This study presents a stereophotogrammetry approach to achieve full-field displacement measurements of helicopter rotor blades. The method is demonstrated in the wind tunnel test of a 2 m diameter rotor, conducted at the
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Aeroacoustic Wind Tunnel of China Aerodynamics Research and Development Center (CARDC). By arranging the retroreflective targets on the special hat installed directly above the rotor hub, the dynamic motion of the rotor shaft was tracked accurately, and a unified coordinate system was established on the rotor. Therefore, three-dimensional coordinates of instantaneously measured targets attached on the blade could be transformed to the unified rotor coordinate system, thereby providing a basis for consistently calculating the blade displacements at different test conditions. Moreover, location deviations of the blade caused by the vibration of the measuring system or the rotor due to freestream and rotor rotation were also effectively corrected through coordinate transformation. Comparisons of experimental and simulation results for a range of hover and forward flight conditions show good magnitude and trend agreements.
When the aircraft opens the bay door to let the landing gear either drop or retract, the incoming flow will result in a significant amount of coupling noise from the bay and the landing gear. Here, an experimental study was reported to characterise the acoustic performance and flow field at low subsonic speeds. Also, we examined a passive control method leading-edge chevron spoiler to suppress the noise. The experiment was performed in a low-speed aeroacoustic wind, the bay was simplified as a rectangular cavity and the spoiler was mounted to the leading edge. Both acoustic and aerodynamic measurements were performed through two microphone arrays, pressure transducers and particle image velocimetry. It was found that installation of the landing gear model can attenuate cavity oscillation noise to some extent by disturbing the shear layer of the cavity leading edge. Moreover, acoustic measurement confirmed the noise control when the spoiler was used. In addition, a parametric study on the effects of chevron topology was performed, and an optimised value was found for each parameter. From the aerodynamic measurement, the noise reduction was explained from the perspective of fluid dynamics. It was observed that installation of the chevron can raise the leading-edge shear layer and break up the large-scale vortices, thereby controlling the Rossiter mode noise and the landing gear model noise at certain frequencies.
In order to improve the aerodynamic performance of the wing at post-stall conditions, the experimental comparative investigations on the flow separation control over an ONERA 212 airfoil using steady and unsteady plasma actuators are carried out at Reynolds number of 3.1 × 105. The duty-cycle ratio is fixed at 80% and the non-dimensional unsteady frequency F+ is varied from 0.04 to 1. The lift coefficients are increased by 39.6% and 66% respectively after steady and unsteady operations (F+=0.08) at an angle of attack of 18{degree sign}, which indicates that the unsteady actuation is more efficient than steady operation. Meanwhile, the study provides new insight into understanding the post-stall separation flow controlling mechanism. First, different from the general view that the injection of momentum is the controlling mechanism of steady operation, flow control using the steady actuation experiences four stages, namely, flow separation, promoting the instability of the separated shear layer to produce large-scale spanwise vortices, flow re-attachment, and the continuous generation of small-scale vortices in the separated shear layer. Secondly, flow control with the unsteady operation consists of several quasi-periodic flow processes. Each quasi-cycle is composed of three stages, namely, flow separation, promoting the separation of shear layer instability to produce large-scale spanwise vortices, and flow re-attachment. The off-time of the plasma actuator plays an important role in realizing the control effect of the unsteady actuation, and an effective strategy to promote the control effect of the unsteady operation is proposed based on the propagation time of the induced spanwise vortex.
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