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An assessment methodology has been developed for - use during the conceptual/preliminary design phase to quantify the effectiveness of newly designed aircraft. The effectiveness is measured by a squadron Sortie Generation Rate (SGR). Key elements of this methodology were the establishment of link parameters between design synthesis and the main effectiveness disciplines. These were Reliability and Maintainability (R&M), Survivability/Vulnerability and Acquisition Cost. A programmable solid modeller was used to create a solid CAD assembly of the aircraft critical components. A ray tracing technique has been used to develop an interactive vulnerability assessment tool. A Mission Simulation Model (MSM) has been developed which typically simulates the operation of a squadron of aircraft and gives the operational activities such as flying sorties and maintenance actions. The methodology has been validated based on real data from recent conflicts. The application aspects of the methodology have been demonstrated by quantifying the effectiveness of two recent combat aircraft.

This thesis presents the development of a Baseline Configuration for an Advanced Short Take-Off and Vertical Landing (ASTOVL) Combat Aircraft, the Design Synthesis and coding of this Baseline Congiguration (Code VERTI), the interfacing of the Design Synthesis Code VERTI with the Optimizer code RQPMIN and the optimization of the Baseline Configuration. The background and the objectives of this Research Programme are initially examined. The evaluation of the ASTOVL Combat Aircraft Baseline Configuration is then described, including all the problems, assumptions, choices and compromises that led to the specific configuration. The development of the Design Synthesis and the Code VERTI then follow, where the methodology used, the techniques adopted and the code operation are explained. A full description of the Design Synthesis is included as an appendix. Finally, the interfacing of Code VERTI with the optimizer RQPMIN and the optimization of the Baseline configuration are presented. The problems and difficulties of the RQPMIN operation are thoroughly discussed. The RQPMIN- VERTI code is used to optimize the initial Baseline Configuration and an optimization example is provided in appendix form. The optimized Baseline Configuration is partly validated against two ASTOVL combat aircraft designs. In addition to the optimization with the aircraft empty mass as objective function, a search for a better objective function is attempted.

Aerodynamic investigations have shown' that variable camber wings (VCW) for transport aircraft have considerable potential in terms of improving aircraft performance and enhancing their operational flexibility. In order to justify these benefits it is essential that the camber varying system is structurally and mechanically feasible. This research examined the feasibility of providing variable camber to two supercritical aerofoil sections of different'characteristics. The unique method of camber vaTiation was applied by rotating the forward and aft regions of the aerofoil on a circular arc and keeping the surface continuous and matching at their attachment to the main wing box. The change in camber thus increased the chord due to translational motion of the aforementioned regions. The geometries required for varying the forward camber by this method presented formidable design difficulties and no immediate solutions could be found. As a result, an alternative geometry was devised which accepts camber by simply drooping the nose region. A novel idea was developed for aft camber variation, which is considered to be universal for all supercritical aerofoil sections. The system utilises a tracking mechanism which guides a trailing edge element on a continuous arc. Surface continuity is provided by a flexible skin on the upper side and a spring loaded hinged panel on the under side. The flexible skin remains attached to the trailing edge element through a series of roller link arrangement which locate the skin in a separate guide rail. The large moment arm and therefore the increased torsional loads created due to the translational motion of the trailing edge element necessitated investigation of alternative deployment geometries. As a result two additional geometries were schemed. One had reduced radius of rotation and therefore reduced extension, while the other changed camber by drooping the aft region without any chordal extension. Since there was no aerodynamic evidence on the possible benefits offered by these geometries it was decide to postpone them until such information was available. Some detailed aspects of the proposed concept for aft camber variation were considered by applying the system to a modem transport aircraft wing. This resulted in a design which is practically feasible. Justification of this concept was made by designing and testing a half scale structural model of one trailing edge segment. Three dimensional (3-D) geometric investigation showed that the camber-varying elements ride on a frustum of a cone and therefore their deployment is skewed to the line of flight. The 3-D geometric implications of variable camber clearly suggested that the camber variation by rotation on a circular arc, on a tapered wing can be possible if the rotating element is made to flex and twist or it utilises a pin jointed arrangement. To provide the necessary flexibility to the trailing edge element, its structural box best be made from fibre reinforced plastic material. The deployment of the trailing edge element on the structural model was made possible by designing it in laminated wood. Comparison of the proposed variable camber system with a conventional single slotted flap arrangement suggests that the two systems could be equally complex but the variable camber could be slightly heavier., Further systems investigations are required to quantify overall aerodynamic, mass, and cost implications of the use of VCW on transport aircraft.

This thesis investigates the application of Multi-Disciplinary Design, Analysis and Optimisation to the design of a large civil airliner, similar in size as the future A3XX. For the first time structural optimisation, manufacturing cost and aerodynamic effects are simultaneously integrated within a realistic, complex aircraft design problem: the wing box of such a large airliner. A novel multi-level system was developed to incorporate structural effects and manufacturing cost: mass is treated at a top-level while costs are treated at a structural sub-level. It allows a designer to study cost changes with respect to design changes and the interaction of cost with other disciplines such as structures and aerodynamics. The flexibility of the system allows companies to import their own results or cost data and to perform cost studies based on historical data or highly novel processes. Structural optimisation of the wing box using MSc/NASTRAN and STARS, the development of a metal and composite cost model and the overall MDO methodology are being discussed.

