Thesis (Ph. D.)--University of Washington, 1997. / Vita. Includes bibliographical references (leaves [165]-173).

M.Ing. (Electrical and Electronic Engineering) / This treatise presents an investigation into the application of multivariable frequency domain techniques in the modelling and control of a helicopter aircraft in forward flight. The presentation is structured in the following sectioned format: I Hypotheses are stated which deal with the use of linear, multivariable, frequency domain theory in the modelling and control of helicopter aircraft. II The stated hypotheses are investigated by the application of relevant theories and techniques to a reference case plant - a single rotor helicopter in forward flight. III Conclusions drawn from the results are used to assess the validity of the hypotheses. The subject matter of the presentation may be summarized as follows: The hypotheses are initially placed in perspective by a discussion of the incentives for their formulation. In essence, the hypotheses state that helicopter dynamics, in a multivariable systems characterization, can be modelled and an appropriate flight control system designed by the use of linear frequency domain theory. The plant in reference to which the hypotheses are investigated is a single rotor utility helicopter - the Aerospatiale Alouette III. A single flight condition - a typical cruising condition - is considered. A comprehensive, nonlinear digital computer simulation of the aircraft is used as a substitute for the actual plant in the execution of the modelling and control design processes. The plant is modelled in terms of a linear model structure, in the form of the frequency response function, by linearization of its highly nonlinear dynamics about an operating point (datum flight condition). The frequency response function model parameters are identified by power spectral density analysis procedures. This method, based on random signal excitation of the plant, provides a valuable quantitative measure of the accuracy of the linearization performed in the identification. The measure, the coherence function, is used as a criterion for the robustness required of a control system of which the design is based on a linear model of a nonlinear plant.

A method for computing the skin-stiffener interface stresses in stiffened composite panels is developed. Both geometrically linear and nonlinear analyses are considered. Particular attention is given to the flange termination region where stresses are expected to exhibit unbounded characteristics. The method is based on a finite-element analysis and an elasticity solution. The finite-element analysis is standard, while the elasticity solution is based on an eigenvalue expansion of the stress functions. The eigenvalue expansion is assumed to be valid in the local flange termination region and is coupled with the finite-element analysis using collocation of stresses on the local region boundaries. In the first part of the investigation the accuracy and convergence of the local elasticity solution are assessed using a geometrically linear analysis. It is found that the finite-element/local elasticity solution scheme produce a very accurate interface stress representation in the local flange termination region. The use of 10 to 15 eigenvalues, in the eigenvalue expansion series, and 100 collocation points results in a converged local elasticity solution. In the second part of the investigation, the local elasticity solution is extended to include geometric nonlinearities. Using this analysis procedure, the influence of geometric nonlinearities on skin-stiffener interface stresses is evaluated. It is found that in flexible stiffened skin structures, which exhibit out-of-plane deformation on the order of 2 to 4 times the skin thickness, inclusion of geometrically nonlinear effects in the calculation of interface stresses is very important. Thus, the use of a geometrically linear analysis, rather than a nonlinear analysis, can lead to considerable error in the computation of the interface stresses. Finally, using the analytical tool developed in this investigation, it is possible to study the influence of stiffener parameters on the state of interface stresses. / Ph. D.

It is the intent of this paper to investigate how the convergence rate of a fully stressed design method using a stress-ratio algorithm may be affected by manipulating various parameters. The parameters studied are: (1) selection of a design variable, (2) relationships between the variables of moment of inertia, area, and section modulus, (3) use of overrelaxation, (4) tolerance on scale factors, (5) initial input, and (6) number and type of loading conditions. Design examples illustrating the discussion of the various parameters are given. The optimality of a fully stressed design approach and its use in industry is presented in a literature review, and discussion of the stress-ratio algorithm and the feasibility (convergence) of the fully stressed design method is also included. / Master of Science

As the UH-1 aircraft continue to age, there is growing concern for their structural integrity. With corrosion damage becoming a bigger part of the sustainment picture with increasing maintenance burden and cost, it is becoming increasingly important for corrosion management to be updated with more advanced techniques. The current find-and-fix technique for handling corrosion has many shortfalls, spurring the recent interest in early detection through structural health monitoring. This condition based technique is becoming more prevalent and is recognized for the potential to greatly reduce maintenance cost. Through corrosion monitoring, structural and environmental conditions can be closely observed, preventing excessive maintenance action and saving cost. Searches for corrosion monitoring system designs revealed several commercial companies with prototype systems installed on commercial aircraft, however, details on system design and data analysis were scarce. This study attempted to bridge the gap in literature by providing insight into the development of a corrosion damage prediction model and the design of a corrosion monitoring system. This study attempted to use aircraft maintenance data to make prediction models for determining what corrosion damage an aircraft can expect, given varying operating conditions. Although a reliable prediction model could not be created, trends observed in the data were still valuable for identifying problematic areas of the aircraft. In order to create reliable models, more accurate corrosion data is needed. This can be accomplished through the implementation of a corrosion monitoring system. A custom corrosion monitoring system was designed for the UH-1 aircraft. Commercial off-the-shelf products were fit to the design and a benefits-to-cost analysis was performed for the monitoring system, evaluating the system based on criteria developed from user requirements. The system proved to meet and exceed expectation, making it an ideal choice for the UH-1 aircraft.

