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  • About
  • The Global ETD Search service is a free service for researchers to find electronic theses and dissertations. This service is provided by the Networked Digital Library of Theses and Dissertations.
    Our metadata is collected from universities around the world. If you manage a university/consortium/country archive and want to be added, details can be found on the NDLTD website.
381

Intersection and filleting of non-uniform B-spline surfaces

Jones, Robert W. 24 April 2009 (has links)
Preliminary aircraft design codes requIre a more complete and integrated geometry definition than that used by conceptual design codes. This thesis documents the design and creation of an interactive CAD system which converts the geometry descriptions commonly used in conceptual aircraft design codes to descriptions that meet the requirements of preliminary design systems. In particular, the conversion of ACSYNT Hermite surface data of aircraft models to the non-uniform hi-cubic B-Spline surface representation is addressed. The topics discussed in this thesis include the design and development of an interactive graphics user interface, the design and coding of an intersection method for non-uniform hi-cubic B-Spline surfaces utilizing subdivision techniques and the development of a one-dimensional filleting algorithm for blending surfaces along iso-parametric curves. / Master of Science
382

Model-following control applications to nonlinear mechanical systems

Barlas, Mustafa Remzi 31 October 2009 (has links)
Model-following control design methodology is introduced for nonlinear plants and models. The plant equations are considered to be linear in the control input. Dynamic matching conditions are presented and the resulting error dynamics are given. The stability of error dynamics is ensured, using Liapunov's second theorem; by modifying the model state rates, which effectively introduces error feedback. The methodology is applied to two problems. Motion control of an n-link manipulator with torque controllers on each linkage, and control of an aircraft lacking direct control of lift and side force. The former represents the systems where all of the degrees of freedom can be controlled, and the latter represents the systems where only some of the degrees of freedom can be controlled. The aircraft control problem is analyzed in more detail. The resulting control law does not require any explicit gain scheduling, but instead, requires estimates of the stability and control derivatives. A method is proposed to compensate for actuator dynamics. The control law is then verified by simulating some maneuvers on the aircraft model provided for the AIAA Controls Design Challenge, which includes nonlinear and full-envelope aerodynamic and engine models, and rate and position limited controls. The maneuvers simulated include a level acceleration and a 3-g turn. / Master of Science
383

Structural efficiency study of composite wing rib structures

Swanson, Gary D. 29 April 2010 (has links)
A series of short stiffened panel designs which may be applied to a preliminary design assessment of an aircraft wing rib is presented. The computer program PASCO is used as the primary design and analysis tool to assess the structural efficiency and geometry of a tailored corrugated panel, a corrugated panel with a continuous laminate, a hat stiffened panel, a blade stiffened panel, and an unstiffened nat plate. To correct some of the shortcomings in the PASCO analysis when shear is present a two-step iterative process using the computer program VICON is used. The loadings considered include combinations of axial compression, shear, and lateral pressure. The loading ranges considered are broad enough such that the designs presented may be applied to other stiffened panel applications. An assessment is made of laminate variations, increased spacing. and non-optimum geometric variations, including a beaded panel. on the design of the panels. / Master of Science
384

Response of a nonlinear two-degree-of-freedom system subjected to an impact loading

Theisen, Jerome G. 04 August 2009 (has links)
The solution of the equations of motion of an aircraft fuselage-landing gear configuration during landings is of interest to the designer who must predict the landing loads which an airplane encounters in service. In general such solutions are difficult because of the highly nonlinear characteristics of the oleo-pneumatic shock strut which couples the lower mass of the landing gear to the fuselage. In the past, attempts to obtain solutions by linearization of these shock strut characteristics have resulted in unrealistic predictions of landing gear motions. Therefore, it has been necessary to carry out most of the theoretical analysis associated with landing gears by means of numerical integration procedures. These numerical methods are tedious, and as a result a large portion of design work has been carried out by means of trial and error drop testing of a system of masses representative of an airplane and landing gear. This in turn has proved to be time consuming and expensive. This paper presents a method for obtaining an analytical solution of the equations of motion for a basically nonlinear system which closely resembles an actual airplane and landing gear configuration. The nonlinear system considered has two degrees of freedom and is composed of a large mass representative of the fuselage-wing combination connected by an oleo-pneumatic shock strut to a wheel. The shock strut is assumed to have velocity-squared hydraulic damping and coulomb friction forces on the strut bearings. The nonlinear spring characteristic of the tire is represented by a sectionally linear spring. In the first part of this paper the equations of motion for this nonlinear system are derived making use of a few simplifications which previous papers have shown to be justified. Also, the degree to which these assumptions limit the results is discussed. Next these equations of motion are solved in analytical form by a method which may be called "equivalent non linearisation." It is shown that this solution is exact only for a specific combination of impact parameters, but that for a wide range of parameters the solution describes the motion of the system adequately for design purposes. Finally, a few analytical solutions are compared with solutions obtained by numerical integration methods; and the results are compared with experimental data for a typical impact. / Master of Science
385

