Minimum mass designs are obtained for a simple box beam structure subject to bending, torque and combined bending/torque load cases. These designs are obtained subject to point strain and linear buckling constraints. The present work differs from previous efforts in that special attention is payed to including the effects of subcomponent panel interaction in the optimal design process. Two different approaches are used to impose the buckling constraints. When the global approach is used, buckling constraints are imposed on the global structure via a linear eigenvalue analysis. This approach allows the subcomponent panels to interact in a realistic manner. The results obtained using this approach are compared to results obtained using a traditional, less expensive approach, called the local approach. When the local approach is used, in-plane loads are extracted from the global model and used to impose buckling constraints on each subcomponent panel individually. In the global cases, it is found that there can be significant interaction between skin, spar, and rib design variables. This coupling is weak or nonexistent in the local designs. It is determined that weight savings of up to 7% may be obtained by using the global approach instead of the local approach to design these structures. Several of the designs obtained using the linear buckling analysis are subjected to a geometrically nonlinear analysis. For the designs which were subjected to bending loads, the innermost rib panel begins to collapse at less than half the intended design load and in a mode different from that predicted by linear analysis. The discrepancy between the predicted linear and nonlinear responses is attributed to the effects of the nonlinear rib crushing load, and the parameter which controls this rib collapse failure mode is shown to be the rib thickness. The rib collapse failure mode may be avoided by increasing the rib thickness above the value obtained from the (linear analysis based) optimizer. It is concluded that it would be necessary to include geometric nonlinearities in the design optimization process if the true optimum in this case were to be found. / Master of Science

The objective of this research is to investigate the techniques and payoffs of integrated aircraft design. Lifting line theory and beam theory are used for the analysis of the aerodynamics and the structures of a composite sailplane wing. The wing is described by 33 - 34 design variables which involve the planform geometry, the twist distribution, and thicknesses of the spar caps, spar webs, and the skin at various stations along the wing. The wing design must satisfy 30 â 31 aeroelastic, structural, aerodynamic, and performance constraints. Two design procedures are investigated. The first, referred to as the iterative, sequential procedure, involves optimizing the aerodynamic design for maximum average cross-country speed at E1 constant structural weight, and then optimizing the the structural design of the resulting wing geometry for minimum weight. This value is then used in another aerodynamic optimization, and the process continues iteratively until the weight converges. The other procedure, the integrated one, simultaneously optimizes the aerodynamic and the structural design variables for either maximum average cross-country speed or minimum weight. The integrated procedure was able to improve the value of the objective function obtained by the iterative procedure in all cases. This shows The objective of this research is to investigate the techniques and payoffs of integrated aircraft design. Lifting line theory and beam theory are used for the analysis of the aerodynamics and the structures of a composite sailplane wing. The wing is described by 33 - 34 design variables which involve the planform geometry, the twist distribution, and thicknesses of the spar caps, spar webs, and the skin at various stations along the wing. The wing design must satisfy 30 â 31 aeroelastic, structural, aerodynamic, and performance constraints. Two design procedures are investigated. The first, referred to as the iterative, sequential procedure, involves optimizing the aerodynamic design for maximum average cross-country speed at E1 constant structural weight, and then optimizing the the structural design of the resulting wing geometry for minimum weight. This value is then used in another aerodynamic optimization, and the process continues iteratively until the weight converges. The other procedure, the integrated one, simultaneously optimizes the aerodynamic and the structural design variables for either maximum average cross-country speed or minimum weight. The integrated procedure was able to improve the value of the objective function obtained by the iterative procedure in all cases. This shows that definite benefits can be gained from taking advantage of aerodynamic/structural interactions during the design process. / Master of Science

M.S.

A Numerical Solution is presented for the problem of flow past a highly swept, slender wing with sharp leading edges. The lifting surface is modelled as a bound vortex sheet, while the wake is modelled as a force-free vortex sheet. The solution is obtained by the use of a unsteady Vortex-Lattice Method which includes the effect of leading edge separation. Numerical predictions for the aerodynamic loads and pressure distributions are compared with experimental data. A 75° Delta wing and a 60° Delta wing with Leading Edge Vortex flaps in uniform, symmetric and steady flow are studied. Uniform and cosine distributions are used to determine the effect of lattice shape on the solution. The results show that good aerodynamic load predictions are obtained by this Vortex-lattice method. The results also indicated that fewer cosine distribution control points predict pressures as well as the use of a larger number of uniform distribution control points. The numerical results for wings with LEVFs show good agreement with experimental data away from the trailing edge. This may be due to the viscous effects in the experiment not modelled in this method. It is also apparent that the size of the wake, trailing and leading edge wakes, is the important factor effecting computation times. / M.S.

A method is presented for calculating the shape sensitivity of a wing aeroelastic response with respect to changes in geometric shape. Yates’ modified strip method is used in conjunction with Giles' equivalent plate analysis to predict the flutter speed, frequency, and reduced frequency of the wing. Three methods are used to calculate the sensitivity of the eigenvalue. The first method is purely a finite difference calculation of the eigenvalue derivative directly from the solution of the flutter problem corresponding to the two different values of the shape parameters. The second method uses an analytic expression for the eigenvalue sensitivities of a general complex matrix, where the derivatives of the aerodynamic, mass, and stiffness matrices are computed using a finite difference approximation. The third method also uses an analytic expression for the eigenvalue sensitivities but the aerodynamic matrix is computed analytically. All three methods are found to be in good agreement with each other. The sensitivities of the eigenvalues were used to predict flutter speed, frequency , and reduced frequency. These approximations were found to be in good agreement with those obtained using a complete reanalysis. However, it is recommended that higher order terms be used in the calculations in order to assure greater accuracy. / Master of Science / incomplete_metadata

