Experimental Investigation of Dynamic Stall on a Pitching Rotor Blade ...

Experimental Investigation of Dynamic Stall on a Pitching Rotor Blade Tip C.B. Merz, C.C. Wolf, K. Richter, K. Kaufmann, and M. Raffel

Abstract 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.

1 Introduction The term ‘dynamic stall’ refers to the unsteady separation on lifting surfaces encountered due to a dynamically changing effective angle of attack beyond the static stall angle. Dynamic stall generally results in an increase of lift well beyond the maximum static lift but also in large excursions in the drag and pitching moments. On helicopter main rotors, dynamic stall often occurs on the retreating blade during highly loaded maneuvers or fast forward flight where the low inflow velocity leads to high effective angles of attack. Because of the prohibitive loads on the pitch links induced by the sharp peak in the pitching moment and increased vibrations of the structure, the flight envelope of modern helicopters is limited such that dynamic stall does not occur on large parts of the rotor blade. A well suited model for wind tunnel tests of dynamic stall at scales relevant to helicopter main rotors is a 2D airfoil pitching at a comparable reduced frequency, as observed in [7]. To investigate the mutual influence of the blade tip vortex and the dynamic stall, the model has been extended to a pitching finite wing. Comprehensive studies including a range of Mach and Reynolds numbers, amplitudes, frequencies and sweep angles have been C.B. Merz · C.C. Wolf · K. Richter · K. Kaufmann · M. Raffel German Aerospace Center (DLR), Institute of Aerodynamics and Flow Technology, Bunsenstraße 10, 37073 Göttingen, Germany, e-mail: [email protected]

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C.B. Merz, C.C. Wolf, K. Richter, K. Kaufmann, and M. Raffel

performed using surface pressure measurements as well as hot film sensors [5] or micro-tufts on the model surface [10]. Further studies have included flow visualization such as the smoke wire technique [1] and dye visualization [11]. A more recent study [4] used phase averaged PIV at three sections at constant streamwise positions, one section of Laser Doppler Velocimetry (LDV) and four sections of surface pressure measurements at constant span to investigate the spanwise propagation of the dynamic separation and reattachment near the blade tip as well as the interaction between the tip vortex and the dynamic stall. All studies reported a significant change in the dynamic stall behavior near the blade tip compared to 2D pitching airfoils. The presence of a tip vortex appears to delay the formation, separation and convection of a dynamic stall vortex. The flow further inboard is frequently described as similar to a 2D dynamic stall case in terms of the formation and propagation of the dynamic stall vortex. In case of an untwisted wing, the dynamic stall vortex starts to form close to the wing root, where the first occurrence of leading edge separation takes place subsequently (cf. [6, 12]). Further studies combining time-resolved quantitative flow field measurements with surface pressure distributions are needed to relate aerodynamic loads to coherent structures in the flow field and to establish additional data bases for the validation of Computational Fluid Dynamics (CFD) calculations. Furthermore, the investigation of a dynamic separation occurring first close to the blade tip and subsequently spreading both inboard and outboard can produce valuable insight to the dynamic stall as it appears on helicopter rotors. The current study is part of the DLR project Stall and Transition on Elastic Rotor Blades (STELAR). The dynamic stall on a pitching finite wing is investigated at several spanwise locations using surface pressure measurements as well as tuft visualization, deformation measurements and high-speed PIV. An overview of the experimental setup and the test matrix is given, and one test case is discussed in detail.

2 Experimental setup A wind tunnel model was designed to investigate the dynamic stall near a free blade tip. The model has an aspect ratio of six and a DSA-9A airfoil. The span is 1620 mm and the chord is 270 mm, except for the parabolic blade tip which starts at y = 1404 mm and along which the chord reduces to 90 mm according to the specifications of the SPP8 blade tip without anhedral [9]. The downwash induced by the tip vortex reduces the effective angle of attack. This effect increases when moving closer to the blade tip. The maximum effective angle of attack for an untwisted and unswept wing is therefore at the wing root. Consequently, the stall interferes with the wind tunnel wall, which is undesirable for the current investigation. To move the maximum effective angle of attack and thereby the location of the onset of flow separation to the outboard section of the wing, the model has a positive linear twist of 5.5◦ . To transfer the aerodynamic loads to the shaft of the model, the airfoil thickness of the innermost 100 mm of the span increases from 9% of the chord to 18%

