This paper deals with the problem of determining anoptimal controller which minimises thenorm of a given aircraftwhile guaranteeing that an upper limit of its norm is not exceeded. Theproblem is tackled by means of a linear matrixinequality formulation, which allows one toconstrain the eigenvalues of the model to within afixed region of the complex plane. Key features ofthe method are the possibility of simultaneously

• 1. keeping the maximum value of the inputdemand under control

• 2. taking flying quality requirements intoaccount

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

The approach is simple to handle and is well suited tothe design of aircraft stability augmentationsystems. A discussion of two case studiesdemonstrates the effectiveness of the procedure.

The start condition of a second-throat ejector-diffuser(STED) system has been studied by solving theaxisymmetric Navier-Stokes equations with a highorder conservative supra-characteristics method(CSCM). According to the conventional concept, whena STED is started, the flow passing through thesecond-throat contraction should be supersonic sothat the normal shock swallowing condition applyingfor a supersonic windtunnel holds. The presentnumerical results, however, suggest that a STED canbe started with a normal shock forming ahead of thesecond-throat contraction because the high speedflow in the region close to the wall of the diffuserserves to release the mass blocked by the normalshock, leading to a decrease in chamber pressure aswell as the start of the system. This not onlycontradicts the conventional concept of the start ofa STED, but also provides an explanation for theexperimental discrepancy between the startconditions of a STED and a supersonicwindtunnel.

Recent advances in two lifting surface methods arereported. The present subsonic constant pressuremethod improves upon the program robustness of thedoublet lattice method. The present unifiedsupersonic-hypersonic method extends the range ofapplicability of lifting surface theory to theNewtonian limit. Based on a uniformly-valid seriesformulation, the method accurately approximates thenon-linear thickness effect, whereas this effect isneglected by the linear theory and overestimated bypiston theory. In fact, the present resultsgenerally agree well with the trends of severalEuler solutions. Among other findings, casescomputed by the unified method confirm that theeffect of thickness is to reduce the supersonicflutter speed.

The paper describes the aeroelastic analysis for thedesign, the ground verification tests and the flighttest programme, for the Ranger 2000 trainingaircraft, developed jointly by Dasa and RockwellInternational between 1991 and 1994, as a competitorfor the next generation US Air Force and Navy JointPrimary Advanced Training System (JPats). Specialefforts with respect to the aeroelastic stabilitywere required for the T-tail configuration, for thedesign of the manual flight control system, and forthe establishment of sufficient mass balance for thecontrol surfaces. Several ground vibration testswere performed for the complete aircraft and forindividual components for all major designimprovements during the flight test evaluation. Tominimise the required time for these tests, highlymodular test equipment was required. For anefficient flutter flight test programme a reliableexcitation system was chosen. This system consistsof slotted rotating cylinders, mounted on smallvanes, which can be attached to any aerodynamicsurface. This equipment creates defined unsteadyaerodynamic forces to excite the eigenmodes of thestructure.

Quasi-on-line frequency and damping data evaluationbetween consecutive flight test points was madepossible by installing the required hardware andsoftware directly at the flight test quick-lookcontrol room, for the direct use of telemetry data.With this approach only a small number of dedicatedflutter flights was required.

Some earlier investigations on the flows in S-shapeddiffusing ducts suggested that the first bend wasthe primary source of pressure loss and engine-faceflow distortion, e.g., Goldsmith. By increasing thecurvature ratio (ratio of the duct curvature to theduct inlet cross-sectional diameter) of the firstbend and by incorporation of a straight diffuserjoining the end of the first bend and the inlet ofthe second bend, a significant improvement in thetotal pressure recovery at the engine face could beachieved (Fig. 1). A large proportion of the gain inthe total pressure was attributed to the centralstraight diffuser. Results at a lower inlet speed(incompressible; inlet Reynolds number about 60 000)in the same ducts by Ho, Myring andLivesey(2) revealed the same trend asin the original high inlet speed reported byGoldsmith(1) (compressible; inlet Machnumber from 0·4 to 0·8; inlet Reynolds number overone million). However, changing the curvature ratioin the first bend also changes the direction of theaxis of the inlet plane, which should be aligned inthe same direction as the axis of the exit plane(cf Fig. 1). Thus, the previousexperiments were actually comparing ducts atdifferent angles of attack relative to the axis ofthe exit plane; in relation to a typical jetaircraft intake installation. Hence, theeffectiveness of the straight diffuser at zero angleof attack (relative to the axes of the inlet andexit planes) on the overall total pressure recoveryin an S-shaped duct had not been assessed.

Results are presented from low-speed windtunnel testson an isolated generic chined forebody withpneumatic methods of yaw control at high alpha.Measurements of forces and moments were made on theforebody at angles of attack in the range 0° ≤ ɑ ≤60°. Slot blowing through the chine edge at variousangles and upper surface tangential blowing havebeen compared, and optimum configurations have beenidentified. The effects of forebody cross-sectionalshape have also been investigated and have beenfound to be highly critical in determining blowingeffectiveness. Limited flowfield visualisation isalso presented in order to gain some understandingof the flow field characteristics.