The application of quadratic optimal cooperative control synthesis to a CH-47 helicopter

A control-system design method, Quadratic Optimal Cooperative Control Synthesis (CCS), is applied to the design of a Stability and Control Augmentation Systems (SCAS). The CCS design method is different from other design methods in that it does not require detailed a priori design criteria, but instead relies on an explicit optimal pilot-model to create desired performance. The design model, which was developed previously for fixed-wing aircraft, is simplified and modified for application to a Boeing Vertol CH-47 helicopter. Two SCAS designs are developed using the CCS design methodology. The resulting CCS designs are then compared with designs obtained using classical/frequency-domain methods and Linear Quadratic Regulator (LQR) theory in a piloted fixed-base simulation. Results indicate that the CCS method, with slight modifications, can be used to produce controller designs which compare favorably with the frequency-domain approach

Output model-following control synthesis for an oblique-wing aircraft

Recent interest in oblique-wing aircraft has focused on the potential aerodynamic performance advantage of a variable-skew oblique wing over a conventional or symmetric sweep wing. Unfortunately, the resulting asymmetric configuration has significant aerodynamic and inertial cross-coupling between the aircraft longitudinal and lateral-directional axes. Presented here is a decoupling control law synthesis technique that integrates stability augmentation, decoupling, and the direct incorporation of desired handling qualities into an output feedback controller. The proposed design technique uses linear quadratic regulator concepts in the framework of explicit model following. The output feedback strategy used is a suboptimal projection from the state space to the output space. Dynamics are then introduced into the controller to improve steady-state performance and increase system robustness. Closed-loop performance is shown by application of the control laws to the linearized equations of motion and nonlinear simulation of an oblique-wing aircraft

Autonomous take-off and landing of a tethered aircraft: a simulation study

The problem of autonomous launch and landing of a tethered rigid aircraft for airborne wind energy generation is addressed. The system operates with ground-based power conversion and pumping cycles, where the tether is repeatedly reeled in and out of a winch installed on the ground and linked to an electric motor/generator. In order to accelerate the aircraft to take-off speed, the ground station is augmented with a linear motion system composed by a slide translating on rails and controlled by a second motor. An onboard propeller is used to sustain the forward velocity during the ascend of the aircraft. During landing, a slight tension on the line is kept, while the onboard control surfaces are used to align the aircraft with the rails and to land again on them. A model-based, decentralized control approach is proposed, capable to carry out a full cycle of launch, low-tension flight, and landing again on the rails. The derived controller is tested via numerical simulations with a realistic dynamical model of the system, in presence of different wind speeds and turbulence, and its performance in terms of landing accuracy is assessed. This study is part of a project aimed to experimentally verify the launch and landing approach on a small-scale prototype.Comment: This is the longer version of a paper submitted to the 2016 American Control Conference 2016, with more details on the simulation parameter

Cooperative distributed LQR control for longitudinal flight of a formation of non-identical low-speed experimental UAV's

In this paper, an established distributed LQR control methodology applied to identical linear systems is extended to control arbitrary formations of non-identical UAV's. The nonlinear model of a low-speed experimental UAV known as X-RAE1 is utilized for simulation purposes. The formation is composed of four dynamically decoupled X-RAE1 which differ in their masses and their products of inertia about the xz plane. In order to design linear controllers the nonlinear models are linearized for horizontal flight conditions at constant velocity. State-feedback, input and similarity transformations are applied to solve model-matching type problems and compensate for the mismatch in the linearized models due to mass and symmetry discrepancies among the X-RAE1 models. It is shown that the method is based on the controllability indices of the linearized models. Distributed LQR control employed in networks of identical linear systems is appropriately adjusted and applied to the formation of the nonidentical UAV's. The applicability of the approach is illustrated via numerous simulation results