Experimental Investigation of Dynamic Stall on a Pitching Rotor Blade Tip

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chord with constant curvature transitions. The model is equipped with 100 Kulite (XCQ-093 series) unsteady pressure transducers for the measurement of the surface pressures at three chord-wise cuts around the entire airfoil and three span-wise sections on the suction side (see Fig. 1). The surface pressures at the three chord-wise cuts, indicated by the red lines, can be integrated to obtain the sectional forces and moments. The distribution of the pressure transducers was optimized to achieve a low discretization error throughout a pitching cycle including deep dynamic stall. The resulting maximum discretization error was estimated based on unsteady threedimensional Reynolds-averaged Navier-Stokes (RANS) computations to approximately 0.015 in cl , less than 0.003 in cd and less than 0.005 in cm as long as the flow is attached to the wing section. During parts of the pitching cycle where the flow is massively separated, the uncertainty increases to 0.03 in cl , approximately 0.02 in cd and less than 0.02 in cm . The model is made from carbon fiber reinforced plastic to achieve a high stiffness at a low weight. Nevertheless, elastic deformation was observable during the experiments – especially a flapping motion and to a lesser extent also a torsional motion for cases with increased pitching moments. A finite element analysis has been performed to identify the eigenmodes and to find optimal positions for six PCB 352C22 acceleration sensors inside the model. The first four eigenmodes corresponding to blade flapping (first and third mode), lead-lag motion (second mode) and blade torsion (forth mode) can be monitored. An additional flutter analysis revealed no unstable modes.

Fig. 1: Sketch of the wind tunnel model including the locations of the pressure taps on the suction side of the model (•) and the fields of view for the PIV measurements at y = {750, 1100, 1400} mm span. An electric test rig was developed to perform sinusoidal pitching motions as well as angle of attack ramps and higher harmonic oscillations. The system consists of a Parker 190ST6M motor with a 4:1 ASD SPG040.2-M-4-00 gearbox in an on-axis configuration. The motor controller has a built-in cascaded controller for current, rotational speed and position. An external iterative learning control for the position input was added to improve the repeatability of the motion. The resulting maximum standard deviation of the sinusoidal motion for deep dynamic stall cases is approx-

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imately 0.08◦ . The largest cycle-to-cycle variations occur in massively separated flow conditions with strong negative peaks in the pitching moment. The experiments were carried out in the Side Wind Test Facility Göttingen (SWG), a closed-loop, low speed wind tunnel with a rectangular test section of 2.4 m width and 1.6 m height. The maximum free stream velocity is 65 m/s. The free stream velocity throughout this study was set to 55 m/s resulting in a Mach number of approximately 0.16 and a chord-based Reynolds number of about 900 000. A number of different mean angles of attack as well as amplitudes for the sinusoidal motion of the airfoil were tested, ranging from attached flow throughout the entire pitching cycle to deep dynamic stall conditions. The reduced frequency k = π f c/U∞ was varied between 0.025 and 0.1. Additionally, ramps with non-dimensional pitch ˙ rates α ∗ = αc/2U ∞ between 0.01 and 0.06 and quasi-steady motion were performed from fully attached flow to complete flow separation. The flow field in several planes on the suction side of the airfoil was measured by means of stereoscopic PIV. Two PCO.dimax cameras and a Litron LDY304 laser were used at an acquisition rate of 1 kHz per image pair. The camera resolution was reduced to 1680x1472 pixels to achieve the desired acquisition frequency. The presented results were obtained using a multi-grid evaluation with a final interrogation window size of 32x32 pixels and an overlap of 50% resulting in a velocity vector spacing of 3.6 mm. Assuming a precision of 0.1 pixel for the determined DEHS particle displacements, the uncertainty in the velocity is approximately 3% of U∞ . The same hardware setup, with a slightly different spatial resolution was used to obtain the flow field in the wake including the tip vortex at three planes perpendicular to the inflow. Additionally, the flap and torsional motion of the model was measured between y = 1122 mm and y = 1363 mm span by means of Stereo Pattern Recognition (SPR), a marker based photogrammetry method for deformation measurements. Three HardSoft IL-106X Illuminator high power LEDs and two additional PCO.dimax cameras were used for illumination and image acquisition of the deformation measurements. The SPR measurements were synchronized to the PIV acquisitions. Surface pressures, model accelerations and the angle of attack at the wing root were sampled at 20 kHz and related to the PIV recordings by means of a feedback signal and a fast photodiode. The surface pressures at y = 800 mm, y = 1100 mm and y = 1400 mm span were integrated using the in-house cp2cl plusm tool [3] to obtain the sectional lift, drag and pitching moment.

3 Results and Discussion First, results of the model dynamics are presented including deflections and eigenmodes. Next, the integral loads at three spanwise locations are compared, followed by a discussion of the onset, propagation and termination of dynamic stall using tuft visualizations and PIV.


