Aircraft Cruise Performance is examined by using energy-state modelling to investigate fuel-range optimal trajectories. Chattering controls are considered appropriate when the hodograph is non-convex. Classical steady-state cruise, simple chattering-cruise and the extended chattering-cruise models are studied as constrained parameter-optimization problems. The term "extended chattering" refers to vehicle system modelling extended to maintain vertical equilibrium only on the average. Numerical solution is obtained using a variable-metric gradient-protection algorithm and computational results are presented for three different aircraft. This study shows that simple chattering cruise for certain specific energies can result in substantial fuel savings over classical steady-state cruise. However extended chattering cruise results in only marginal fuel savings when compared to simple chattering cruise. / Master of Science

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

This thesis presents an effective control design methodology using a one-step-ahead prediction adaptive control law and an adaptive control law based on a Lyapunov function. These control law were applied to a highly maneuverable high performance aircraft, in particular, a modified F/A-18. An adaptive controller is developed to maneuver an aircraft at a high angle of attack even if the aircraft is required to fly over a highly nonlinear flight regime. The adaptive controller presented in this thesis is based on linear, bilinear, and nonlinear prediction models with input constraints. It is shown that the linear, bilinear, and nonlinear adaptive controllers can be constructed to minimize the given cost function or Lyapunov function with respect to the control input at each step. The control is calculated such that the system follows the reference trajectory, and such that control signal remains within its constraints. From several simulation results, the nonlinear controller is controller is better than the linear controller. A nonlinear adaptive control law based on a Lyapunov function is designed such that control inputs are smoother than for the one-step-ahead prediction adaptive controller. / Graduation date: 1994

In this thesis, A nonlinear methodology for the control of the highly maneuverable, high performance aircraft HARV (F-18) is studied by using sliding-mode control (SMC). This control law, which takes a continuous function when the input constraints are not considered, satisfies the reachability condition by which concerned states are driven to their sliding surfaces. Especially, this SMC is generalized for the so-called (square) uncoupled multi-input multi-output (MIMO) system by the use of error dynamics and applied directly to the nonlinear aircraft system without linearizing the system. For the practical application of the SMC to aircraft with input constraints, two control schemes are used, considering that variations of pitch rate q directly affect variations of the angle of attack. The first scheme consists of an ��-q control for the fast response. The second scheme consists an a-control for making the output approach its sliding surface slowly by setting boundary layers and adjusting reachable speeds to the sliding surface. Robustness to parameter uncertainties and disturbances is also studied for the SMC. Especially, when the effect of parameter uncertainties is severe, then multiple boundary layers are set in the neighborhood of sliding surface, in which different reachable speeds to the sliding surface are used to sustain the concerned state within the boundary layer and to reduce the effect of chattering. / Graduation date: 1994

An adaptive pitch axis autopilot design procedure is presented. The design procedure is applicable to both stable and unstable pitch axis models and to those having nonminimum phase. The design approach assumes the adaptive autopilot is activated after achieving level flight. It is shown a rate-feedback compensator can be designed to ensure stable level flight pitch axis operation for the entire desired flight regime. The adaptive control loop design utilizes a pole-placement algorithm. The closed-loop characteristic polynomial is designed to have dominant poles of that of an ideal second order system to obtain the desired transient response. The identification of the system uses a modified least-squares algorithm with a variable forgetting factor. The nonlinear pitch axis model is used in simulations to evaluate the design. Command response tests include the step response and the ramp command response. Simulation results indicate that the adaptive pitch axis autopilot is capable of tracking altitude commands after activation. The closed-loop system response is close to that of the ideal second order system having the dominant poles. / Graduation date: 1991

No description available.

All aircraft must meet controllability requirements to be certified for commercial use or adopted by the mi1itary. Aircraft maneuverability is often 1imited by control authority. Thus, it is essential for designers to evaluate a candidate concept's control authority early in the conceptual design phase. In this thesis, a methodology for rapid control power evaluation of preliminary design configurations against requirements at the key flight conditions is established. First, a collection of critical flight conditions to be considered using this methodology is identified. To examine a variety of aircraft configurations and accelerate the process of estimating stability and control derivatives, a FORTRAN program using the Vortex-Lattice Method was written to estimate subsonic, low angle-of-attack aerodynamics. Then, a spreadsheet processes the aerodynamic data to check whether the design configuration possesses adequate control power to satisfy the requirements of the critical flight conditions. / Master of Science

A digital model reference scheme is presented for the control of the longitudinal modes of an EBF-STOL aircraft. The method used is based on Liapunov's direct method and is used to update feedback and feedthrough matrices. An equilibrium point of the EBF-STOL is chosen about which to constrain performance to follow that of a linearized decoupled DC-9 model. The change of control matrix elements is controlled by (1) an error vector equal to the difference between state variables of the model and those of the STOL, (2) the present values of the pilot inputs and STOL state variables, (3) a P matrix obtained from the equation A_mᵀPA_m - P = - I where z(k + 1) = A_m z(k) + B_m u_p is the equation of motion for the model, (4) selected gains, and (5) a matrix B₋ᵤʳ⁻¹ whose elements are determined by the inverse of an n x n partition of the linearized input matrix B₋ᵤ of the STOL. A method for smoothing control action is discussed. / Master of Science

The problem of determining cruise-dash trajectories is examined for the case of time-fuel optimization using a linear combination of time and fuel as the performance index. These trajectories consist of a transient arc followed by a steady-state arc. For cases where the steady-state arc is flown with full throttle the associated skeletal transient trajectories are also flown with full throttle, and approach the cruise-dash points monotonically in an asymptotic fashion. When the steady-state arc is flown at an intermediate throttle setting, the transient trajectories follow a singular control law and exhibit a complex structure that is different from the full-throttle transients. Singular transients in the vicinity of singular cruise-dash points are confined to a bounded singular surface. In state-space these trajectories trace out asymptotic spirals on the singular surface as they approach the steady-state arc. If the initial operating point lies outside the singular surface, then the transient trajectories are composites consisting of a full-throttle or zero-throttle segment flown until the singular surface is met, followed by a singular segment that fairs into the cruise-dash point. Addressing the question of optimality of the steady-state arc, it was found that although steady-state cruise fails a Jacobi-type condition, steady-state cruise-dash can satisfy this condition if the emphasis on time is sufficiently large. The outcome of the Jacobi-type test appears to be connected with the eigenstructure of the linearized state-adjoint system. / PH. D.

Thesis. 1975. M.S.--Massachusetts Institute of Technology. Dept. of Electrical Engineering and Computer Science. / Bibliography: leaves 258-259. / by Christopher S. Greene. / M.S.