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The introduction of composite materials and the ability to tailor these materials to improve aerodynamic and structural performance is having a distinct effect upon aircraft design. In order to optimize the efficiency of the design procedure, a design process which is more integrated than the traditional approach is required. Currently the utilization of such design procedures produces enormous computational costs. An ongoing effort to reduce these costs is the development of efficient methods for cross-disciplinary sensitivities and approximate optimization techniques. The present research concentrates on investigating the integrated design optimization of a subsonic, forward-swept transport wing. A modular sensitivity approach for calculating the cross-sensitivity derivatives is employed. These derivatives are then used to guide the optimization process. The optimization process employed is an approximate technique due to the complexity of the analysis procedures. These optimization results are presented and the impact of the modular technique is discussed. / Master of Science

Past research at this university has proven the feasibility of the multi-disciplinary design of a complex system involving the complete interaction of aerodynamics and structural mechanics. Critical to this design process, is the ability to accurately and efficiently calculate the sensitivities of the involved quantities (such as drag and dynamic pressure) with respect to the design variables. These calculations had been addressed in past research, but it was felt that insufficient accuracy had been obtained. The focus of this research was to improve the accuracy of these sensitivity calculations with a thorough investigation of the computational aspects of the problem. These studies led to a more complete understanding of the source of the errors that plagued previous results and provided substantially improved sensitivity calculations. Additional research led to an improvement in the aerodynamic-structural interface which aided in the accuracy of the sensitivity computations. Furthermore, this new interface removed discontinuities in the calculation of the drag which the previous model tended to yield. These improvements were made possible with the application of shape functions in surface deflection analysis, instead of the previous ‘zonal’ approach. Other factors which led to accuracy improvements were changes to the aerodynamic model and the paneling scheme. Final studies with the optimization process demonstrated the ability of the improved sensitivities to accurately approximate the design problem and provided useful results. Additional studies on the optimization process itself provided information on move limit restrictions and various constraint problems. / Master of Science

The dissertation is focused on the large computational costs of integrated multidisciplinary design. Efficient techniques are developed to reduce the computational costs associated with integrated structural-aerodynamic design. First efficient methods for the calculations of the derivatives of the flexibility matrix and the aerodynamic influence coefficient matrix are developed. An adjoint method is used for the flexibility sensitivity, and a perturbation method is used for the aerodynamic sensitivity. Second a sequential optimization algorithm that employs approximate analysis methods is implemented. Finally, a modular sensitivity analysis, corresponding to the abstraction of a system as an assembly of interacting black boxes, is applied. This method was developed for calculating system sensitivity without modifying disciplinary black-box software packages. The modular approach permits the calculation of aeroelastic sensitivities without the expensive calculation of the derivatives of the flexibility matrix and the aerodynamic influence coefficient matrix. / Ph. D.

