This study looks at the feasibility of using structural optimization techniques to address the problem of designing composite fuselage frames for crashworthiness. A key feature of any optimization strategy for increasing structural crashworthiness is a progressive failure analysis. Currently, the most widely used analysis methods for progressive failure of composite structures are considered too expensive computationally for practical optimization in today's computing environment. Developing an efficient analysis method for progressive failure of composite frames is a first step in the optimization for crashworthiness. In the current work a progressive failure analysis for thin-walled open cross-section curved composite frames is developed using a Vlasov type beam theory. A curved thin-walled composite beam theory is developed and a finite element implementation of the beam theory is used for progressive failure analysis. The accuracy and limitations of this analysis method are discussed. A model for progressive failure of the composite fuselage frame is developed from an extension of the laminate progressive failure analysis of Tsai-Wu. Comparisons based on a limited amount of available experimental data are encouraging. The first major failure event is captured by the theory, and the prediction of total energy absorbed follows the trend of the experimental data. It is believed that this accuracy is sufficient for preliminary design and optimization for crashworthiness. This progressive failure analysis is then incorporated into a frame optimization for crashworthiness based on the genetic algorithm method. The optimization methodology is demonstrated analytically to obtain frame designs with substantially increased crashworthlness. Laminate stacking sequence and cross-section shape are design variables for optimization / 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.

The object of this thesis is to determine the directional stability of a tricycle-geared aircraft during the landing roll. The motion of the aircraft is simulated by a computer program based on the appropriate equations of motion. Empirical aircraft and tire data are utilized in order to improve the simulation process. The stability of the aircraft is obtained by analyzing the motion of the vehicle after an initial disturbance about the vertical axis of the aircraft. The influence of the aircraft's velocity and other parameters on the stability is then determined. For the single-engine Cessna, the results show that a steady-state yaw angle is obtained after an initial disturbance. The results are presented graphically to show the effect of various parameters on the aircraft's stability. / Master of Science

In recent years, the use of computer-aided design (CAD) systems for conceptual aircraft design has greatly increased. As a result, new and better methods for creating surface models of aircraft geometry using dimensional parameters are needed. One such method, the Rule-Based Fuselage method, was suggested by Lockheed. The Rule-Based Fuselage method allows an aircraft designer to define complex aircraft fuselage geometry by specifying the fuselage profile and individual parametric cross-sections along the fuselage. This thesis describes the Rule-Based Fuselage method and discusses the implementation of the method in an interactive, object-oriented environment. Also included in this system is the Spine and Cross-Section method for creating arbitrarily shaped aircraft components. The design and implementation of both the Rule-Based Fuselage and Spine and Cross-Section methods are described. The integration of these methods with the conceptual aircraft design code, ACSYNT, is also discussed. / 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

A general numerical simulation of closely coupled lifting surfaces in steady and unsteady ground effects was developed. This model was coupled with the equations of motion to simulate aerodynamic-dynamic interaction. The resulting model was then coupled with a feedback-control law to form a general nonlinear unsteady numerical simulation of control of an aircraft in and out of ground effect. The aerodynamic model is based on the general unsteady vortex-lattice method and the method of images. It is not restricted by planform, angle of attack, sink rate, dihedral angle, twist, camber, etc. as long as stall or vortex bursting does not occur. In addition, it has the versatility to model steady and unsteady aerodynamic interference. The present model can be used to simulate any prescribed flare and to model the effects of cross and/or head winds near the ground. The present results show the influences of various parameters on the aerodynamic coefficients for both steady and unsteady flows. Generally, the ground increases the aerodynamic coefficients; the greater the sink rates, the stronger the effects. Increasing the aspect ratio increases both the steady and unsteady ground effects. An exception is a large aspect-ratio wing with large camber. The present results are generally in close agreement with limited exact solutions and experimental data. In the aerodynamic-dynamic simulation, the equations of motion were solved by Hammlng's predictor-corrector method. The aircraft, air stream, and control surfaces were treated as a single dynamic system. The entire set of governing equations was solved simultaneously and interactively. The aerodynamic-dynamic model was used to study a configuration that resembles a Cessna 182 airplane. The ground lowers the effectiveness of the tail in controlling pitch, increases the lift and drag, and makes the hinge-moment less negative. Proportional and rate control laws were used in a feedback system to control pitch. One set of gains was used in and out of ground effect. For the same control input, the pitch angle responds faster and overshoots more near the ground than it does far from the ground. The present results demonstrate the feasibility of using the current simulation to model more complicated motions and the Importance of including the unsteady ground effects when analyzing the performance of an airplane during a landing maneuver. / Ph. D.

Mathematical representation of military operations have long fascinated analysts and practitioners. In 1916 English mathematician Frederick W. Lanchester represented the attrition rates of two opposing forces in the form of two differential equations, functions of the size and combat effectiveness of each side. Lanchester's model was an intellectual breakthrough in the analysis of warfare insofar as it provided a deep insight into the possibilities inherent in simple models of combat. Interestingly enough, Lanchester's representation of the problem as a dynamic system is precisely the approach used in the system dynamics methodology employed here. In system dynamics, differential equations are converted to difference equations and there is virtually no limit to the number that can be employed to represent the known and complex details of a system. The attrition model developed here describes the interaction between twelve types of U.S. combat aircraft and twelve types of U.S.S.R. combat aircraft and indicates the winner or the loser at the end of an engagement or a battle during wartime. To guide the peacetime preparations, a generic baseline and modified aircraft are utilized and compared using an adaptation of the attrition model, so as to decide if the proposed modification of U.S. aircraft should be undertaken or not. Two measures of effectiveness are presented to evaluate the overall performance of the modified aircraft compared to the baseline aircraft -- decreased program life cycle cost and increased payload delivered to target per aircraft lost. Scenario analyses are performed to assess the combat aircraft effectiveness under changes to endogenous and exogenous parameters. / Ph. D.

Thesis: B.S., Massachusetts Institute of Technology, Department of Physics, 1982 / Vita. / Includes bibliographical references (leaf 185). / by Osa Edward Fitch. / B.S. / B.S. Massachusetts Institute of Technology, Department of Physics

Thesis (M.S.)--Massachusetts Institute of Technology, Dept. of Electrical Engineering and Computer Science; and, (M.S.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 1982. / MICROFICHE COPY AVAILABLE IN ARCHIVES AND ENGINEERING / Includes bibliographical references. / by Marc Bodson. / M.S.

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