Christoph B. Merz

Differential Infrared Thermography for Unsteady Boundary-Layer Transition Measurements

Unsteady laminar/turbulent boundary-layer transition, dynamic flow separation, and reattachment are of special interest in recent propeller and helicopter aerodynamics, since modern rotor blade design methodologies increasingly take dynamic characteristics into account. Sophisticated experimental techniques for the determination of the skin-friction distribution (thereby also detecting transition) have to be developed in order to improve and validate numerical prediction and design methods. The complex hot-film anemometry has successfully been used in order to investigate the unsteady boundary-layer transition on pitching airfoils and large-scale helicopter rotor models in forward-flight configurations.

Blade-Tip Vortex Detection in Maneuvering Flight Using the Background-Oriented Schlieren Technique

The background-oriented schlieren technique was used to visualize the blade-tip vortices of a Eurocopter AS532UL Cougar helicopter in maneuvering flight. The test program covered a large part of the flight envelope, including maneuvers such as hover flight, fast forward flight, flare maneuvers, and high-speed turns. For selected flight conditions, the aerodynamic results are presented here. It is shown that, with the reference-free background-oriented schlieren method, the detection of vortex filaments up to vortex ages of ψv=540  deg is possible. The visualization of the vortex system is used to detect aerodynamic phenomena such as blade–vortex interactions, vortex–airframe interactions, and the occurrence of smooth sinuous disturbances. A detailed description of the applied reference-free background-oriented schlieren setup is given, and the suitability of different natural backgrounds for the background-oriented schlieren method is analyzed.

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

Dynamic Stall Simulations on a Pitching Finite Wing

The aerodynamic effects of dynamic stall on pitching wings and airfoils differ significantly from the static case, as soon as the static stall angle is exceeded. The large drag and pitching moment peaks associated with this phenomenon make it impossible to operate helicopters in flight conditions which trigger dynamic stall over large areas of the rotor blade. To simplify the problem, most of the research on dynamic stall focuses on two-dimensional pitching airfoils. This has lead to an understanding of the process and mechanism of dynamic stall, but has a limited insight into the three-dimensional nature of dynamic stall. There have been a few experimental and numerical studies on full helicopter configurations. These revealed the very complex flow field around the rotor blades, but the combination of downwash, blade-wakevortex interactions and the elasticity of the rotor blades leads to a limited comparability. An approach between these configurations is to use stiff pitching finite wings with defined boundary conditions. In these configurations the blade tip vortex reduces and delays the occurrence of dynamic stall in the surrounding region leading to a strongly three-dimensional flow, see e.g. Le Pape et al. and Lorber.

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...

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.