Integral aerodynamic-structural-control wing design

Rais-Rohani, Masoud 14 October 2005 (has links)
The aerodynamic-structural-control design of a simplified wing and a forward-swept composite wing are studied. In the first example, the wing is modeled as a beam with a control surface near the wing tip. The torsional stiffness is the only physical property varying along the span. The aerodynamic model is based on strip theory, and the control model is based on output feed-back control. With the structural-control interaction being the main focus, two different approaches are taken for the simplified wing design: (1) a sequential approach, (2) an integrated approach. In each approach the wing is designed for minimum weight subject to divergence and control deflection constraints. The results of this study indicated that while the integrated approach produced a better design than the sequential approach, the difference was minimal. In the second example, a forward-swept composite wing is designed for a high subsonic transport aircraft. The structural analysis is based on finite-element method. The aerodynamic calculations are based on vortex-lattice method, and the control calculations are based on output feed-back control. The wing is designed for minimum weight subject to structural, aerodynamic/performance and control constraints. Efficient methods are used to calculate the control deflection and efficiency sensitivities which appear as second order derivatives in the control constraint equations. To suppress the aeroelastic divergence of the forward-swept wing, and to reduce the gross weight of the design aircraft, two separate cases are studied: (1) combined application of aeroelastic tailoring and active controls, (2) aeroelastic tailoring alone. The results of this study indicated that, for this particular example, aeroelastic tailoring is sufficient for suppressing the aeroelastic divergence, and the use of active controls was not necessary. / Ph. D.
386

System reliability optimization of aircraft wings

Yang, Ju-Sung January 1989 (has links)
System reliability based design of aircraft wings is studied. A wing of a light commuter aircraft designed according to the FAA regulations is compared with one designed by system reliability optimization. Both the level III, and the advanced first order, second moment (AFOSM) method are employed to evaluate the probability of failure of each failure element of the system representing the wing. In the level III method the statistical correlation between failure modes is neglected. The AFOSM method allows to evaluate the sensitivity derivatives of the system safety index analytically. Furthermore, it accounts for the statistical correlation between failure modes. The results demonstrate the potential of stochastic optimization, and the importance of accounting for the statistical correlation between failure modes. Finally, it is shown that the problem associated with discontinuity of sensitivity derivatives, encountered when using second order Ditlevsen upper bounds to estimate the system failure probability, is circumvented if a penalty function method is used for optimization. / Ph. D.
387

An investigation into the finite element modelling of an aircraft tyre and wheel assembly

Guo, H. January 2014 (has links)
This thesis reports the investigation into the modelling and simulation of an aircraft tyre and wheel assembly in finite element environment. The finite element simulations basing on aircraft tyre test and operational scenarios could predict the loads transferred from tyre and the stresses distributed to the wheel rim. The virtual analysis could assess the safety criteria of different tyre structures, which would lead to the cost and time circle reduction in tyre R&D process. An H41x16.0R20 radial ply aircraft test tyre and its corresponding test wheel, provided by Dunlop Aircraft Tyres Limited, are adopted as the subject of this research. The material properties, especially the rubber and fabric materials, have been investigated. The finite element hyperelastic models have been utilized to represent rubbers and been correlated to experimental data. The 2D and 3D finite element tyre models, along with the finite element wheel models are created in the commercial finite element code, LS-Dyna. The finite element models have been validated with either industrial standardised simulation results or experimental data. Basing on the validated models, simulations that duplicating static test and dynamic operational scenarios have been developed. The researches have provided knowledge in comparing single and double bead tyre designs with respect to wheel loading mechanisms. The computational model also allowed manufacturers to assess the performance and safety criteria of a particular tyre at its design stage. The development of such models would add to the general drive towards the use of more virtual prototypes in an area traditionally reliant on experimental testing.
388

Influence of vane sweep on rotor-stator interaction noise.

