A computational method has been developed toinvestigate the aerodynamic interaction between ahelicopter rotor and the fuselage in steady-statelevel forward flight. In order to model a compoundhelicopter the fuselage is fitted with wings. Themodel uses deformable vortex rings for the rotorwake, a source panel representation of the fuselageand a lifting surface method for the wings which canhandle arbitrary geometries. The wake representationis based on a free wake analysis, relaxing the nodalpoints on the vortex rings and wing wake elementsaccording to the induced velocities in the flowfield after every time step. In this way, neitherwake is restricted to planar shape. At each timestep the fuselage and wing forces were calculated.The trim attitude of the fuselage was determined bythe pressure field surrounding it using a potentialflow model for the external flow and a fullydeveloped turbulent boundary layer model on thefuselage surface. Parametric studies were carriedout to determine the influence of wing position,wing span, wing angle of incidence and advance ratioon the results. The computer program incorporates agraphical postprocessing output routine allowing thevisualisation of the wake shape, the inducedvelocity and the angles of incidence on the rotordisc to be made, together with the velocities atvarious wing stations and the lift coefficientsalong the span.

A rapid quasi-analytical method — low orderperturbation panel method — is presented forcalculating aerodynamic sensitivity derivatives forsubsonic wings. The method is based on the low orderpanel method with the internal Dirichlet problemformulation and analytical differentiations cascadedand inverted. In terms of doublet strengthsensitivity to configuration geometry, the computingcost of the present method for the partialderivative matrix calculation is less than one orderof magnitude than the existing high orderperturbation panel method, but the accuracy iscomparable. Furthermore, by applying a physicalinterpolation instead of a weighted geometricalinterpolation, the sensitivity derivatives ofpressure, lift and pitching moment coefficient withrespect to configuration geometry are derived andcalculated in the present paper. Such sensitivityderivatives are lacking in high order perturbationpanel method. A few calculative examples aregiven.

The high enthalpy and high Mach number effects on flatplate boundary layer flow are examined. Inparticular, the behaviour of the density profileunder high enthalpy, real gas conditions isconsidered. Mach-Zehnder interferograms of the flowover a flat plate are analysed using atwo-dimensional Fourier transform procedure in orderto obtain the density field. The comparison betweentheory and experiment is only fair. It must bepointed out, however, that the real gas effects onthe density profiles were found to be small.Estimation of the velocity boundary layer thicknessas obtained from a measurement of the thermalboundary layer thickness is seen to compare wellwith a theoretical prediction.