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The maximum deflection at y = 1363 mm span, calculated from the SPR results for a dynamic stall test case of αroot (t) = 9◦ + 6◦ sin(2π f t) with f = 3.2 Hz (k = 0.05), is approximately 27.2 mm and the minimum about 8.2 mm compared to wind-off conditions. In comparison, the maximum deflection is around 21.0 mm and the minimum about 6.2 mm at y = 1122 mm span. The high negative pitching moment peaks also lead to a measurable torsional deformation of the model. The maximum local change in the geometric angle of attack between y = 1122 mm and y = 1363 mm span is less than 0.3◦ compared to the measured angle of attack at the blade root. For the same test case, the first bending eigenmode of the system lies a little below 20 Hz and is detectable in both SPR measurements and model accelerations. It was not possible to identify a frequency for the second eigenmode corresponding to a lead-lag motion of the wing. However, the second flap eigenmode at 83 Hz as well as the first torsional eigenmode at 106 Hz could be detected from a spectral analysis of the model accelerations. As expected from the flutter analysis, all modes were damped. If the mean angle of attack was reduced such that the flow was attached throughout the entire pitching cycle, no eigenmodes were detectable in the SPR recordings and no clear peaks could be detected with a spectral analysis of the accelerations. Figure 2 depicts the phase averaged coefficients of lift, pitching moment and drag for the same dynamic stall test case. The angle of attack is corrected for the local twist at each spanwise location. A significant hysteresis is present in all coefficients and at all spanwise locations indicating the presence of dynamic stall. However, there are differences in the characteristics of the integral loads between the three spanwise positions. Most notable is the difference between the two inboard sections and the position near the parabolic blade tip. While the two inboard sections show large hystereses in lift, indicating deep dynamic stall conditions, the hysteresis in the outboard section as well as the reduced peaks in drag and pitching moment point to a less pronounced stall in that area. This agrees well with the findings in [4] for the section at 0.6 chord lengths from the blade tip compared to the mid-span section. The lift at the outmost position is significantly lower throughout the upstroke. A

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then the lift remains higher during the stalled stage compared to the other sections. The lift loops are similar for the two inner sections except for the characteristics near the maximum lift. At y = 800 mm, there are two distinct peaks, the second is interrupted by the reversal of the pitching motion leading to an abrupt loss of lift. At y = 1100 mm, only one peak can be observed and lift reduces already before the turning point of the pitch oscillation. During the upstroke, before the start of flow separation, the pitching moment is nearly zero for all positions. As the dynamic stall vortex progresses towards the trailing edge of the airfoil, there is a strong negative peak in pitching moment. A gradual recovery occurs during the entire downstroke, with a small overshoot to positive pitching moments. The peak negative pitching moment is greatest for the section at y = 1100 mm and is significantly smaller at the most outboard section. The pressure drag increases at all sections with increasing angle of attack. With the onset of massive flow separation, a spike appears in all drag curves, followed by a hysteresis during the downstroke before the flow is again fully attached to the wing surface. As with the pitching moment, the maximum load is at y = 1100 mm span, whereas the drag at the section further inboard is slightly lower and the drag at the outboard section is significantly reduced. The hysteresis is also most pronounced for the section at y = 1100 mm. A qualitative investigation of the stalling process for the same airfoil motion is shown in Fig. 3, with all images taken during the same pitching cycle. The tufts on the suction side of the wing indicate regions of flow reversal and stall. The separation starts at the trailing edge over much of the wing span as visible in Fig. 3(a). Then, while still on the upstroke, the separated region moves upstream and reaches the leading edge around the position indicated with a red arrow in Fig. 3(b). At the same time, an outboard motion of the flow can be observed near the trailing edge on the blade tip as well as the roll-up of the tip vortex. After the reversal of the pitching motion, the flow separates over most of the wing as depicted in Fig. 3(c). Only over part of the parabolic blade tip the flow appears to be attached. At the time instance of Fig. 3(d), the flow is mostly attached to the model surface again. A last area of separated flow is visible slightly further outboard from the position of the first complete separation. The phase averaged pressure distributions at αroot = 14.4◦ ↑, corresponding to the snapshot in Fig. 3(b), are depicted in Fig. 4. The onset of stall for the section at y = 1100 mm is also evident from the pressure distribution. The suction peak is already significantly reduced and the standard deviations of the pressure signals on the suction side are strongly increased. Further inboard, the suction peak is near its maximum of c p = −10.00, yet the increased standard deviations of the pressures near the trailing edge indicate a beginning trailing edge separation. This is in contrast to the results obtained with an untwisted wing (see e.g. [6]), where the breakdown of the suction peak and the onset of stall occurs first on the most inboard section of the wing. On the most outboard section, the suction peak is still increasing slightly to a maximum of c p = −7.92 shortly after the snapshot shown here, and there are no increased standard deviations of the pressure signals near the trailing edge. An analysis of the course of dynamic stall at three different spanwise locations is presented in Fig. 5. The top row depicts the sectional lift coefficients


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Fig. 3: Visualization of the flow separation using tufts. All images are recorded during the same pitching cycle.

