The notion of airplane stability and control being a balancing act between stability and control has been around as long as aeronautics. The Wright brothers’ first successful flights were born of the debate, and were successful at least in part because they spent considerable time teaching themselves how to control their otherwise unstable airplane. This thesis covers four aspects of handling for large transport aircraft: large size and the accompanying low frequency dynamics, the way in which lifting surfaces and control system elements are modelled in flight dynamics analyses, the cockpit feel characteristics and details of how pilots interact with them, and the dynamic instability associated with Pilot Induced Oscillations. The dynamics associated with large transport aircraft are reviewed from the perspective of pilot-in-the-loop handling qualities, including the effects of relaxing static stability in pursuit of performance. Areas in which current design requirements are incomplete are highlighted. Issues with modelling of dynamic elements which are between the pilot’s fingers and the airplane response are illuminated and recommendations are made. Cockpit feel characteristics are examined in detail, in particular, the nonlinear elements of friction and breakout forces. Three piloted simulation experiments are described and the results reviewed. Each was very different in nature, and all were designed to evaluate linear and nonlinear elements of the cockpit feel characteristics from the pilot’s point of view. These included understanding the pilot’s ability to precisely control the manipulator itself, the pilot’s ability to command the flight path, and neuro-muscular modelling to gain a deeper understanding of the range of characteristics pilots can adapt to and why. Based on the data collected and analyzed, conclusions are drawn and recommendations are made. Finally, a novel and unique PIO prediction criterion is developed, which is based on control-theoretic constructs. This criterion identifies unique signatures in the dynamic response of the airplane to predict the onset of instability.

Plasma actuators are devices which function by creating a discharge in air at atmospheric conditions. These devices have been demonstrated to effectively delay flow separation and enhance the lift- drag characteristics of wing sections. They have also been shown to have potential applications in controlling dynamic stall, flow separation control over turbine blades, flow vectoring, boundary layer manipulation and bluff body flow control. This study examines the characteristics of the plasma actuator, its working and the optimization of its configuration for its use as a lift enhancing device. A single actuator connected to a high-voltage, high-frequency power supply was studied in quiescent conditions. It was demonstrated by means of flow visualization experiments and hot-wire anemometry that the plasma actuator functions by inducing a flow, thus behaving as a source of momentum flux in any system that it is introduced into. Further, it was inferred that the flow induced is a wall jet and that the magnitude of the velocity achieved is maximum within a few millimeters of the surface of the actuator. A parametric investigation of the actuator was conducted next. The variation of the peak velocity induced in quiescent conditions with the variation of configuration parameters was studied by means of photographic studies and hot-wire anemometry. These experiments indicated that there is a strong correlation between the visible extent of the plasma along the direction of the induced _ow (plasma width) and the peak velocity achieved. The peak velocity achieved is found to increase with the increase in the plasma width as long as the discharge created is in the uniform glow discharge regime. The development of localized high intensity streamers, which destroy the uniformity of the plasma, lead to a loss in the peak velocity. Hot-wire tests indicated that the peak velocity increases with a decrease in the spanwise overlap of the electrodes, with the other parameters kept constant. Also, in the uniform glow discharge regime, the velocity increases with the increase in the thickness of the dielectric placed between the two electrodes. After a particular optimum thickness, further increase of the thickness leads to formation of streamers. The velocity increased with a decrease in streamwise overlap, with the maximum being reached for a overlap of approximately 2mm, after which it remained a constant. It was observed that the absence of overlap leads to a loss of uniformity of the discharge created. The velocity was found to be independent of the variations in the electrode widths. Particle Image Velocimetry (PIV) was conducted to study the characteristics of the jet produced. It was observed that when the actuator is switched on, a low pressure region is created near the surface of the actuator, vertically above it, leading to a flow towards this region from above the actuator. Furthermore, a vortex is shed, which is convected downstream, after which a wall jet is established close to the dielectric surface.

In this work, capturing Near-Earth Asteroids (NEAs) into Near-Earth orbits is investigated. A general optimization strategy is employed whereby a genetic algorithm is used to seed a sequential quadratic programming (SQP) method for the first step, and then nearby solutions seed further SQP runs. A large number of solutions are produced for several asteroids with varying levels of thrust, and under the effects of various perturbations. Solutions are found over a range of epochs and times of flight as opposed to many traditional methods of optimizing point solutions. This methodology proved effective, finding low-thrust capture solutions within 10% of the required Delta V for analytically estimated transfers, and matching results from other optimization programs such as MALTO. It is found that the effects of solar radiation pressure (SRP) and n-body effects do not have a significant impact on the optimized transfer costs, nor do the perturbations significantly affect the shapes and trends of the optimized solution space. <p><br></p> <p>These optimized results are then used to develop analytic models for both optimized transfer costs and flight times. These models are then used to estimate the transfer costs and flight times for all listed Near Earth Asteroids from the JPL Small Body Database. This analysis is then used to determine the nominal properties of potentially capturable asteroids. The characteristics are then related to a series of different asteroid transfer technologies, elucidating each technology's capabilities and potential capture targets. Finally, this analysis concludes with a brief roadmap of the major decisions mission designers should consider for future asteroid capture missions.</p>

