles of attack are identified one of them with the occurrence of two stall cells. During dynamic stall, the leading edge vortex formation starts near the maximum effective angle of attack and the pitching moment peak resulting from the passage of the dynamic stall vortex is higher at this section. Further inboard the maximum aerodynamic loads are of comparable magnitude whereas the outboard section shows reduced peaks due to the influence of the wing tip vortex.

Towards Density Reconstruction of Helicopter Blade Tip Vortices from High-Speed Background-Oriented Schlieren Data

A method for the reconstruction of 3D density distributions of blade tip vortices is presented. The reconstruction is performed using a filtered back-projection on 2D density gradient data from a high-speed LS-BOS (laser-speckle-illuminated background-oriented Schlieren) setup, captured during the STAR (smart-twisting active rotor) project. Density profiles of blade tip vortices are presented, as well as comparisons of the size and development of the vortices. In addition the comparison to a PIV (particle image velocimetry) measurement is shown.

Aerodynamic Study on Efficiency Improvement of a Wing Embedded Lifting Fan Remaining Open in Cruise Flight

An aerodynamic study on efficiency improvement of a wing embedded lifting fan with a diameter of 0.36 m remaining open and not closed by doors in cruise flight has been conducted. Two of these fans are used to enable an unconventional unmanned transport aircraft of 16 kg maximum take-off mass and almost 2 m wing span to take-off and land vertically. A preliminary study showed that a step for flow deflection on the lower side in front of the lifting fan duct opening increases the lift-to-drag ratio while even a large change of the duct inlet lip curvature radius has only a negligible impact on cruise flight lift-to-drag ratio. For validation and quantification of these effects, an experimental wind tunnel investigation was completed in the German Aerospace Center (DLR) Seitenwindversuchsanlage Gottingen (SWG) using surface pressure, force, and stereo particle image velocimetry measurement techniques. Wind tunnel data showed an increase of the lift-to-drag ratio of 20% of the overall model by the step for all positive angle of attack. Only minor improvement could be achieved due to a larger inlet lip curvature. The wind tunnel measurement data were used for validation of numerical fluid simulation parameters. Using these validated simulation settings, the step height and inclination was optimized for 2°, 4° and 6° angle of attack, the original model size and model sizes scaled by 1:5 and 5:1 to investigate and quantify the influence of a change of the Reynolds number. At a 4° angle of attack of the original size, the lift-to-drag ratio was increased from 4.7 without a step to 6.6 with a step, compared to a lift-to-drag ratio of 20.75 with smoothly closed lifting-fan-ducts.