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.

A conceptual design methodology was produced and subsequently coded into a Visual C++ (GUI) environment to facilitate the rapid comparison of several possible configurations to satisfy High Altitude Long Endurance (FIALE) unmanned aircraft (UAV) missions in the Low Speed (propeller driven aircraft) regime. Several comparative studies were performed to verify the applicability of traditional design methods. The traditional computational design methodologies fail in several areas such as high aspect ratio wing weight estimation and design, low Reynolds number wing design, high altitude engine performance, low Reynolds number drag estimation, unmanned aircraft design, and the conceptual design of unconventional configurations. The methodology developed for this thesis was robust enough to allow not only for consideration of these areas of inadequacy in traditional methods, but also to allow for the inclusion of advancements in the relevant technologies as they become more widely available. The following configurations were evaluated for suitability to the Low Speed HALE UAV application: conventional, canard, twin boom, multiple fuselage (conventional or canard), tandem wing, multiple fuselage tandem wing or flying wing configuration. The configurations were compared on the basis of aircraft endurance for takeoff weights ranging from 2,000 to 20,000 pounds and wing loadings ranging from 5 to 25 lbs1fe. Initial drag estimates were made using traditional parabolic drag estimation techniques. A more refined drag buildup was performed using a vortex lattice drag estimation for the lift induced drag (for all lifting components) and calculated skin friction coefficients for the parasite drag. Statistically based methods were used for other components of drag having much smaller contributions. In addition, a statistical approach was taken to the weight estimation of the major aircraft components. However, this approach made comparison of alternative configurations more difficult. Thus wing bending moments trends were evaluated and utilized in the development of weight saving values for multiple fuselage wing weight estimation. The comparative performance of each configuration is justified with direct reference to the terms in the Breguet Endurance equation. Validation was performed where possible on all modules and segments associated with the methodology, as well as for the macroscopic results. In addition, parametric studies on endurance were performed for the conventional configuration for geometric characteristics and operating conditions directly and indirectly effecting the calculated endurance and generalized results presented. Finally, a case study was performed to demonstrate this capability. A new relation was developed for aircraft empty weight prediction, a low speed airfoil figure of merit was proposed, and new constants were offered for UAV fuselage length prediction. In addition, horizontal and vertical tail volume coefficients were proposed for all of the Low Speed HALE UAV configurations considered. It was determined that the multiple fuselage configurations showed comparatively superior endurance performance across a range of takeoff weights, with several other configurations demonstrating marginal endurance improvements. Finally, a highly flexible and robust computer based conceptual design methodology was developed and validated enabling the quick comparison of a greater number of possible configurations to satisfy a given mission for Low Speed HALE UAV's and providing detailed drag and weight breakdown data.

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Environmental factors, such as noise and emissions have begun to play a significant role in the design of new aircraft. Although advances in propulsion technology have reduced source noise levels significantly over the past few decades, it is becoming increasingly difficult to project similar advances for the next few decades. It is likely however that some noise benefits may come from improvements in aircraft performance and from changes in operational procedures. In order for such developments to be analysed at the conceptual design stage, an integrated conceptual aircraft design and aircraft noise model is required that enables the designer to rapidly assess the effect of key design parameters on reference noise levels and noise contour area.

Thesis (M.S.)--Brigham Young University. Dept. of Electrical and Computer Engineering, 2004. / Includes bibliographical references (p. 57-59).

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