This paper deals with the problem of determining an optimal controller which minimises the norm of a given aircraft while guaranteeing that an upper limit of its norm is not exceeded. The problem is tackled by means of a linear matrix inequality formulation, which allows one to constrain the eigenvalues of the model to within a fixed region of the complex plane. Key features of the method are the possibility of simultaneously

• 1. keeping the maximum value of the input demand under control

• 2. taking flying quality requirements into account

• 3. assuring a minimum level of system robustness against uncertainties of the model.

The approach is simple to handle and is well suited to the design of aircraft stability augmentation systems. A discussion of two case studies demonstrates the effectiveness of the procedure.

The start condition of a second-throat ejector-diffuser (STED) system has been studied by solving the axisymmetric Navier-Stokes equations with a high order conservative supra-characteristics method (CSCM). According to the conventional concept, when a STED is started, the flow passing through the second-throat contraction should be supersonic so that the normal shock swallowing condition applying for a supersonic windtunnel holds. The present numerical results, however, suggest that a STED can be started with a normal shock forming ahead of the second-throat contraction because the high speed flow in the region close to the wall of the diffuser serves to release the mass blocked by the normal shock, leading to a decrease in chamber pressure as well as the start of the system. This not only contradicts the conventional concept of the start of a STED, but also provides an explanation for the experimental discrepancy between the start conditions of a STED and a supersonic windtunnel.

Recent advances in two lifting surface methods are reported. The present subsonic constant pressure method improves upon the program robustness of the doublet lattice method. The present unified supersonic-hypersonic method extends the range of applicability of lifting surface theory to the Newtonian limit. Based on a uniformly-valid series formulation, the method accurately approximates the non-linear thickness effect, whereas this effect is neglected by the linear theory and overestimated by piston theory. In fact, the present results generally agree well with the trends of several Euler solutions. Among other findings, cases computed by the unified method confirm that the effect of thickness is to reduce the supersonic flutter speed.

The paper describes the aeroelastic analysis for the design, the ground verification tests and the flight test programme, for the Ranger 2000 training aircraft, developed jointly by Dasa and Rockwell International between 1991 and 1994, as a competitor for the next generation US Air Force and Navy Joint Primary Advanced Training System (JPats). Special efforts with respect to the aeroelastic stability were required for the T-tail configuration, for the design of the manual flight control system, and for the establishment of sufficient mass balance for the control surfaces. Several ground vibration tests were performed for the complete aircraft and for individual components for all major design improvements during the flight test evaluation. To minimise the required time for these tests, highly modular test equipment was required. For an efficient flutter flight test programme a reliable excitation system was chosen. This system consists of slotted rotating cylinders, mounted on small vanes, which can be attached to any aerodynamic surface. This equipment creates defined unsteady aerodynamic forces to excite the eigenmodes of the structure.

Quasi-on-line frequency and damping data evaluation between consecutive flight test points was made possible by installing the required hardware and software directly at the flight test quick-look control room, for the direct use of telemetry data. With this approach only a small number of dedicated flutter flights was required.

Some earlier investigations on the flows in S-shaped diffusing ducts suggested that the first bend was the primary source of pressure loss and engine-face flow distortion, e.g., Goldsmith. By increasing the curvature ratio (ratio of the duct curvature to the duct inlet cross-sectional diameter) of the first bend and by incorporation of a straight diffuser joining the end of the first bend and the inlet of the second bend, a significant improvement in the total pressure recovery at the engine face could be achieved (Fig. 1). A large proportion of the gain in the total pressure was attributed to the central straight diffuser. Results at a lower inlet speed (incompressible; inlet Reynolds number about 60 000) in the same ducts by Ho, Myring and Livesey(2) revealed the same trend as in the original high inlet speed reported by Goldsmith(1) (compressible; inlet Mach number from 0·4 to 0·8; inlet Reynolds number over one million). However, changing the curvature ratio in the first bend also changes the direction of the axis of the inlet plane, which should be aligned in the same direction as the axis of the exit plane (cf Fig. 1). Thus, the previous experiments were actually comparing ducts at different angles of attack relative to the axis of the exit plane; in relation to a typical jet aircraft intake installation. Hence, the effectiveness of the straight diffuser at zero angle of attack (relative to the axes of the inlet and exit planes) on the overall total pressure recovery in an S-shaped duct had not been assessed.

Results are presented from low-speed windtunnel tests on an isolated generic chined forebody with pneumatic methods of yaw control at high alpha. Measurements of forces and moments were made on the forebody at angles of attack in the range 0° ≤ ɑ ≤ 60°. Slot blowing through the chine edge at various angles and upper surface tangential blowing have been compared, and optimum configurations have been identified. The effects of forebody cross-sectional shape have also been investigated and have been found to be highly critical in determining blowing effectiveness. Limited flowfield visualisation is also presented in order to gain some understanding of the flow field characteristics.