An L-shaped tab was tested at the trailing edge of an oscillating airfoil to evaluate its effects on blades aerodynamic performance. The tests were conducted on a NACA 23012 pitching airfoil in deep dynamic stall conditions with the L-shaped tab fixed in two different positions. When deployed the tab is attached to the airfoil upper surface so that the end prong protrudes at the airfoil trailing edge. In retracted position the tab features an angle of 9.1 with the airfoil upper surface, since its prong tip touches the airfoil trailing edge. The airloads time histories during a pitching cycle were evaluated by pressure measurements carried out on the airfoil midspan contour. The phase-averaged flow field at the trailing edge region was investigated by means of particle image velocimetry to evaluate the detailed flow physics involved in the use of the device. The experimental results indicate that the use of such a pivoting L-shaped tab can introduce similar effects to those that can be obtained by the use of an active Gurney flap. Thus, the L-shaped tab can be considered an attractive device due to its easier integration on helicopter blades.

An extensive experimental investigation was conducted on an oscillating NACA 23012 airfoil to study the flow structures and the consequent performances in dynamic stall conditions. The testing activity involved two different measurement techniques: fast unsteady pressure measurements and particle image velocimetry. The analysis of the experimental data set made possible to achieve a deep insight in the mechanism of the dynamic stall phenomena for the NACA 23012 airfoil in the different dynamic stall regimes. In particular, the flow velocity field measured on the airfoil upper surface described in detail the mechanism of the formation, migration and shedding of strong vortical structures characteristic of the deep dynamic stall. In addition, Gurney flap effects were investigated. The experimental results showed that it would be advantageous to deploy active Gurney flaps to improve helicopter rotor blade performances. The whole set of experimental results can be considered as a reference to validate computational fluid dynamics tools.

In this study, experiments were performed to investigate the aerodynamic interaction between a helicopter and ground obstacles. A new experimental set-up was realised and validated. The motorised helicopter model, which included the fuselage, was positioned in different positions relative to a model building in order to replicate different hovering configurations. The use of a helicopter model with a six-component balance and a building model with several pressure taps allowed a database to be compiled for the loads on the helicopter and obstacle. First several tests were performed without the building in order to develop a reference database and assess the experimental set-up through a comparison with results in the literature. The measured loads were analysed to investigate the interference effects of the building model on the helicopter performance. A physical interpretation of the flow phenomena was obtained through analysis of the obstacle pressure measurements and particle image velocimetry surveys of relevant configurations.

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NOMENCLATURE a speed of sound, ms -1 c blade chord, m u′ 2 root mean square of velocitiy longitudinal component v′ 2 root mean square of velocitiy spanwise component w′ 2 root mean square of velocitiy vertical component M WT Mach number of wind-tunnel flow, M WT = V WT /a M MR Mach number of blade tip in hover, main rotor, M MR = (ΩR) MR /a M TR Mach number of blade tip in hover, tail rotor, M TR = (ΩR) TR /a R blade radius, m α S rotor shaft angle-of-attack α fus fuselage angle-of-attack μ advance ratio M WT /M MR ψ MR main rotor azimuth angle (0° = reference blade above the fuselage) Ω rotor rotational frequency, rad/s ABSTRACT The GOAHEAD (Generation of an Advanced Helicopter Experimental Aerodynamic Database for CFD code validation) consortium was created in the frame of an EU-project in order to create an experimental database for the validation of 3D-CFD and comprehensive aeromechanics methods for the prediction of unsteady viscous flows. This included the rotor dynamics for complete helicopter configurations, i.e. main rotor -fuselage -tail rotor configurations with emphasis on viscous phenomena like flow separation and transition from laminar to turbulent flow. The wind tunnel experiments have been performed during two weeks in the DNW-LLF on a Mach-scaled model of a modern transport helicopter consisting of the main rotor, the fuselage, control surfaces and the tail rotor. For the sake of controlled boundary conditions for later CFD validation, a closed test section has been used. The measurement comprised global forces of the main rotor and the fuselage, steady and unsteady pressures, transition positions, stream lines, position of flow separation, velocity profiles at the test section inlet, velocity fields in the model wake, vortex trajectories and elastic deformations of the main and tail rotor blades.