The design of a model helicopter rotor blade incorporating a swept tip is described. The swept tip is chosen to provide a coupling between the lift at the tip and the blade twist thus achieving a variation of blade tvist both with azimuth and with forward speed. The design is the first stop in an investigation of aeroelastic tailoring as a means of reducing helicopter vibration and increasing rotor performance. The first prototype blade encountered stability problems and further designs were evaluated using a new modes/stability computer program developed within RAE. Comparisons are made between the stresses measured on a second stable swept tip blade and a dynamically similar rectangular blade. The results show that a beneficial twisting of the swept tip blade is achieved which enhances rotor perforawtnee and reduces the flatvise bending and torsional moments. Comparisons are made between the experimental results and those predicted by the RAE/VOL coupled modes analysis. The predictions are. reasonably accurate for the flatwise moments but a more representative model of the hub is needed to produce acceptable predictions of the torsional moments. The design of a further set of blades is discussed, the aim being to investigate the effects of introducing a strong coupling between the flap and torsion notions of the blade by sweeping back the shear centre. An analysis of-the results shows that there are large gains in blade stability with no severe adverse effects on blade loads. A theoretical investigation has been undertaken to observe the effects of tip sweep on the performance of a full size rotor. The results show that aft tip sweep can reduce control loads and rotor power for a rotor with a cambered aerofoil, section.

The aircraft design process is characterised by the application of a wide range of knowledge across many disciplines based upon a certain degree of judgement and experience of the designer. A two pass approach has been taken towards the development of an aircraft design expert system based on the requirements of two conceptually different design steps namely, wing design and aircraft configuration. The current status of the work is one where an actual program for wing design exists with supporting documentation, and a very effective examination of the knowledge base performed based on the detail investigation of overall aircraft design process with particular emphasis on the wing design and the aircraft configuration design steps. The approach taken accomplishes the objectives of the current research in defining the knowledge base, providing tools and specifications for tools to be used within an aircraft design expert system closely following the problem-solving techniques utilised by the design expert.

The results from a series of low speed wind tunnel tests on two half model highly swept wings (a symmetrical aerofoil section and a highly cambered aerofail section) are presented in order to examine the trailing edge flow separation mechanism and its development with wing sweep between 30' and 60'. The tests involved surface oil flow visualisation, smoke flow visualisation, surface static pressure and force balance measurements at streamwise chord Reynolds numbers from 1.5 x 105 to 5.2 x 106 and Mach number from 0.09 to 0.17. These results are used to assess two viscous-inviscid interaction CFD methods (BVGK and VFP) and two boundary layer methods (TAPERBL and WAKELAG) used to predict the flow over the highly cambered wing. A parametric study using cropped delta vane vortex generators in a co-rotating array was conducted on the 40' swept wing to investigate the effect of vane chordwise position, vane orientation, vane height relative to the boundary layer thickness and vane spacing on the prevention of the trailing edge separation. The performance of these flow control devices is assessed in terms of changes in; the wing surface flowfield, lift curve slope and the lift-dependant drag factor. In addition comparisons are made between the clean wing and flow control wing measured pressure distributions. The results and analysis show that the performance of the vortex generators is improved when the height of the vortex generator is approximately equal to that of the local boundary layer thickness and when the vane angular deflection to the local upstream flow direction is between 14' and 21'. The performance is also seen to depend on the vanes position ahead of separation and on the adverse pressure gradient to be restored and may also depend on a vane spacing made non-dimensional on the wing normal chord rather than the vane height. Similar performance improvements are observed with the wing swept to 50' using the positioning guidelines from this optimisation study. The performance of concave slats, canted cropped delta vanes, 'bent'wires and sub-boundary layer wires as vortex generating devices are seen to be not as effective as upright cropped delta vane vortex generators. To assist in the interpretation of the parametric vortex generator study a low speed wind tunnel technique is developed using shear stress sensitive liquid crystals to investigate the downstream development of vortices from cropped delta vane vortex generators. The results show that -- i) submerged vortices have less influence on the surface flow with downstream distance than vortices closer to the edge of the boundary layer, and ii) the primary increase in skin ffiction arises in the flow adjacent to the upflow side of the vortex. This area increases with vortex size. The results from this research programme are finally shown to be applicable in two market areas. The first is as a performance improvement on current highly swept winged military aircraft and the second is as flight controls on future aircraft from making the vortex generating devices active. The possible customers in these two areas are identified and marketing strategies developed for each case.

The research reported in this thesis is concerned with the structural weight optimization of aircraft lifting surfaces when subjected to the satisfaction of flutter requirements. The main text is intended primarily as an expository account on the work and as such it aims at introducing and defining the subject of research and presenting the results. Accordingly, the mathematics have been simplified to the utmost in the main text and heavy theoretical treatments are revealed in the appendices. As the aim of this work is not directed at in-depth studies of the physical nature of flutter nor for a comprehensive treatment of structural optimization, the basic concepts of these two subjects are touched upon in the beginnings of chapters II and III respectively. We concluded these two chapters by clarifying the class of flutter, constraints and design variables for which the program we developed is designed. We endeavored to keep the problem to within certain practical boundaries without loosing too much of either its generality or its applicability to structures in realistic operational environments.