The interactions of trailing vortex pairs shed from the tips of two rectangular wings have been studied through helium bubble flow visualizations and extensive hot wire velocity measurements made between 10 and 30 chord lengths downstream. The wings were placed tip to tip at equal and opposite angles of attack, generating pairs of co-rotating and counter rotating vortices. Meaningful hot wire measurements could be made because the vortices were found to be insensitive to probe interference and experienced very small wandering motions. The co-rotating pairs were observed to rotate around each other and merge. Upstream of the merging location, the vortices have approximately elliptical cores. These are surrounded by the two wing wakes which join together around the two cores. Flow in the vicinity of the cores appears fully developed. During the merging process, the cores rotate rapidly about each other, winding the wing wakes into a fine spiral structure. Merger roughly doubles the core size and appears to produce turbulence over abroad range of frequencies. The counter rotating pairs move sideways under their mutual induction and slightly apart; their flow structure changing little with downstream location. These cores remain fairly circular and do not become fully developed within 30 chord lengths of the measurements. / Master of Science

The aircraft industry has long been concerned with the increase of drag on airplanes due to fouling of the wings by insects. The present research studied the effects of surface energy and surface roughness on the phenomenon of insect sticking. Aluminum plates of different roughnesses were coated with thin films of polymers with varying surface energies. The coated plates were attached to a custom jig and mounted on top of an automobile for insect collection. Contact angle measurements, x-ray photoelectron spectroscopy and specular reflectance infrared spectroscopy were used to characterize the surfaces before and after the insect impact experiments. Scanning electron microscopy showed the topography of insect residues on the exposed plates. Moments were calculated in order to find a correlation between the parameters studied and the amount of bugs collected on the plates. An effect of surface energy on the sticking of insect residues was demonstrated. / M.S.

This aircraft industry is concerned with the increase of drag on planes due to the sticking of insects on critical airfoil areas. The objectives of the present study were to investigate the effects of surface energy and elasticity on the number of insects sticking onto the polymer coatings on a modified aircraft wing and to determine the mechanism by which insects stick onto surfaces during a high-Velocity impact. Analyses including scanning electron microscopy (SEM), electron spectroscopy for chemical analysis (ESCA) and contact angle measurements of uncoated and polymer-coated aluminum surfaces have been performed. An air-gun was designed to accelerate insects to high speeds and impact them onto modified wing surfaces in a laboratory environment. A direct relation between the number of insects sticking on a sample and its surface energy was obtained. Since the sticky liquid from a burst-open insect will not spread on the low energy surface, it will ball up providing poor adhesion between the insect debris and the surface. The incoming air How can easily blow oH' the insect debris thus reducing the number of insects that remain stuck on the surface. Also a direct relation between the number of insect sticking onto sample surfaces and their moduli of elasticity was obtained. The deceleration of an insect impacting onto an elastomer reduces in proportional to the modulus of elasticity of the material. As a consequence, the rate of change of momentum is lower and the force and pressure exerted on the body of the insect is reduced if it impacts onto a material with a low modulus of elasticity. This lessens the chance of bursting the i insect exoskeleton. / Master of Science

Using a general, unsteady, nonlinear vortex lattice method, the aerodynamic loads have been found on a 75° delta wing with and without leading edge vortex flaps. The flap had an area approximately 26 percent of the wing area with a constant chord of 6.7 percent of the wing mean aerodynamic chord and was deflected at 30°. Results for lift, drag, axial force, and pitching moment coefficients are compared with experimental data and show very good agreement. Individual pressure difference coefficients along the wing and flap are also presented and compared with experimental data. Overall, the method shows the leading edge vortex flap to be very effective in reducing drag while maintaining lift comparable to that of the plain wing. / Master of Science

Engineering design is hierarchical in nature, and if no attempt is made to benefit from this hierarchical nature, design optimization can be very expensive. There are two alternatives to taking advantage of the hierarchical nature of structural design problems. Multi-level optimization techniques incorporate the hierarchy at the formulation stage, and result in the coordinated optimization of a hierarchy of subsystems. The use of multi-level optimization techniques often necessitates the use of equality constraints. These constraints can sometimes cause numerical difficulties during optimization. Single-level decomposition techniques take advantage of the hierarchical nature to reduce the optimization cost. In this research the decomposition approach has been followed to reduce the computational effort in a single-level design space. A decoupling technique has been developed that retains the advantages of a partitioned system of smaller independent subsystems without an increase in the total number of design variables. A penalty function formulation using Newton's method for the solution of a sequence of unconstrained minimizations was employed. The optimization of the decoupled system is cheaper due to (i) cheaper evaluation of the hessian matrix by taking advantage of its sparsity, (ii) fewer global analyses for constraint derivative calculations, and (iii) utilizing the decoupled nature of the hessian matrix in the solution process. Further, the memory requirements of the decoupled system are much less than that of the original coupled system. These benefits increase substantially for design problems with larger and larger number of detailed design variables. Orthotropic material properties as stiffness global variables have been shown to be effective as global variables for panels in a hierarchical wing design formulation. The proposed decoupling technique was implemented to minimize the volume of a portal frame and a wing box. Computational savings of up to 50 percent have been obtained for medium sized problems. The savings increase as the size of the problem and the amount of decoupling is increased. The procedure is simple to implement. For truly large systems this decoupling technique provides the necessary reduction of computational effort to make the optimization process viable. / Ph. D.