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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

A rotor test facility was developed to investigate dynamic stall under optimized boundary conditions compared to conventional hover chambers. lt features a defined axial inflow, reduction of ground effect and recirculation of the rotor wake and good optical access to apply non-intrusive measuring techniques. The rotor consists of two blades with an aspect ratio of 6.8 and a tip radius of 0.65m. lt was operated at a Mach number of 0.21 and a chord based Reynolds number of 350000, both at 75% radius. The flow and blade deformation were analyzed by means of unsteady blade pressure transducers, particle image velocimetry and tip deflection measurements covering the whole azimuth and different radii. Three measurement approaches to detect flow separation: tufts, differential infrared thermography and surface pressure analysis showed comparable results and flow separation over one half of the cycle. A radiusdependent separation point in time and an altering of the dynamic stall behavior near the tip were observed. The recirculation area remained close to the blade's surface with radial jet flow increasing towards the tip. A maximum increase of the effective angle of attack of 1.4° was observed due to blade deformation. For a fully attached flow case the laminar-turbulent boundary layer transition differed along the radius and moved over 70% of the blade's chord.

Laminar-turbulent boundary-layer transition is investigated on the suction side of Mach-scaled helicopter rotor blades in climb and analyzed in view of the effect of rotational forces. Transition p...

The tip vortex system downstream of a four-bladed instrumented rotor was investigated experimentally through the application of stereoscopic particle image velocimetry (PIV). A dynamic stall test case was facilitated by a high cyclic pitch setting of the swashplate, with additional attached-flow and constant-pitch test cases for comparison reasons. The phase-locked PIV system and a rotation of the swashplate assembly allowed for an acquisition of the tip vortex system over the entire dynamic stall cycle and vortex ages up to at least 235°. The vortex structure and its relation to the blade shear layers were studied by means of both phase-averaged flow fields and the identification of vortex properties such as circulation and swirl velocity distributions. When approaching dynamic stall, a breakdown of the vortex structure started at high vortex ages, accompanied by the entrainment of turbulent structures from the passing blade shear layers into the tip vortices. After the flow over the blade is fully separated and during large parts of the downstroke, the wake of the rotor tips appears as a highly turbulent area with no individual tip vortices traceable, before reestablishing an ordered tip vortex structure shortly before the minimum blade pitch angle.

The blade-tip vortex of a rotating and pitching DSA-9A blade was investigated numerically and experimentally. Numerical computations were performed using DLR, German Aerospace Center’s finite volum...

A pressure-sensitive paint (PSP) system is presented to measure global surface pressures on fast rotating blades. It is dedicated to solve the problem of blurred image data employing the single-shot lifetime method. The efficient blur reduction capability of an optimized double-shutter imaging technique is demonstrated omitting error-prone post-processing or laborious de-rotation setups. The system is applied on Mach-scaled DSA-9A helicopter blades in climb at various collective pitch settings and blade tip Mach and chord Reynolds numbers ( $$M_{\text {tip}}$$ = 0.29–0.57; $$Re_{\text {tip}}$$ = 4.63–9.26 $$\times 10^5$$ ). Temperature effects in the PSP are corrected by a theoretical approximation validated against measured temperatures using temperature-sensitive paint (TSP) on a separate blade. Ensemble-averaged PSP results are comparable to pressure-tap data on the same blade to within 250 Pa. Resulting pressure maps on the blade suction side reveal spatially high resolved flow features such as the leading edge suction peak, footprints of blade-tip vortices and evidence of laminar–turbulent boundary-layer (BL) transition. The findings are validated by a separately conducted BL transition measurement by means of TSP and numerical simulations using a 2D coupled Euler/boundary-layer code. Moreover, the principal ability of the single-shot technique to capture unsteady flow phenomena is stressed revealing three-dimensional pressure fluctuations at stall.

The flight envelope of modern helicopters is limited due to unsteady flow phenomena such as dynamic stall. Most investigations regarding this topic, both experimental and numerical, were done on pitching two-dimensional airfoils. Therefore, it is of great interest to investigate the dynamic stall on rotating blades in order to get a better understanding of the unsteady aerodynamics and to improve future rotor designs. A rotor test facility is being built at the DLR in Gottingen. The test rig is integrated into an existing wind tunnel facility and the surrounding conditions were defined. Critical load cases were derived from numerical flow simulations of the rotor blade, and a finite element analysis including a fatigue analysis was performed. A telemetry system is integrated into the rotor head and unsteady pressure sensors are installed in the rotor blades. The development process with its validation and the design of the rotor head and blades are described in this paper.

Differential infrared thermography (DIT) was investigated and applied for the detection of unsteady boundary layer transition locations on a pitching airfoil and on a rotating blade under cyclic pitch. DIT is based on image intensity differences between two successively recorded infrared images. The images were recorded with a high framing rate infrared camera. A pitching NACA0012 airfoil served as the first test object. The recorded images were used in order to investigate and to further improve evaluation strategies for periodically moving boundary layer transition lines. The measurement results are compared with the results of unsteady CFD simulations based on the DLR-TAU code. DIT was then used for the first time for the optical measurement of unsteady transition locations on helicopter rotor blade models under cyclic pitch and rotation. Image de-rotation for tracking the blade was employed using a rotating mirror to increase exposure time without causing motion blur. The paper describes the challenges that occurred during the recording and evaluation of the data in detail. However, the results were found to be encouraging to further improve the method toward the measurement of unsteady boundary layer transition lines on helicopter rotor models in forward flight.

Numerical investigations of three-dimensional dynamic stall on a two-bladed Mach-scaled rotor (R = 0.65m, Ma75 = 0.21, Re75 = 3.5 × 105) with 1/rev cyclic pitch control are presented and compared to experimental surface pressure and PIV data. In addition to URANS simulations using the finite-volume flow solvers FLOWer and TAU, a delayed detached-eddy simulation (DDES) with Menter SST as underlying RANS model is carried out with FLOWer. Facing dynamic stall and flow separation, the DDES reproduces high-frequency load fluctuations, cycle-to-cycle variations and small-scale vortical structures as seen in the experiment, which is not the case with URANS. However, common hybrid RANS-LES issues – grid-induced separation and the grey area problem – play a role in this DDES and influence loads severely. FLOWer SST simulations yield load peaks of the same magnitude as individual, non-phase-averaged measurements. With TAU SST the dynamic stall event is delayed and weakened compared to FLOWer SST and experimental results. FLOWer and TAU results using the SA turbulence model are fairly comparable but in bad agreement with the experiment at the outboard station at r/R = 0.77, where they exhibit no dynamic stall at all.