The practical aspects of an advanced schlieren technique, which has been presented by Meier (1999) and Richard et al (2000) and in a similar form by Dalziel et al (2000), are described in this paper. The application of the technique is demonstrated by three experimental investigations on compressible vortices. These vortices play a major role in the blade vortex interaction (BVI) phenomenon, which is responsible for the typical impulsive noise of helicopters. Two experiments were performed in order to investigate the details of the vortex formation from the blade tips of two different helicopters in flight: a Eurocopter BK117 and a large US utility helicopter. In addition to this, simultaneous measurements of velocity and density fields were conducted in a transonic wind tunnel in order to characterize the structure of compressible vortices.The background oriented schlieren technique has the potential of complementing other optical techniques such as shadowgraphy or focusing schlieren methods and yields additional quantitative information. Furthermore, in the case of helicopter aerodynamics, this technique allows the effect of Reynolds number on vortex development from blade tips to be studied in full-scale flight tests more easily than through the use of laser-based techniques.

The paper presents an experimental investigationof the effect of the trailing edge vortex shedding on the steady and unsteady trailing blade pressure distribution of a turbine blade at high subsonic Mach number M2,is=0.79 and high Reynolds number RE=2.8×106. The vortex formation and shedding process is visualized using a high-speed schlieren camera and a holographic interferometric density measuring technique. The blade is equipped with a rotatable trailing edge cylinder instrumented side-by-side with a pneumatic pressure tap and a fast response pressure sensor for detailed measurements of the trailing edge pressure distribution. The experiments demonstrate that contrary to the isobaric dead air region demonstrated at low subsonic Mach numbers the data reveal the existence of a highly nonuniform trailing edge pressure distribution with a strong pressure minimum at the center of the trailing edge. This finding is significant for the determination of the base pressure coefficient that is in general measured with a single pressure-sensing hole at the trailing edge center. The paper investigates further the effect of the vortex shedding on the blade rear suction side and discusses the superposition of unsteady effects emanating from the trailing edge and from the neighboring blade. The experimental data are a unique source for the validation of unsteady Navier-Stokes codes.

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Blade tip vortices are the dominant vortical structures of the helicopter flow field. The inherent complexity of the vortex dynamics has led to an increasing interest in full-scale in situ experiments, where the near field, closely behind the blade, is of particular interest, since measures of vortex control mostly target this initial stage of development. To examine the near field, threecomponent particle image velocimetry (PIV) measurements of blade tip vortices of a full-scale helicopter in simulated hover flight in ground effect were conducted. A feasible and robust evaluation procedure was developed to minimise the shortcomings of full-scale PIV applications, such as a moderate spatial resolution and an elevated measurement noise level. At vortex ages ranging from w v ¼ 1 to 30°, a pronounced aperiodicity and asymmetry of the vortex were observed in -sections perpendicular to the vortex axes. At w v ¼ 1 , a preferential orientation of the vortex was observed. For increasing wake age, vortex wandering increased while the asymmetry of the vortex cores decreased. The high level of aperiodicity and core asymmetry must be taken into account when considering phase-averaged vortex characteristics in the near wake region. List of symbols