Helicopter flight modeling and robust autonomous control with uncertain dynamics

Helicopter flight control has gained greater visibility in the last decades due to its characteristics. It is a task of high difficulty due to the system being changeable. Helicopter control is the junction of several variables such as the flying qualities and performance, skill of the pilot, weather conditions, etc. For a given configuration of a helicopter (military, civil, SAR, etc.), we need precise control parameters, and is in flying qualities that we focus our attention. This paper is then dedicated to autonomous modeling and control of conventional rotarywing platforms, so that there is a good balance between robustness and performance of the system, so that when there are disturbances, it can stabilize more quickly and effectively and, to verify this, a particular helicopter is used for the validation of the methods elaborated: DRA research ZD559 Lynx (Lynx Mk7), still serving the UK Army Air Corps. In helicopter flight dynamics, will be focused autorotation phase and different flight speed phase and two specific control systems will be compared: normal LQR and robust LQR. Is then designed the controllers previously mentioned, comparing their results through the dynamic model already linearized in order to verify which one is most appropriate for the platform. When there is a change in the balance of the aircraft, it can return to the same position as quickly as possible. By the obtained results, it’s verified in both cases, success in stabilizing the aircraft and controlling it’s trajectory given different reference speeds, a controller is most evident than the other.O controlo de helicópteros tem vindo a adquirir nas últimas décadas maior visibilidade devido à característica desta aeronave. É uma tarefa de elevada dificuldade devido ao próprio sistema ser, já por si, inconstante. O controlo de helicópteros é a junção de várias variáveis como as qualidades de voo e performance, a perícia do piloto, condições atmosféricas, etc. Seja qual for a missão do helicóptero (militar, civil, SAR, etc), precisaremos de parâmetros de controlo precisos, e é nas qualidades de voo que iremos focar a nossa atenção. O presente trabalho é dedicado então à modelação e controlo autónomo de plataformas de asa rotativa convencional, de forma a que haja uma boa relação entre robustez e performance do sistema para que, quando hajam perturbações, o sistema possa estabilizar eficazmente e, para se verificar isso mesmo, um helicóptero específico é utilizado para a validação dos métodos elaborados: DRA research Lynx ZD559 (Lynx Mk7), ainda ao serviço do UK Army Air Corps. Na dinâmica do voo do helicóptero, irá ser focada a fase de autorrotação e fase de voo para diferentes velocidades e dois sistemas de controlo irão ser comparados: LQR normal e LQR robusto. Projectou-se então os controladores referidos anteriormente comparando os seus resultados através da dinâmica e navegação do modelo já linearizado, para se verificar qual dos dois será mais apropriado para que a plataforma, no momento em que existe uma alteração na dinâmica da aeronave, possa regressar a essa mesma posição o mais brevemente possivel. Através então dos resultados obtidos, verifica-se em ambos os casos, o sucesso em estabilizar a aeronave e controlar o voo dada uma determinada referência para diferentes velocidades, mas um controlador evidencia-se mais que o outro

Helicopter mathematical models and control law development for handling qualities research

Progress made in joint NASA/Army research concerning rotorcraft flight-dynamics modeling, design methodologies for rotorcraft flight-control laws, and rotorcraft parameter identification is reviewed. Research into these interactive disciplines is needed to develop the analytical tools necessary to conduct flying qualities investigations using both the ground-based and in-flight simulators, and to permit an efficient means of performing flight test evaluation of rotorcraft flying qualities for specification compliance. The need for the research is particularly acute for rotorcraft because of their mathematical complexity, high order dynamic characteristics, and demanding mission requirements. The research in rotorcraft flight-dynamics modeling is pursued along two general directions: generic nonlinear models and nonlinear models for specific rotorcraft. In addition, linear models are generated that extend their utilization from 1-g flight to high-g maneuvers and expand their frequency range of validity for the design analysis of high-gain flight control systems. A variety of methods ranging from classical frequency-domain approaches to modern time-domain control methodology that are used in the design of rotorcraft flight control laws is reviewed. Also reviewed is a study conducted to investigate the design details associated with high-gain, digital flight control systems for combat rotorcraft. Parameter identification techniques developed for rotorcraft applications are reviewed

Utilization of Differential Thrust for Lateral/Directional Stability of a Commercial Aircraft with a Damaged Vertical Stabilizer

