The performance of a helicopter rotor has always been dogged by the dissymmetry of air velocity over the blades caused by its translation in a direction substantially parallel to the plane of the rotor disc. Furthermore, in order to make the rotor even usable, this dissymmetry of lift has caused the introduction of major mechanical complications to the rotor hub. In forward flight the rotor is aerodynamically split into two halves either side of the flight direction. The advancing side, where the rotor rotation is in the same sense as the forward velocity, will see a greater dynamic head than the retreating side, where the rotor rotation and forward speed are in opposition. A rigid rotor would consequently suffer a major rolling moment and if a viable proposition is to be achieved the rolling moment must be avoided.

Two numerical methods are presented for the computation of steady and unsteady Euler flows. These are applied to steady and unsteady flows about the NACA 0012 aerofoil, using structured grids generated by the transfinite interpolation technique. An explicit central-difference scheme is produced based on the cell-vertex method of Ni modified by Hall. The method is second-order accurate in time and space, and with flow quantities stored at boundaries the boundary conditions are simple to apply. This is a definite advantage over the cell-centred approach of Jameson, where extrapolation of the flow quantities is required at the boundaries, making unsteady boundary conditions difficult to apply. An explicit upwind-biased scheme is also produced, based on the flux-vector splitting of van Leer. The method adopts a three stage Runge-Kutta time-stepping scheme and a high-order spatial discretisation which is formally third-order accurate for one-dimensional calculations. The upwind scheme is shown to be slightly more accurate than the central-difference scheme for steady aerofoil flows, but it is not clear which is the more accurate for unsteady aerofoil flows. However, the central-difference scheme requires less than half the CPU time of the upwind-difference scheme, and hence is attractive, especially when considering three-dimensional flows. The transfinite interpolation technique is ideal for generating moving structured grids due to its simplicity, and grid speeds are available algebraically by the same interpolation as grid points. The method is also ideal for use in a multi-block approach.

The variation of the stress intensity factor (SIF) for cracks growing in integrally machined stiffened skin panels have been studied in an isotropic material. The stringer-skin panel has been modelled using the Nastran finite element system with allowance for crack tip singularity by use of crack tip elements. To allow for variations in stringer and skin geometry, three stiffness ratios (stringer stiffness-skin stiffness) were investigated.

The stress intensity factors have been converted into nondimensionalised values of stress intensity factor coefficients (K1/K0). A damage tolerant and residual strength analysis has been carried out using an in-house computer program. Results indicate very little difference for the different stiffness ratios examined.

The dynamic stability derivatives of blunt cones for small variations in angles of attack have been previously derived by the current author. However, no account of the unsteady nature of the boundary layer was made in that work. In this paper closed form expressions for the increment in dynamic stability due to the presence of the boundary layer are derived by considering the pressure distribution perturbations as the boundary layer continuously adjusts to the enclosed oscillating body. The theory provides a first hand estimate of the complete pitch derivative damping without having to resort to more rigorous and expensive computational methods. Calculations performed at Mach numbers of 7 to 10 with axis positions ranging from 0·5 to 0·7 of the chord length for cones of semi-angle 10° and 20°, indicate that the effect of boundary layer is to slightly reduce the magnitude of the inviscid damping derivative. For blunt cones at angles of attack less than 5°, this was in good agreement with the limited experimental data available.