This thesis compares the results of two of the more popular flexible aircraft modeling formulations, the mean-axes method and the fixed-axes method, for application in real-time motion simulators. First, the time-domain equations of motion for an elastic body using the fixed-axes are derived. Subsequently, the mean-axes equations are derived by making a few assumptions from the fixed-axis equations. The two formulations are then implemented for a scaled-up beam model of a Cessna Citation aircraft, with the deformations represented by the modal expansion of the whole aircraft from their respective constrained and free-free finite element solutions. Time-domain results, consisting of the acceleration, velocity, and attitude of a point on the aircraft body, are obtained in both models at two beam-stiffness configurations using a quasi-steady aerodynamic model for a single maneuver at one flight condition. The two methods produced similar results with the fixed-axes formulation producing slightly more accurate results.

This thesis compares the results of two of the more popular flexible aircraft modeling formulations, the mean-axes method and the fixed-axes method, for application in real-time motion simulators. First, the time-domain equations of motion for an elastic body using the fixed-axes are derived. Subsequently, the mean-axes equations are derived by making a few assumptions from the fixed-axis equations. The two formulations are then implemented for a scaled-up beam model of a Cessna Citation aircraft, with the deformations represented by the modal expansion of the whole aircraft from their respective constrained and free-free finite element solutions. Time-domain results, consisting of the acceleration, velocity, and attitude of a point on the aircraft body, are obtained in both models at two beam-stiffness configurations using a quasi-steady aerodynamic model for a single maneuver at one flight condition. The two methods produced similar results with the fixed-axes formulation producing slightly more accurate results.

This thesis examines the relative importance of the inertial and geometric nonlinearities in modelling the dynamics of a flexible aircraft. Inertial nonlinearities are derived by employing an exact definition of the velocity distribution and lead to coupling between the rigid body and elastic motions. The geometric nonlinearities are obtained by applying nonlinear theory of elasticity to the deformations. Peters' finite state unsteady aerodynamic model is used to evaluate the aerodynamic forces. Three approximate models obtained by excluding certain combinations of nonlinear terms are compared with that of the complete dynamics equations to obtain an indication of which terms are required for an accurate representation of the flexible aircraft behavior. A generic business jet model is used for the analysis. The results indicate that the nonlinear terms have a significant effect for more flexible aircraft, especially the geometric nonlinearities which leads to increased damping in the dynamics.

This thesis examines the relative importance of the inertial and geometric nonlinearities in modelling the dynamics of a flexible aircraft. Inertial nonlinearities are derived by employing an exact definition of the velocity distribution and lead to coupling between the rigid body and elastic motions. The geometric nonlinearities are obtained by applying nonlinear theory of elasticity to the deformations. Peters' finite state unsteady aerodynamic model is used to evaluate the aerodynamic forces. Three approximate models obtained by excluding certain combinations of nonlinear terms are compared with that of the complete dynamics equations to obtain an indication of which terms are required for an accurate representation of the flexible aircraft behavior. A generic business jet model is used for the analysis. The results indicate that the nonlinear terms have a significant effect for more flexible aircraft, especially the geometric nonlinearities which leads to increased damping in the dynamics.

A multidisciplinary design optimization framework is developed that integrates control system design with aerostructural design for a highly-deformable wing. The objective of this framework is to surpass the existing aircraft endurance limits through the use of an active load alleviation system designed concurrently with the rest of the aircraft. The novelty of this work is two fold. First, a unified dynamics framework is developed to represent the full six-degree-of-freedom rigid-body along with the structural dynamics. It allows for an integrated control design to account for both manoeuvrability (flying quality) and aeroelasticity criteria simultaneously. Secondly, by synthesizing the aircraft control system along with the structural sizing and aerodynamic shape design, the final design has the potential to exploit synergies among the three disciplines and yield higher performing aircraft. A co-rotational structural framework featuring Euler--Bernoulli beam elements is developed to capture the wing's nonlinear deformations under the effect of aerodynamic and inertial loadings. In this work, a three-dimensional aerodynamic panel code, capable of calculating both steady and unsteady loadings is used. Two different control methods, a model predictive controller (MPC) and a 2-DOF mixed-norm robust controller, are considered in this work to control a highly flexible aircraft. Both control techniques offer unique advantages that make them promising for controlling a highly flexible aircraft. The control system works towards executing time-dependent manoeuvres along with performing gust/manoeuvre load alleviation. The developed framework is investigated for demonstration in two design cases: one in which the control system simply worked towards achieving or maintaining a target altitude, and another where the control system is also performing load alleviation. The use of the active load alleviation system results in a significant improvement in the aircraft performance relative to the optimum result without load alleviation. The results show that the inclusion of control system discipline along with other disciplines at early stages of aircraft design improves aircraft performance. It is also shown that structural stresses due to gust excitations can be better controlled by the use of active structural control systems which can improve the fatigue life of the structure.