Envia, Edmane. January 1988 (has links)
In this dissertation the influence of vane sweep on rotor-stator interaction noise is investigated. In an analytical approach, the interaction of a convected gust, representing the rotor viscous wake, with a cascade of finite span swept airfoils, representing the stator, is analyzed. The analysis is based on the solution of the exact linearized equations of motion. High-frequency convected gusts for which noise generation is concentrated near the leading edge of the airfoils are considered. In a preliminary study, the problem of an isolated finite span swept airfoil interacting with a convected gust is analyzed. Using Fourier transform methods and the Wiener-Hopf technique, an approximate solution for this problem is developed. Closed form expressions for the acoustic farfield are obtained and used in a parametric study to assess the effect of airfoil sweep on noise generation. Results indicate that sweep can substantially reduce the farfield noise levels for a single airfoil. Utilizing the single airfoil model, an approximate solution to the problem of noise radiation from a cascade of finite span swept airfoils interacting with a convected gust is derived. Only upstream radiated noise is considered. Neglecting the weak coupling between the adjacent leading edges at high frequencies, the cascade solution is constructed as a superposition of acoustic farfields emanating from an infinite number of isolated airfoils. A parametric study of noise generated by gust-cascade interaction is then carried out to assess the effectiveness of vane sweep in reducing rotor-stator interaction noise. The results of the parametric study show that, over a fairly wide range of conditions, sweep is beneficial in reducing noise levels. One conclusion of particular importance is that rotor wake twist or circumferential lean substantially influences the effectiveness of vane sweep. The orientation of the vane sweep must be chosen to enhance the natural phase lag caused by wake lean, in which case rather small sweep angles substantially reduce the noise levels.
389

NONLINEAR AERODYNAMICS OF CONICAL DELTA WINGS.

SRITHARAN, SIVAGURU SORNALINGAM. January 1982 (has links)
Steady, inviscid, supersonic flow past conical wings is studied within the context of irrotational, nonlinear theory. An efficient numerical method is developed to calculate cones of arbitrary section at incidence. The method is fully conservative and implements a body conforming mesh generator. The conical potential is assumed to have its best linear variation inside each cell; a secondary interlocking cell system is used to establish the flux balance required to conserve mass. In regions of supersonic cross flow, the discretization scheme is desymmetrized by adding the appropriate artificial viscosity in conservation form. The algorithm is nearly an order of magnitude faster than present Euler methods. It predicts known results as long as the flow Mach numbers normal to the shock waves are near 1; qualitative features, such as nodal point lift-off, are also predicted correctly. Results for circular and thin elliptic cones are shown to compare very well with calculations using Euler equations. This algorithm is then implemented in the design of conical wings to be free from shock waves terminating embedded supersonic zones adjacent to the body. This is accomplished by generating a smooth cross-flow sonic surface by using a fictitious gas law that makes the governing equation elliptic inside the cross-flow sonic surface. The shape of the wing required to provide this shock-free flow, if such a flow is consistent with the sonic surface data, is found by solving the Cauchy problem inside the sonic surface using the data on this surface and, of course, the correct gas law. This design procedure is then demonstrated using the simple case of a circular cone at angle of attack.
390

Pilot estimates of glide path and aim point during simulated landing approaches

Acree, Cecil Wallace January 1978 (has links)
Thesis. 1978. E.A.A.--Massachusetts Institute of Technology. Dept. of Aeronautics and Astronautics. / MICROFICHE COPY AVAILABLE IN ARCHIVES AND AERONAUTICS. / Bibliography: leaves 130-133. / by C.W. Acree, Jr. / E.A.A.

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