The first part of the history of composites in aerospace emphasised materials with high specific strength and stiffness. This was followed by a quest for reliable manufacturing techniques that guaranteed sufficiently high fibre volume fractions in complex structural parts with reasonable cost. Further improvements are still possible leading, ultimately to an extension of the functionality of composite structures to non-mechanical functions. Reduction of material scatter and a more probability-based design approach, improved material properties, higher post buckling factors, improved crashworthiness concepts and improved NDI techniques are some of the evolutionary measures that could improve the performance of current composite structures. Modular design, increased co-curing, hybrid material structures, hybrid fabrication methods, innovative structural concepts and reduced development times are more revolutionary steps that could bring today’s solutions further. Manufacturing engineering is also important for achieving revolutionary change. Function integration such as embedded deicing, morphing,, and boundary-layer suction are among the next steps in weight and cost reduction, but now on the system level.

Military aircrafts are often subjected to severe flight maneuvers with high Angles-of -Attack (AOA) and Angles of Sideslip (AOSS). These flight attitudes induce non-uniform in flow conditions to their gas turbine engines which may include distortion of inlet total pressure and total temperature at the Aerodynamic Interface Plane (AIP). Operation of the downstream compression system with distorted inflow typically results in reduced aerodynamic performance, reduced stall margin, and increased blade stress levels. In the present study the steady state total pressure distortion induced to the Aerodynamic Interface Plane due to the aircraft’s flight attitude have been estimated in terms of distortion descriptors. The distorted conditions at the interface between the intake and the engine have been predicted by using Computational Fluid Dynamics (CFD), where 33 different aircraft flight attitudes have been tested. Based on the obtained results the effect of Angle-of-Attack (AOA) and Angle of Side Slip (AOSS) on the distortion descriptors have been studied. The results showed that the distortion effect becomes more pronounced whenever this specific airframe configuration is exposed to incoming flow with an AOSS. Among the tested cases, the greatest total pressure defect at the AIP in terms of difference from the average value and of circumferential extent was calculated for the flight attitudes of 0·35M flight with 0° AOA and 8° AOSS and 0·35M fight with 16° AOA and 16° AOSS.

The paper describes in detail an older method than Relaxation of approximating to the solution of equations of the Laplace and Poisson type. The corresponding fourth order equations are discussed briefly, and a method of dealing with certain non-linear equations is indicated. A description is also given of the propagation of errors in the fields due to various causes.

It has been found that the addition of a whirl velocity component to an otherwise axial flow in a conical diffuser can lead to improvements in static pressure recovery and in efficiency, according to a restricted definition. Diffusers of total conical angle 10, 20 and 30 deg. were tested in detail; improvements were evident only in the last two, which would otherwise have suffered from flow separation problems caused by their relatively large divergence angles. The agency for improvement seems to be a beneficial redistribution of axial velocities.

During the last fifty years much experimental and theoretical work has been carried out on the subject of the failure of struts, i.e., rods under axial compression. Most of this work has been devoted to the case of struts pin-jointed at both ends and only a small part examined the behaviour of rigidly encastred struts. From these investigations we know that in case of very slender struts the critical load for rigid encastrement is the fourfold of that for pin-jointed ends. This difference becomes less with decreasing slenderness of the strut, but even for slenderness ratios usually attained in aircraft framework it is considerable, if high tensile steel is used. In actual framework, compression members being more or less rigidly joined to nodes which may twist elastically, they represent a case between both forementioned extremes and may best be regarded as elastically encastred. Since according to above statements the load which may be sustained by a strut obviously depends a great deal upon end constraint, it is of great importance for design purposes to determine exactly the effect of elastic encastrement.

Recently, grid fins have been receiving increasing attention as a practical and efficient means of controlling missile trajectory. Preliminary studies at IAR have demonstrated that modern CFD methods can be used for computing flows past complex grid fin type configurations, and that these methods are more soundly physically based than the earlier vortex lattice and/or shock expansion methods.

