A computational method has been developed to investigate the aerodynamic interaction between a helicopter rotor and the fuselage in steady-state level forward flight. In order to model a compound helicopter the fuselage is fitted with wings. The model uses deformable vortex rings for the rotor wake, a source panel representation of the fuselage and a lifting surface method for the wings which can handle arbitrary geometries. The wake representation is based on a free wake analysis, relaxing the nodal points on the vortex rings and wing wake elements according to the induced velocities in the flow field after every time step. In this way, neither wake is restricted to planar shape. At each time step the fuselage and wing forces were calculated. The trim attitude of the fuselage was determined by the pressure field surrounding it using a potential flow model for the external flow and a fully developed turbulent boundary layer model on the fuselage surface. Parametric studies were carried out to determine the influence of wing position, wing span, wing angle of incidence and advance ratio on the results. The computer program incorporates a graphical postprocessing output routine allowing the visualisation of the wake shape, the induced velocity and the angles of incidence on the rotor disc to be made, together with the velocities at various wing stations and the lift coefficients along the span.

A rapid quasi-analytical method — low order perturbation panel method — is presented for calculating aerodynamic sensitivity derivatives for subsonic wings. The method is based on the low order panel method with the internal Dirichlet problem formulation and analytical differentiations cascaded and inverted. In terms of doublet strength sensitivity to configuration geometry, the computing cost of the present method for the partial derivative matrix calculation is less than one order of magnitude than the existing high order perturbation panel method, but the accuracy is comparable. Furthermore, by applying a physical interpolation instead of a weighted geometrical interpolation, the sensitivity derivatives of pressure, lift and pitching moment coefficient with respect to configuration geometry are derived and calculated in the present paper. Such sensitivity derivatives are lacking in high order perturbation panel method. A few calculative examples are given.

The high enthalpy and high Mach number effects on flat plate boundary layer flow are examined. In particular, the behaviour of the density profile under high enthalpy, real gas conditions is considered. Mach-Zehnder interferograms of the flow over a flat plate are analysed using a two-dimensional Fourier transform procedure in order to obtain the density field. The comparison between theory and experiment is only fair. It must be pointed out, however, that the real gas effects on the density profiles were found to be small. Estimation of the velocity boundary layer thickness as obtained from a measurement of the thermal boundary layer thickness is seen to compare well with a theoretical prediction.