This thesis investigates the utilization of differential thrust to help a commercial aircraft with a damaged vertical stabilizer regain its lateral/directional stability. In the event of an aircraft losing its vertical stabilizer, the consequential loss of the lateral/directional stability is likely to cause a fatal crash. In this thesis, the damaged aircraft model is constructed, and the lateral/directional dynamic stability and frequency domain analyses are conducted. The propulsion dynamics of the aircraft are modeled as a system of differential equations with engine time constant and time delay terms to study the engine response time with respect to a differential thrust input. The novel differential thrust control module is presented to map the rudder input to differential thrust input. Then, the differential thrust based control strategies such as linear quadratic regulator (LQR), model reference adaptive system (MRAS), and H∞ loop-shaping based robust control system are proposed to be utilized to help maintain stability and control of the damaged aircraft. For each type of control system design, robustness and sensitivity analysis is also conducted to test the performance of each control system in the presence of noise and uncertainty. Results demonstrate successful applications of such control methodologies as the damaged aircraft can achieve stability under feasible control efforts and without any actuator saturation. Finally, a comparison study of three control systems is conducted to investigate the merits and limits of each control system. Overall, the H∞ loop-shaping based robust control system was found to have the most remarkable results for stabilizing and saving the damaged aircraft

Coupled flight dynamics and CFD - demonstration for helicopters in shipborne environment

The development of high-performance computing and computational fluid dynamics methods have evolved to the point where it is possible to simulate complete helicopter configurations with good accuracy. Computational fluid dynamics methods have also been applied to problems such as rotor/fuselage and main/tail rotor interactions, performance studies in hover and forward flight, rotor design, and so on. The GOAHEAD project is a good example of a coordinated effort to validate computational fluid dynamics for complex helicopter configurations. Nevertheless, current efforts are limited to steady flight and focus mainly on expanding the edges of the flight envelope. The present work tackles the problem of simulating manoeuvring flight in a computational fluid dynamics environment by integrating a moving grid method and the helicopter flight mechanics solver with computational fluid dynamics. After a discussion of previous works carried out on the subject and a description of the methods used, validation of the computational fluid dynamics for ship airwake flow and rotorcraft flight at low advance ratio are presented. Finally, the results obtained for manoeuvring flight cases are presented and discussed

Optimal fault-tolerant flight control for aircraft with actuation impairments

Current trends towards greater complexity and automation are leaving modern technological systems increasingly vulnerable to faults. Without proper action, a minor error may lead to devastating consequences. In flight control, where the controllability and dynamic stability of the aircraft primarily rely on the control surfaces and engine thrust, faults in these effectors result in a higher extent of risk for these aspects. Moreover, the operation of automatic flight control would be suddenly disturbed. To address this problem, different methodologies of designing optimal flight controllers are presented in this thesis. For multiple-input multiple-output (MIMO) systems, the feedback optimal control is a prominent technique that solves a multi-objective cost function, which includes, for instance, tracking requirements and control energy minimisation. The first proposed method is based on a linear quadratic regulator (LQR) control law augmented with a fault-compensation scheme. This fault-tolerant system handles the situation in an adaptive way by solving the optimisation cost function and considering fault information, while assuming an effective fault detection system is available. The developed scheme was tested in a six-degrees-of-freedom nonlinear environment to validate the linear-based controller. Results showed that this fault tolerant control (FTC) strategy managed to handle high magnitudes of the actuator’s loss of effciency faults. Although the rise time of aircraft response became slower, overshoot and settling errors were minimised, and the stability of the aircraft was maintained. Another FTC approach has been developed utilising the features of controller robustness against the system parametric uncertainties, without the need for reconfiguration or adaptation. Two types of control laws were established under this scheme, the H∞ and µ-synthesis controllers. Both were tested in a nonlinear environment for three points in the flight envelope: ascending, cruising, and descending. The H∞ controller maintained the requirements in the intact case; while in fault, it yielded non-robust high-frequency control surface deflections. The µ-synthesis, on the other hand, managed to handle the constraints of the system and accommodate faults reaching 30% loss of effciency in actuation. The final approach is based on the control allocation technique. It considers the tracking requirements and the constraints of the actuators in the design process. To accommodate lock-in-place faults, a new control effort redistribution scheme was proposed using the fuzzy logic technique, assuming faults are provided by a fault detection system. The results of simulation testing on a Boeing 747 multi-effector model showed that the system managed to handle these faults and maintain good tracking and stability performance, with some acceptable degradation in particular fault scenarios. The limitations of the controller to handle a high degree of faults were also presented