The current paper addresses the issue of the grid fin size, in terms of both the panel thickness and the frontal shape. The study covers three thicknesses for the grid fin panel, with front shapes having a simple blunt square face, as well as a sharp knife-edge shape. In addition, an important aspect of the present investigation is to quantify the aerodynamic effect of the ramp fairing installed immediately upstream of the blunt base upon which the grid fin assembly resides. A comparison of flow field characteristics and aerodynamic coefficients of the grid fin assembly, with and without the fairing ramp, would provide a direct means of evaluating the effect of the ramp. At this stage, the investigations are based on Euler calculations. The present study focuses on a standard grid fin configuration mounted on a generic cylindrical body.

After some general remarks upon the constant flight characteristics of birds belonging to any given species, the author describes his method of obtaining, approximately, a cinematographic reproduction of three geometrical projections of the wing-beat. Dealing with one particular case he determines the trajectory in space of a pigeon's wing tip during an entire beat, and, also, the speed and acceleration reached by it.

Ladies and gentlemen, it is an honour and a great pleasure to present this Lanchester Memorial Lecture. I thank the Royal Aeronautical Society and its Aerodynamic Committee for inviting me. In preparing this lecture I greatly enjoyed the added significance of presenting it in the historical context set by Lanchester. But the real pleasure is to be here among many good friends with whom I worked together for shorter or longer periods.

“A body that in its motion through a fluid does not give rise to a surface of discontinuity.” So Lanchester defined a ‘ streamline body’ in his standard work Aerodynamics. With ‘ discontinuity’ the boundary is meant between the outer flow and the dead water region formed by fluid that departs from the surface as illustrated nicely in Fig. 1 for the flow around a cylinder.

A numerical method has been developed recently to solve the Navier-Stokes equations for laminar compressible flow around delta wings. A large-scale Navier-Stokes solution on a mesh of 129 × 49 × 65 points for transonic flow Mx = 0·85, α = 10° and Rex = 2·38 × 106 around a 65° swept delta wing with round leading edge is presented and compared with a correspondingly large-scale Euler solution. The viscous results reveal the presence of primary, secondary, and even tertiary vortices. The starting location of the primary vortex is seen to be quite different in the two solutions. In the viscous solution it starts at the wing apex, but in the Euler results it starts about one quarter chord downstream. The secondary reparations are also different, due to the up-lifting of the boundary layer in the viscous results, but to a cross-flow shock in the Euler computation. Comparison with experiment shows that the interaction between the primary and secondary vortices in the Navier-Stokes computation is obtained correctly and that these results are a more realistic simulation than the one given by the Euler equations.

The present paper addresses the problem of extracting indicial response functions from solutions of the Navier-Stokes equations. Two approaches are considered.

In the first, the indicial response is computed directly by modifying the local grid velocity to reflect the step change in plunge velocity. The computed data are then fitted using an assumed form for the response function and numerical optimisation techniques.

In the second (indirect) approach, the general form of the lift transfer operator is determined analytically from the functional form of the assumed response. Using the lift transfer operator explicit solutions of the in-phase and quadrature components of the normal force time history can be obtained. The unknown coefficients in the response function are then extracted from computed data for harmonic motion by fitting the time history using the explicit solution.

Computed response functions are presented for a range of Mach numbers and are compared with analytical results and indicial models currently used within the rotorcraft community. For the cases considered there is no significant difference between the response functions determined using the direct or indirect approach. Comparison of the computed results with existing models suggests that generalisation of the response to step changes in incidence using a two-pole approximation of the circulatory response together with the Prandtl-Glauert factor is not possible. However, comparison with higher order semi-empirical models is much improved, providing some justification for the use of such models. In addition to these observations, the extracted response functions suggest improvements to piston theory to account for thickness effects. The proposed modification of piston theory is based upon properties of the static pressure distribution and so can be used in conjunction with both CFD and experimental data.