The rapid air travel growth during the last three decades, has resulted in runway congestion at major airports. The current airports infrastructure will not be able to support the rapid growth trends expected in the next decade. Changes or upgrades in infrastructure alone would not be able to satisfy the growth requirements, and new airplane concepts such as the NASA proposed Super Short Takeo and Landing and Extremely Short Takeo and Landing (ESTOL) are being vigorously pursued. Aircraft noise pollution during Takeoff and Landing is another serious concern and efforts are aimed to reduce the airframe noise produced by Conventional High Lift Devices during Takeoff and Landing. Circulation control technology has the prospect of being a good alternative to resolve both the aforesaid issues. Circulation control airfoils are not only capable of producing very high values of lift (Cl values in excess of 8.0) at zero degree angle of attack, but also eliminate the noise generated by the conventional high lift devices and their associated weight penalty as well as their complex operation and storage. This will ensure not only satisfying the small takeoff and landing distances, but minimal acoustic signature in accordance with FAA requirements. The Circulation Control relies on the tendency of an emanating wall jet to independently control the circulation and lift on an airfoil. Unlike, conventional airfoil where rear stagnation point is located at the sharp trailing edge, circulation control airfoils possess a round trailing edge, therefore the rear stagnation point is free to move. The location of rear stagnation point is controlled by the blown jet momentum. This provides a secondary control in the form of jet momentum with which the lift generated can be controlled rather the only available control of incidence (angle of attack) in case of conventional airfoils. The use of Circulation control despite its promising potential has been limited only to research applications due to the lack of a simple prediction capability. This research effort was focused on the creation of a rapid prediction capability of Circulation Control Aerodynamic Characteristics which could help designers with rapid performance estimates for design space exploration. A morphological matrix was created with the available set of options which could be chosen to create this prediction capability starting with purely analytical physics based modeling to high fidelity CFD codes. Based on the available constraints, and desired accuracy metamodels has been created around the two dimensional circulation control performance results computed using Navier Stokes Equations (Computational Fluid Dynamics). DSS2, a two dimensional RANS code written by Professor Lakshmi Sankar was utilized for circulation control airfoil characteristics. The CFD code was first applied to the NCCR 1510-7607N airfoil to validate the model with available experimental results. It was then applied to compute the results of a fractional factorial design of experiments array. Metamodels were formulated using the neural networks to the results obtained from the Design of Experiments. Additional validation runs were performed to validate the model predictions. Metamodels are not only capable of rapid performance prediction, but also help generate the relation trends of response matrices with control variables and capture the complex interactions between control variables. Quantitative as well as qualitative assessments of results were performed by computation of aerodynamic forces and moments and flow field visualizations. Wing characteristics in three dimensions were obtained by integration over the whole wing using Prandtl's Wing Theory. The baseline Super STOL configuration was then analyzed with the application of circulation control technology. The desired values of lift and drag to achieve the target values of Takeoff and Landing performance were compared with the optimal configurations obtained by the model. The same optimal configurations were then subjected to Super STOL cruise conditions to perform a tradeoff analysis between Takeoff and Cruise Performance. Supercritical airfoils modified for circulation control were also thoroughly analyzed for Takeoff and Cruise performance and may constitute a viable option for Super STOL and STOL Designs. The prediction capability produced by this research effort can be integrated with the current conceptual aircraft modeling and simulation framework. The prediction tool is applicable within the selected ranges of each variable, but methodology and formulation scheme adopted can be applied to any other design space exploration.

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

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.

A variable-complexity design strategy with combined aerodynamic and structural optimization procedures is presented for the high speed civil transport design (HSCT). Variable-complexity analysis methods are used to reduce the computational expense. A finite element-model based structural optimization procedure with flexible loads is implemented to evaluate the wing bending material weight. Static aeroelastic effects, evaluated through the comparison of rigid and flexible wing models, are found to be small in the HSCT design. The results of structural optimization are compared with two quasi-empirical weight equations. Good correlation is obtained between the structural optimization and one of the weight equations. Based on this comparison, an interlacing procedure is developed to combine both the simple weight equations and structural optimization in the HSCT design optimization, at modest computational cost. HSCT designs based on the interlacing procedure reveal that the aerodynamic optimizer may take advantage of weaknesses in weight equation. However, the optimizer may be unable to escape the local minimum due to the noisy of aerodynamic response and the lack of derivative information for the interlacing procedure. / Ph. D.

A novel combined piezoelectric-composite actuator configuration is proposed and analytically modeled in this work. The actuator is a low complexity, active compliant mechanism obtained by coupling a modified star cross sectional configuration composite beam with a helicoidal bimorph piezoelectric actuator coiled around it. This novel actuator is a good candidate as a hinge tension-torsion bar actuator for a helicopter rotor blade flap or blade tip and mirror rotational positioning. In the wing tip case, the tip deflection angle is different only according to the aerodynamic moment depending on the hinge position of the actuator along the chord and applied voltage because there is no centrifugal force. For an active blade tip subject to incompressible flow and 2D quasi steady airloads, its twist angle is related not only to aerodynamic moment and applied voltage but also to coupling terms, such as the trapeze effect and the tennis racquet effect. Results show the benefit of hinge position aft of the aerodynamic center, such that the blade tip response is amplified by airloads. Contrary to this effect, results also show that the centrifugal effects and inertial effect cause an amplitude reduction in the response. Summation of these effects determines the overall blade tip response. The results for a certain hinge position of Xh=1.5% chord aft of the quarter chord point proves that the tip deflection target design range[-2,+2] can be achieved for all pitch angle configurations chosen.