<p>The long-term sustainable human exploration of Mars is approached via the design and analysis of an orbital logistics architecture as part of a robust logistics infrastructure. In this investigation, we analyze the advantages of an orbital logistics node around Mars (which we call Mars Spacedock), which plays a crucial role to support the transport of vehicles and resupply of cargo to a base on the surface. The Mars Spacedock serves as one of the many logistics nodes at different locations between Earth and Mars that support the continuous movement of crew and cargo to and from Mars for the next several decades. The need of multiple nodes at strategic locations is supported by lessons learned from terrestrial analogs of complex missions such as military, Antarctic exploration, and the International Space Station. The Mars Spacedock is envisaged to have at least aggregation, refueling, resupply and refurbishing capabilities. The stationing orbit of the Spacedock is one of the primary design drivers in determining the associated propellant requirement and surface accessibility. The stationing orbit is selected from a range of Mars orbits such that it best accommodates (delta V  cost being a major determinant) arrival from a variety of interplanetary approaches, capture into Mars orbit, deorbit and entry into Mars atmosphere, surface accessibility, launch from surface to stationing orbit, and departure to Earth. A variety of mission types are evaluated over a 15-year cycle as follows: long-stay crewed missions, short-stay crewed missions, cargo transfer missions on low-thrust and ballistic trajectories. The perturbation of orbits due to aspherical gravity of Mars and timeline of missions are found to be crucial factors in selection of orbit. The Low Mars Orbits are found to be comparable to the Highly Elliptical Mars Orbits in total delta V requirement. The optimal stationing orbit is selected by minimizing a combination of mission propellant mass and transfer time for a given set of mission parameters. The sensitivity of the optimal solution to various mission parameters (landing site latitude, propellant, refueling capability in Mars orbit, deorbit method, mission type, and frequency of different mission types) is assessed. The analysis on orbit considerations aids mission designers in selecting suitable stationing orbit for a set of mission parameters and assessing the long term impacts of mission design choices on the logistics requirements. Finally, the viability of the Spacedock is analyzed in terms of landing site accessibility, station-keeping requirement, and initial mass in cislunar staging orbit. Here also Low Mars Orbits have accessibility over a wider range of landing sites compared to 1 sol orbit. The station-keeping requirement is found to be insignificant over the scale of the missions. The Spacedock refuel capability leads to lower mass in cislunar staging orbit, about 60 Mg lower per crewed MTV mission, and compensates for the higher capture and departure delta Vs.</p> <p>   </p> <p>   A logistics architecture stationed in a strategic orbit around Mars would enable long term sustainable operations for human exploration, reduce the logistics footprint of the exploration campaigns, and aid in transitioning to an eventual permanent presence on Mars. </p>

<p>Increasingly, space flight missions are planned to traverse regions of space with complex dynamical environments influenced by multiple gravitational bodies. The nature of these systems produces motion and regions of sensitivity that are, at times, unintuitive,</p> <p>and the accumulation of trajectory dispersions from a variety of sources guarantees that spacecraft will deviate from their pre-planned trajectories in this complex environment, necessitating the use of a targeting process to generate a new feasible reference path. To ensure mission success and a robust path planning process, trajectory designers require insight into the interaction between the targeting process, the baseline trajectory, and the dynamical environment. In this investigation, the convergence behavior of these targeting processes is examined. This work summarizes a framework for characterizing and predicting the convergence behavior of perturbed targeting problems, consisting of a set of constraints, design variables, perturbation variables, and a reference solution within a dynamical system. First, this work identifies the typical features of a convergence basin and identifies a measure of worst-case performance. In the absence of an analytical method, efficient numerical discretization procedures are proposed based on the evaluation of partial derivatives at the reference solution to the perturbed targeting problem. A method is also proposed for approximating the tradespace of position and velocity perturbations that achieve reliable</p> <p>convergence toward the baseline solution. Additionally, evaluated scalar quantities are introduced to serve as predictors of the simulation-measured worst-case convergence behavior based on the local rate of growth in the constraints as well as the local relative change in the targeting-employed partial derivatives with respect to perturbations.</p> <p><br></p> <p>A variety of applications in different dynamical regions and force models are introduced to evaluate the improved discretization techniques and their correlation to the predictive metrics of convergence behavior. Segments of periodic orbits and transfer trajectories from past and planned missions are employed to evaluate the relative convergence performance across sets of candidate solutions. In the circular restricted three-body problem (CRTBP), perturbed targeting problems are formulated along a distant retrograde orbit and a near-rectilinear halo orbit (NRHO) in the Earth-Moon system. To investigate the persistence of results from the CRTBP in an ephemeris force model, a targeting problem applied to an NRHO is analyzed in both force models. Next, an L1 -to-L2 transit trajectory in the Sun-Earth system is studied to explore the effect of moving a maneuver downstream along</p> <p>a trajectory and altering the orientations of the gravitational bodies. Finally, a trans-lunar return trajectory is explored, and the convergence behavior is analyzed as the final maneuver time is varied.</p>

