In a major cooperative research program within existing US-Gennan and US-French Memoranda of Understandings (MoU's) a comprehensive experimental study was conducted with a 40-percent geometrically and dynamically scaled B0-105 main rotor operated in the open-jet anechoic test section of the German-Dutch Wind Tunnel (DNW). The objectives of the program were to improve the physical understanding and the mathematical modelling of the effects of the higher harmonic blade pitch control technique on blade-vortex interaction (BY!) impulsive noise and vibration reduction. A unique set of acoustic, aerodynarnic, dynamic response, performance, and rotor wake data were acquired with a pressure and strain gauge instrumented blade and by application of non~intrusive measurement techniques.

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The flow field around a helicopter is characterised by its inherent complexity including effects of fluidstructure interference, shock-boundary layer interaction, and dynamic stall. Since the advancement of computational fluid dynamics and computing capabilities has led to an increasing demand for experimental validation data, a comprehensive wind tunnel test campaign of a fully equipped and motorised generic medium transport helicopter was conducted in the framework of the GOAHEAD project. Different model configurations (with or without main/tail rotor blades) and several flight conditions were investigated. In this paper, the results of the three-component velocity field measurements around the model are surveyed. The effect of the interaction between the main rotor wake and the fuselage for cruise/tail shake flight conditions was analysed based on the flow characteristics downstream from the rotor hub and the rear fuselage hatch. The results indicated a sensible increment of the intensity of the vortex shedding from the lower part of the fuselage and a strong interaction between the blade vortex filaments and the wakes shed by the rotor hub and by the engine exhaust areas. The pitch-up phenomenon was addressed, detecting the blade tip vortices impacting on the horizontal tail plane. For high-speed forward flight, the shock wave formation on the advancing blade was detected, measuring the location on the blade chord and the intensity. Furthermore, dynamic stall on the retreating main rotor blade in high-speed forward flight was observed at r/R = 0.5 and 0.6. The analysis of the substructures forming the dynamic stall vortex revealed an unexpected spatial concentration suggesting a rotational stabilisation of large-scale structures on the blade. Abbreviations CFDComputational fluid mechanics DEHS Di-ethyl-hexyl-sebacat DS Dynamic stall GOAHEAD Generation of advanced helicopter experimental aerodynamic database for CFD code validation PIV Particle image velocimetry WT Wind tunnel Symbols a Speed of sound, m/s c Blade chord, m L Fuselage length, m L m Measurement volume length, m M Mach number r Radial coordinate, m r v Vortex radius, m r c Vortex core radius, m R Rotor radius, m t Time, s

In the original publication of this article, there are typographical errors in the first equation on page 154. We apologize for this mistake. The correct equation is:Accordingly the following sentences should read ''Therefore, the measurement error can be reduced by minimizing any one of three terms: W max /U max , 3/(DZ · M), and ex.''The online version of the original article can be found at http:// dx

An evaluation is made of extensive three-component (3-C) particle image velocimetry (PIV) measurements within the wake across a rotor disk plane. The model is a 40 percent scale BO-105 helicopter main rotor in forward flight simulation. This study is part of the HART II test program conducted in the German-Dutch Wind Tunnel (DNW). Included are wake vortex field measurements over the advancing and retreating sides of the rotor operating at a typical descent landing condition important for impulsive blade-vortex interaction (BVI) noise. Also included are advancing side results for rotor angle variations from climb to steep descent. Using detailed PIV vector maps of the vortex fields, methods of extracting key vortex parameters are examined and a new method was developed and evaluated. An objective processing method, involving a centerof-vorticity criterion and a vorticity "disk" integration, was used to determine vortex core size, strength, core velocity distribution characteristics, and unsteadiness. These parameters are mapped over the rotor disk and offer unique physical insight for these parameters of importance for rotor noise and vibration prediction. SYMBOLS C rotor blade chord, 0.121 m C T rotor thrust coefficient, thrust/ρπR 2 (ΩR) 2 CV center of vorticity in (x,y) plane, m DNW German-Dutch Wind tunnel DLR German Aerospace Center HART HHC Aeroacoustic Rotor Test HHC Higher Harmonic Control k grid point indices j indices for instantaneous image LDV Laser Doppler Velocimetry LLS Laser Light Sheet n number associated with analytical velocity profile, Eq. 5 R rotor radius, 2 m vortex core radius, m core radius associated with Rankine vortex, m shape factor for vortex velocity profile u,v,w velocity components for x,y,z coordinates, m/s rotor hover tip speed, RΩ (218 m/s) 'spin' velocity at r C , m/s (x,y,z) TUN windtunnel coordinate system: (x TUN positive downstream, y TUN positive starboard, z TUN positive up). x,y,z PIV image frame coordinates, m, Fig. 4. α rotor shaft angle with respect to z TUN axis, deg δ vortex wander parameter, m circulation within , m/s 2 σ standard deviation of parameters derived from instantaneous images Ω rotor rotation frequency, rad/s vorticity normal to (x, y) plane, s −1 Ψ blade azimuth angle (0 0 aft), deg.