M. Tech. Mechanical Engineering / The increased use of composite materials in different industries has led to the realization of some of its benefits and disadvantages. One of the major problems, however, is the availability of biaxial test data for different composite materials. This is because structures during application face multi-axial stress states examples of such stress state scenarios include wind turbine blades and pressure vessels. This has also led to diverse range of test methods and material compositions such as combining different fibbers. The material used in this work is polymer matrix honeycomb sandwich panels. Sandwich panels are fabricated by attaching two thin stiff skins of fiber glass or carbon fiber reinforced laminates to a lightweight core. This work addresses some of the major advantages and disadvantages of this testing method. The main objective of this study is to develop a repeatable, cost effective and time efficient method for multi-axial testing of sandwich panels using existing resources.

This minor thesis examines whether or not the Australian Defence Force (ADF) needs an indigenous aircraft structural testing program for military aircraft. Aircraft structural testing is a major test and evaluation activity currently undertaken by the DSTO to provide structural integrity information about particular aircraft for the ADF. The ADF has required this testing to be undertaken because, in general terms, testing undertaken by an aircraft's manufacturer has not been sufficient to meet the needs of the ADF. The research was divided into five sub-problems, which examined the need for aircraft structural testing, the management of these needs, what structural testing is required, how the testing should be managed and how the testing should be conducted. The research was undertaken by conducting five case studies into past, current and future aircraft structural tests being undertaken by the DSTO. The cases examined were the F/A-18 International Follow-On Structural Test Project, the P-3 service life assessment program, the F-111 wing test, the Hawk lead-in fighter full-scale fatigue test and the potential test program for the joint strike fighter (JSF). The basis of the case studies was the conduct of a series of interviews of the DSTO staff relevant to each case being examined. The information gathered through the interviews was supported by material gathered through making observations and reviewing documentation relevant to each case. The research concluded that an indigenous aircraft testing program is required by the ADF in order to maintain the structural integrity of its aircraft fleet. This minor thesis makes some recommendations for the way aircraft structural testing is conducted by the ADF, and the way general management of T & E in the Department of Defence could be improved. / Thesis (MEng(TestandEvaluation))--University of South Australia, 2004

This minor thesis examines whether or not the Australian Defence Force (ADF) needs an indigenous aircraft structural testing program for military aircraft. Aircraft structural testing is a major test and evaluation activity currently undertaken by the DSTO to provide structural integrity information about particular aircraft for the ADF. The ADF has required this testing to be undertaken because, in general terms, testing undertaken by an aircraft's manufacturer has not been sufficient to meet the needs of the ADF. The research was divided into five sub-problems, which examined the need for aircraft structural testing, the management of these needs, what structural testing is required, how the testing should be managed and how the testing should be conducted. The research was undertaken by conducting five case studies into past, current and future aircraft structural tests being undertaken by the DSTO. The cases examined were the F/A-18 International Follow-On Structural Test Project, the P-3 service life assessment program, the F-111 wing test, the Hawk lead-in fighter full-scale fatigue test and the potential test program for the joint strike fighter (JSF). The basis of the case studies was the conduct of a series of interviews of the DSTO staff relevant to each case being examined. The information gathered through the interviews was supported by material gathered through making observations and reviewing documentation relevant to each case. The research concluded that an indigenous aircraft testing program is required by the ADF in order to maintain the structural integrity of its aircraft fleet. This minor thesis makes some recommendations for the way aircraft structural testing is conducted by the ADF, and the way general management of T & E in the Department of Defence could be improved. / Thesis (MEng(TestandEvaluation))--University of South Australia, 2004

Thesis (M.App.Sc.) - Carleton University, 2002. / Includes bibliographical references (p. 143-147). Also available in electronic format on the Internet.

Thesis (Ph.D.) - Carleton University, 2002. / Includes bibliographical references. Also available in electronic format on the Internet.