A theoretical investigation of short-scale inviscid instability in boundary-layer flow past an obstacle on a surface is described, with the application to trip-wire transition and wall roughness effects in mind. An analytical model and computational results show that the occurrence of an inflexion point in a local velocity profile is necessary but not sufficient for the instability to arise.

The synthesis of discrete-time tracking systems incorporating Lur’e plants with multiple non-linearities is illustrated by the design of a fast-sampling digital controller and associated transducers for the automatic control of the longitudinal motions of the CH-47 helicopter with both gang-collective and differential-collective non-linearities. In particular, it is demonstrated that non-interacting control of the vertical velocity and pitch attitude of the helicopter is readily achievable for large classes of non-linear actuator characteristics such as ‘deadzone’ provided that the controller and transducer parameters are chosen so as to ensure that state-bounded absolutely stable tracking occurs.

In this paper a direct digital control law algorithm is proposed which in addition to providing tight non-interacting tracking behaviour and excellent disturbance-rejection characteristics, provides for the compensation of finite-time delays in implementation. The control-law algorithm is defined and system stability is proved in the case of multi-input multi-output linear systems. The theory is synthetic and leads directly to the determination of the appropriate controller matrices.

The theory is illustrated by the presentation of simulation results in which the aircraft is represented by an analogue computer and the digital control system is implemented on a microprocessor. In the simulation study the controller is required to effect fuselage pitch pointing and vertical translation manoeuvres for the analogue computer representation of the YF-16 aircraft. It is shown that tight non-interacting control is achievable even when the control implemented is delayed by 0·1 seconds.

A computational method is presented for solving numerically the three dimensional Euler equations for transonic flow around practical aircraft forebodies. The Euler solver is pseudo-time dependent, split and cast in finite volume form. Shock waves are captured crisply without the need for additional smoothing by means of an operator-switching facility which more accurately reflects the direction of propagation of signals.

The method is illustrated by examples of computed external, axisymmetric flows and some, simulated, realistic aircraft fore-bodies. The computational meshes employed in the three dimensional cases are essentially of cylindrical polar, flow conforming, type and relatively coarse. Closer attention to the mesh generation is expected to refine the results presented here. The method is versatile, robust, and holds promise for treating complex three dimensional geometries within economically viable run times.