<p dir="ltr">The electrification of pneumatic or hydraulic system on aircraft has been shown effective in reducing the fuel burn. Recently, electrifying propulsive loads has attracted a lot of atten- tion to further improve fuel economy. This work focuses on tools to facilitate more electric aircraft at conceptual design stage, particularly assuming a turbo-generator architecture. Specifically, we develop a simulation tool, mimicking SUAVE [1], which allows mission and fuel burn analysis. Major differences from SUAVE include more detailed models of compo- nents in the electric propulsive branch and degrees of freedom to adjust the velocity profile along the entire mission. Based on the simulator, this work further proposes to leverage a gradient-free optimization technique, which optimizes the optimal velocity profile along the entire mission to minimize fuel burn. Simulation results on two aircraft designs, a con- ventional Boeing 737-800 and NASA-STARC-ABL, verify the effectiveness of the proposed tools.</p>

The interaction of an aircraft's structure and the flight dynamics can degrade the performance of a controller designed only considering the rigid body flight dynamics. These concerns are greater for the next generation adaptive controls. These interactions lead to an increase in the tracking error, instabilities in the control parameters, and significant structural excitations. To improve the understanding of these issues the interactions have been examined using simulation as well as flight testing of a subscale aircraft. The scaling required for such a subscale aircraft has also been examined. For the simulation a coordinate system where the non-linear flight dynamics are orthogonal to the linear structural dynamics was defined. The orthogonality allows the use of separates models for the aerodynamics. For the non-linear flight dynamics, preexisting table lookups with extended vortex lattice are used to determine the aerodynamic forces. Strip theory is then used to determine the smaller, but still important, unsteady aerodynamic forces due to the flexible motion. Because the orientation of the engines is dependent on the structural deformations, the propulsive force is modeled as a non-conservative follower force. The simulation of the integrated dynamics is then used to examine the effects of the aircraft flexibility and resultant ASE interactions on the performance of adaptive controls. For the scaling, the complete similitude of a flexible aircraft was examined. However, this complete similitude is unfeasible for an actual model, so partial similitude is investigated using two approaches. First, the classical approximations of the flight dynamic modes are used to reduce the order of the coupled model, and consequently the number of scaling parameters required to maintain the physics of the system. The second approach uses sensitivity of the response to errors in the aircraft's nondimensional parameters. Both methods give a consistent set of nondimensional parameters which do not have significant influence on the aeroservoelastic interaction. These parameters do not need to be scaled, thus leading to a viable scaled model. A subscale vehicle has been designed which shows significant coupling between the flight dynamics and structural dynamics. This vehicle was used to validate the results of the scaling theory. Output error system identification was used to identify a model from the flight test data. This identified model provides the frequency of the short-period mode, and the effects of the Froude number on the flexibility. / Ph. D.

The NASA Performance Adaptive Aeroelastic Wing (PAAW) project goals include significant reductions in fuel burn, emissions, and noise via efficient aeroelastic design and improvements in propulsion systems. As modern transport airplane designs become increasingly lightweight and incorporate high aspect-ratio wings, aeroservoelastic effects gain prominence in modeling and design considerations. As a result, the influence of the flight dynamics and controls on the optimal structural and aerodynamic design needs to be captured in the design process. There is an increasing interest in more integrated aircraft multidisciplinary design optimization (MDAO) processes that can bring flight control design into the early stage of an aircraft design cycle. So, in this thesis different flight dynamics modeling methodologies are presented that can be integrated within the MDAO framework. MDAO studies are conducted to maximize the controllability and observability of a UAV type aircraft using curvilinear SpaRibs and straight spars and ribs as the internal structural layout. The impulse residues and controllability Gramians are used as surrogates for the control objectives in the MDAO to maximize the controllability and observability of the aircraft. The optimal control designs are compared with those obtained using weight minimization as the design objective. It is found that using the aforementioned control objectives, the resulting aircraft design is more controllable and can be used to expand the flight envelope by up to 50% as compared to the weight minimized design. / Doctor of Philosophy / Over the last two decades, several attempts have been made towards multidisciplinary design analysis and optimization (MDAO) of flexible wings by integrating flight control laws in the wing design so that the aircraft will have sufficient control authority across different flying conditions. However, most of the studies have been restricted to the wing design only using a predefined control architecture approach, which would be very difficult to implement at the conceptual design stage. There is a need for an approach that would be faster and more practical. Including control surface and control law designs at the conceptual design stage is becoming increasingly important, due to the complexity of both the aircraft control laws and that of the actuation and sensing, and the enhanced wing flexibility of future transport aircraft. A key question that arises is, can one design an aircraft that is more controllable and observable? So, in this thesis, a more fundamental approach, in which the internal structural layout of the aircraft is optimized to design an aircraft that is more controllable, is presented and implemented. The approach uses the fundamentals of linear systems theory for maximizing the controllability and observability of the aircraft using an MDAO framework.