A multidisciplinary design optimization framework is developed that integrates control system design with aerostructural design for a highly-deformable wing. The objective of this framework is to surpass the existing aircraft endurance limits through the use of an active load alleviation system designed concurrently with the rest of the aircraft. The novelty of this work is two fold. First, a unified dynamics framework is developed to represent the full six-degree-of-freedom rigid-body along with the structural dynamics. It allows for an integrated control design to account for both manoeuvrability (flying quality) and aeroelasticity criteria simultaneously. Secondly, by synthesizing the aircraft control system along with the structural sizing and aerodynamic shape design, the final design has the potential to exploit synergies among the three disciplines and yield higher performing aircraft. A co-rotational structural framework featuring Euler--Bernoulli beam elements is developed to capture the wing's nonlinear deformations under the effect of aerodynamic and inertial loadings. In this work, a three-dimensional aerodynamic panel code, capable of calculating both steady and unsteady loadings is used. Two different control methods, a model predictive controller (MPC) and a 2-DOF mixed-norm robust controller, are considered in this work to control a highly flexible aircraft. Both control techniques offer unique advantages that make them promising for controlling a highly flexible aircraft. The control system works towards executing time-dependent manoeuvres along with performing gust/manoeuvre load alleviation. The developed framework is investigated for demonstration in two design cases: one in which the control system simply worked towards achieving or maintaining a target altitude, and another where the control system is also performing load alleviation. The use of the active load alleviation system results in a significant improvement in the aircraft performance relative to the optimum result without load alleviation. The results show that the inclusion of control system discipline along with other disciplines at early stages of aircraft design improves aircraft performance. It is also shown that structural stresses due to gust excitations can be better controlled by the use of active structural control systems which can improve the fatigue life of the structure.

This dissertation integrates in a single mathematical formulation the disciplines pertinent to the flight of flexible aircraft, namely, analytical dynamics, structural dynamics, aerodynamics and controls. The unified formulation is based on fundamental principles and incorporates in a natural manner both rigid body motions of the aircraft as a whole and elastic deformations of the flexible components (fuselage, wing and empennage), as well as the aerodynamic, propulsion, gravity and control forces. The aircraft motion is described in terms of three translations (forward motion, sideslip and plunge) and three rotations (roll, pitch and yaw) of a reference frame attached to the undeformed fuselage, and acting as aircraft body axes, and elastic displacements of each of the flexible components relative to corresponding body axes. The mathematical formulation consists of six ordinary differential equations for the rigid body motions and one set of ordinary differential equations for each elastic displacement. A perturbation approach permits division of the problem into a nonlinear "zero-order Problem" for the rigid body motions, corresponding to flight dynamics, and a linear "first-order problem" for the elastic deformations and perturbations in the rigid body translations and rotations, corresponding to "extended aeroelasticity." Due to computational speed advantages, the aerodynamic forces are derived by means of strip theory. The control forces for the flight dynamics problem are obtained by an "inverse" process. On the other hand, the feedback control forces for the extended aeroelasticity problem are derived by means of LQG theory. A numerical example corresponding to steady level flight and steady level turn maneuver is included. / Ph. D.

APSDS is a layered elastic tool for aircraft pavement thickness determination developed and distributed by Mincad Systems and based on the sister software Circly. As aircraft pavement thickness determination remains an empirical science, mechanistic-empirical design tools such as APSDS require calibration to full scale pavement performance, via the S77-1 curve. APSDS provides the unique advantage over other tools that it models all the aircraft in all their wandering positions, negating the need for designers to use pass to cover ratios and acknowledging that different aircraft have their wheels located at difference distances from the aircraft centerline. APSDS requires a range of input parameters to be entered, including subgrade modulus, aircraft types, masses and passes and a pavement structure. A pavement thickness is then returned which has 50% design reliability. Greater levels of reliability are obtained by conservative selection of input values. Whilst most input parameters have a linear influence on pavement thickness, subgrade modulus changes have a greater influence at lower values and less influence at higher values. When selecting input values, designers should concentrate their efforts on subgrade modulus and aircraft mass as these have the greatest influence on the required pavement thickness. Presumptive or standard values are generally acceptable for the less influential parameters. S77-1 pavement thicknesses are of a standard composition with only the subbase thickness varying. Non-standard pavement structures are determined using the principle of material equivalence and the FAA provides range of material equivalence factors, of which the mid-range values are most commonly used. APSDS allows direct modelling of non-standard pavement structures. By comparing different APSDS pavements of equal structural capacity, implied material equivalences can be calculated. These APSDS implied material equivalences lie at the lower end of the ranges published by FAA. In order to obtain consistence between APSDS and the FAA guidance, the following material equivalence values are recommended: * Asphalt for Crushed Rock. 1.3. * Crushed Rock for Uncrushed Gravel. 1.2. * Asphalt for Uncrushed Gravel. 1.6. Proof rolling regimes remain an important part of the design and construction of flexible aircraft pavements. Historically, designers relied on Bousinesq's equation and the assumption of point loads on semi-finite homogenous materials to determine proof rolling regimes using stress as the indicator of damage. The ability of APSDS to generate stress, strain and deflection at any depth and any location across the pavement allows these historical assumptions to be tested. As the design of a proof rolling regime is one of comparing damage indicators modelled under aircraft loads to those under heavy roller loads, the historical simplifications are generally valid for practical design scenarios. Where project specific data is required, APSDS can readily calculate stresses induced by proof rollers and aircraft at any location and depth for comparison. APSDS is a leading tool for flexible aircraft pavement thickness determination due to its flexibility, transparency and being free from bias. However, the following possible areas for improvement are considered worthy of future research and development: * Improvements to the user interface. * Ability to model aircraft masses as frequency distributions. * Ability to copy stress with depth data to Excel(tm) spreadsheets. * Ability to perform parametric runs. * Inclusion of a reliability based design module.