During the past decade, the Navy has considerably improved its capabilities in aircraft-weapon integration. In 1989 it took more than 400 hours of wind-tunnel testing, which cost $1,5000,000, and 20 flights to clear the JSOW from the F-18 to Mach 0·95. This year the Mk-83 JDAM was cleared after only 60 hours of wind-tunnel testing and five flights to the full F-18 aircraft envelope of Mach 1·3. This reduction occurred because the Navy not only learned to test smarter, but also developed an integrated approach to modelling and simulation (M&S), wind-tunnel and flight testing which allowed lessons learned on previous programmes to be applied to new ones. However, the present approach still requires a fairly large commitment of time and financial resources to accomplish the mission.

The present approach has optimised the use of available resources; any further gains will have to come not by improving existing techniques, but by bringing new resources into the process.

The safety target for the Supersonic Transport has been set by the governments sponsoring the project. This is to achieve a safety level at least as good as that of subsonic aircraft at the same time. An analysis of present accident trends is used to give more precision to this statement. If the present rate of improvement in safety continues, the accident rate in the early 1970's can be expected to be some 20% to 25% below today's figures and the target for the SST must be for an improvement of at least 30%. To achieve improvements of this order in a safety record which is already good requires considerable effort by all concerned and necessitates a more scientific approach to safety requirements, especially airworthiness requirements. Some examples of the way this is being done in the United Kingdom and f ranee are given.

If the comparisons on productivity made by Mr. Harvey for the United Kingdom and other countries are accepted the question to be considered is why there should be any difference.

In particular I would like to identify some of the reasons for making comparisons with a principal competitor — the USA industry.

While the comments made reflect experience on airframe manufacture there is little doubt that the situation is no different in the rest of the aerospace industry or for that matter the UK engineering industry in general.

A method for accommodating engine deterioration via a scheduled linear parameter varying quadratic Lyapunov function (LPVQLF)-based controller is presented. The LPVQLF design methodology provides a means for developing unconditionally stable, robust control of linear parameter varying (LPV) systems. The controller is scheduled on the engine deterioration index, a function of estimated parameters that relate to engine health, and is computed using a multilayer feedforward neural network. Acceptable thrust response and tight control of exhaust gas temperature (EGT) is accomplished by adjusting the performance weighting on these parameters for different levels of engine degradation. Nonlinear simulations demonstrate that the controller achieves specified performance objectives while being robust to engine deterioration as well as engine-to-engine variations.

This paper examines the changing role aerodynamic technology plays in the design and development of modern combat aircraft. It reviews several aspects of aerodynamics which contribute to aircraft performance and handling characteristics. It considers the importance of a range of technologies, the contribution each makes to the final integrated solution and comments on the compromises necessary through the design cycle to optimise the overall weapon system, sometimes to the cost of one of the component technologies. The technologies discussed are in no way exhaustive but attempt to encapsulate the breadth of the subject, to illustrate its diversity and to point the way for the future development if aerodynamic technology is to continue to make an important contribution to the design of combat aircraft.

Each of the component technologies are discussed in terms of the contribution it makes, the tools and techniques used to predict, analyse and interpret the technology contribution, and the requirements for the direction of the development of the technology for the future.

Before considering the technologies themselves it is important to understand the environment in which they will be used, which increasingly conditions the manner in which they are employed and suggests the direction for their development.

This paper begins, therefore, by briefly reviewing the changes to the environment in which the fighter pilot operates and the manner by which his requirements, and therefore the specification of the aircraft, are defined before going on to look at the overall contribution each technology makes in the design process.

A numerical method, based on the panel technique, has been developed for calculating the time-averaged separated flow about a wedge. The mathematical model, with constant vorticity in the wake, gives an infinite extension of the wake and constant base pressure while incorporation of the dispersion of vorticity in the wake into the mathematical model leads to its finiteness and non-uniform under-pressure. Experimental tests have been conducted in a low speed wind tunnel on models haying different wedge angles. The theoretical and experimental results agree closely for different configurations and specific values of dispersion factors. However, the agreement for base pressure is poor.

A method for the generation of assumed modes in beams with arbitrary boundary conditions is described. The method is applied to two cantilevered beams using a modified Rayleigh-Ritz vibration analysis. The results obtained show excellent agreement with those obtained by exact methods of analysis.