The study of the motion of manoeuvring aircraft has traditionally considered the aircraft to be rigid. This simplifying assumption has been shown to give quite accurate results for the flight dynamics of many aircraft types. As modern transport aircraft have developed however, there has been a marked increase in the size and weight of these aircraft. This trend is likely to continue with the development of future blended-wing-body and supersonic transport aircraft. This increase in size and weight has brought about a unique set of aeroelastic and handling quality issues. The aerodynamic forces and moments acting on an aeroplane have traditionally been represented using the aerodynamic derivative approach. It has been shown that this quasisteady aerodynamic model inadequately predicts the aircraft’s stability characteristics, and that the inclusion of unsteady aerodynamics “greatly improves the fidelity” of aircraft models. This thesis thus presents a novel numerical simulation of an aeroelastic aeroplane for real-time analysis. The model is built around the standard six degree-of-freedom equations of motion for a rigid aeroplane using the mean-axes system, and includes unsteady aerodynamics and structural dynamics. This is suitable for pilot-in-the-loop simulation, handling qualities and flight loads analysis, and control law development. The dynamics of the structure are modelled as a set of normal modes, and the equations of motion are realised in state-space form. The unsteady aerodynamic forces acting on the aeroplane are described by an indicial state-space model, including unsteady tailplane downwash and compressibility effects. An implementation of the model is presented in the MATLAB/ Simulink environment. The interaction between the flight control system, the aeroelastic system and the rigidbody motion of the aeroplane can result in degraded handling qualities, excessive actuator control, and fatigue problems. The introduction of load alleviation systems for the management of loads due to manoeuvres and gusts is also likely to result in the handling qualities of the aeroplane being degraded. This thesis presents a number of studies into the impact of structural dynamics, unsteady aerodynamics, and load alleviation on the handling qualities of a flexible civil transport aeroplane. The handling qualities of the aeroplane are assessed against a number of different handling qualities criteria and flying specifications, including the Neal-Smith, Bandwidth, and CAP criterion. It is shown that aeroelastic effects alter the longitudinal and lateral-directional characteristics of the aeroplane, resulting in degraded handling qualities. Manoeuvre and gust load alleviation are similarly found to degrade handling qualities, while active mode control is shown to offer the possibility of improved handling qualities.

This thesis considers the problem of designing gain scheduled flight control system (FCS) for large transport aircraft that satisfy handling qualities criteria. The goal is to design a set of local Linear Time Invariant (LTI) controllers to cover the wide non- linear aircraft operation flight envelope from the viewpoint of the handling qualities assessment. The global gain scheduler is then designed that interpolates between the gains of the local controllers in order to transfer smoothly between different equilibrium points, and more importantly to satisfy the handling qualities over the entire flight envelope. The mathematical model of the Boeing 747-100/200 aircraft is selected for the purpose of the flight controller design and handling qualities as- sessment. In order to achieve an attitude hold characteristic, and also improve the dynamic tracking behavior of the aircraft, longitudinal pitch Rate Command-Attitude Hold (RCAH) controllers are designed as the local flight controllers at the specific equilib- rium points in the flight envelope by means of a state space pole placement design procedure. The handling qualities assessment of the aircraft is presented, based on which the scheduler is designed. A number of existing criteria are employed to assess the han- dling qualities of the aircraft, including the Control Anticipation Parameter (CAP), Neal and Smith, and C∗ criteria. The gain scheduled flight controller is found to have satisfactory handling qualities. The global gain scheduler is designed by interpolating the gains of the local flight controllers in order to transfer smoothly between different equilibrium points, and more importantly to satisfy the handling qualities over the flight envelope. The main contribution of this research is the combination of the gain scheduling technique based on the local controller design approach and handling qualities as- sessment. The controllers are designed based at a number of operating points and the interpolation between them (scheduling) takes place through the scheduling scheme functions. The aircraft augmented with gain-scheduled controller performs satisfactorily and meets the requirement of handling qualities. Moreover, the per- formance using the gain-scheduled controller is considerably improved compared to the performance using the fixed one.