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Fig. 4: Phase averaged pressure distributions with standard deviations at three spanwise positions at αroot = 14.4◦ ↑. at y = {800, 1100, 1400} mm. Besides the phase averaged data, the instantaneous lift values corresponding to the time instances of the PIV recordings for one cycle are shown. The subsequent rows show snapshots of the PIV recordings at y = {750, 1100, 1400} mm span. Following the classifications given in [2], four PIV snapshots are selected based on the section at y = 1100 mm, and the corresponding lift values are highlighted in red. The first image is selected as the lift slope increases (αroot = 14.4◦ ↑), and the second image is chosen as the maximum negative pitching moment is reached (αroot = 15.0◦ ↑). The series is concluded with a snapshot during full stall (αroot = 14.1◦ ↓) and a final image as the flow reattaches from the leading edge to the trailing edge (αroot = 7.6◦ ↓). Note that the location of the sectional lift and the PIV recordings at the two outboard sections matches exactly, whereas in the left column, the sectional lift is determined 50 mm further outboard of the presented PIV snapshots because of limited optical access at that spanwise position. To

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identify coherent structures in the flow field above the suction side of the wing, the swirling strength criterion according to [13] has been applied on the two in-plane components of the velocity. The swirling strength λci is multiplied by the sign of the out-of-plane component of vorticity, which is a common procedure to obtain information about the direction of the swirling motion. The new quantity is sometimes referred to as ‘signed swirling strength’ and is given the symbol λci± here. The first row of images at t/T = 0.43 depicts the situation shortly before the dynamic lift overshoot. The boundary layer on the most inboard section appears unstable – especially near the trailing edge – which is visible as a thickening of the area with slightly increased swirling strength compared to the free stream. The flow at the middle section has already started to separate downstream of the suction peak. In contrast to the more inboard locations, the section at y = 1400 mm shows no sign of boundary layer instability or separation. At t/T = 0.47, before the turning point of the pitching motion, the flow is separated from the leading edge in both inboard sections, as can be seen in the velocity vectors pointing away from the airfoil surface right after the leading edge. A shear layer indicated by clockwise rotating vortices (λci± < 0) is visible. The shear layer is further away from the airfoil surface at y = 1100 mm, resulting in a slightly reduced lift at that position. In addition, counter-clockwise rotating structures have developed at this section between the shear layer and the airfoil surface. As described in [8], these counter-rotating vortices develop as a consequence of the interaction of the stall vortex and the reversed flow on the airfoil surface. At the same time, the pitching moment reaches a sharp negative peak at y = 1100 mm, caused by the dynamic stall vortex moving towards the trailing edge. In the case of fully separated flow at the inboard sections (t/T = 0.59), the shear layer is at a greater distance from the airfoil surface. Pressure distributions (not shown) reveal that the suction peak has completely collapsed at this time. The short re-establishment of weaker suction peaks and the subsequent shedding of secondary vortices from the leading edge leads to short interruptions in the loss in lift, as apparent in the instantaneous lift values, especially at y = 1100 mm. Note that there remains a reduced suction peak at the most outboard section. There is also no clear lift overshoot, a more gradual decrease in lift in the post-stall phase and a reduced peak in negative pitching moment at this location. This indicates that no strong dynamic stall vortex is shed from the leading edge, which can be confirmed by the images of the swirling strength. At t/T = 0.79, the flow reattaches at the section at y = 1100 mm. As described in [2], the flow reattachment occurs from the leading edge to the trailing edge. Remaining vortices above the surface are convected with the free stream. A new suction peak has developed, and the lift, drag and pitching moment have recovered at all spanwise sections. While the results at y = 1100 mm show a lot of the characteristics of a deep dynamic stall cycle on a pitching 2D airfoil as discussed in [8], there is a significant difference for the most outboard section. The described flow topology could be interpreted as a light dynamic stall due to the reduced effective angle of attack near the blade tip.


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4 Conclusions A wind tunnel model with a positive linear twist of 5.5◦ , an aspect ratio of six and a parabolic wing tip was designed for the investigation of dynamic stall near a free blade tip. Experiments with various pitching motions were performed. A test case of an angle of attack motion of αroot (t) = 9◦ + 6◦ sin(2π f t) with f = 3.2 Hz (k = 0.05) was discussed in detail. Tuft visualization as well as pressure distributions at three sections with constant span revealed that the onset of complete flow separation occurs approximately two chord lengths inboard of the tip. The analysis of the flow field at that location showed similarities to the deep dynamic stall observed on a 2D airfoil for a similar range of non-dimensional parameters. The flow at 0.8c from the tip exhibited a significant influence of the reduced effective angle of attack. The flow separation resembled more of a light stall case with reduced peaks in the load coefficients and weaker coherent structures in the separated flow field.

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