Dissertations / Theses on the topic 'Helicopter dynamic systems'

This research work provides the development of an Adaptive Flight Control System (AFCS) for autonomous hover of a Rotary-wing Unmanned Aerial Vehicle (RUAV). Due to the complex, nonlinear and time-varying dynamics of the RUAV, indirect adaptive control using the Model Predictive Control (MPC) is utilised. The performance of the MPC mainly depends on the model of the RUAV used for predicting the future behaviour. Due to the complexities associated with the RUAV dynamics, a neural network based black box identification technique is used for modelling the behaviour of the RUAV. Auto-regressive neural network architecture is developed for offline and online modelling purposes. A hybrid modelling technique that exploits the advantages of both the offline and the online models is proposed. In the hybrid modelling technique, the predictions from the offline trained model are corrected by using the error predictions from the online model at every sample time. To reduce the computational time for training the neural networks, a principal component analysis based algorithm that reduces the dimension of the input training data is also proposed. This approach is shown to reduce the computational time significantly. These identification techniques are validated in numerical simulations before flight testing in the Eagle and RMAX helicopter platforms. Using the successfully validated models of the RUAVs, Neural Network based Model Predictive Controller (NN-MPC) is developed taking into account the non-linearity of the RUAVs and constraints into consideration. The parameters of the MPC are chosen to satisfy the performance requirements imposed on the flight controller. The optimisation problem is solved numerically using nonlinear optimisation techniques. The performance of the controller is extensively validated using numerical simulation models before flight testing. The effects of actuator and sensor delays and noises along with the wind gusts are taken into account during these numerical simulations. In addition, the robustness of the controller is validated numerically for possible parameter variations. The numerical simulation results are compared with a base-line PID controller. Finally, the NN-MPCs are flight tested for height control and autonomous hover. For these, SISO as well as multiple SISO controllers are used. The flight tests are conducted in varying weather conditions to validate the utility of the control technique. The NN-MPC in conjunction with the proposed hybrid modelling technique is shown to handle additional disturbances successfully. Extensive flight test results provide justification for the use of the NN-MPC technique as a reliable technique for control of non-linear complex dynamic systems such as RUAVs.

Este trabalho aborda o estudo de um sistema de referência inercial de posição e atitude e um sistema de controle para um helicóptero autônomo utilizando, como base para formulação e testes, o modelo linearizado da aeronave Yamaha R-MAX. Um sistema de referência inercial (INS) e um sistema de referência de atitude e orientação (AHRS) são utilizados para estimar a posição e atitude da aeronave, e estimadores robustos baseados no filtro de Kalman são empregados para minimizar os efeitos de incertezas paramétricas. É utilizada uma estratégia de controle em cascata com três malhas consistindo de uma malha interna para garantir a estabilidade do helicóptero (são utilizadas as técnicas LQR e H \'INFINITO\', separadamente), uma malha intermediária baseada em linearização por realimentação (FLC) para desacoplar os pares entrada/saída e uma malha externa baseada em um controlador proporcional-derivativo (PD) para permitir o rastreamento da trajetória. Resultados de simulação são apresentados para avaliar o desempenho de cada abordagem.
This work concerns the study of an inertial reference system and a control system for an autonomous helicopter using, as basis for the formulation and testing, the linearized mo del of the aircraft Yamaha R-MAX. An inertial navigation system (INS) and an attitude and orientation reference system (AHRS) are used to estimate the position and attitude of the aircraft and robust estimators based on Kalman filter are employed to minimize the effects of parametric uncertainties. A cascaded control architecture with three control methodologies is used, consisting of an inner-loop to ensure stability of the helicopter (the LQR and H \'INFINITE\' techniques are used, separately), a mid-loop based on linearization feedback (FLC) to decouple the dynamics ofthe lateral, longitudinal, vertical and heading axes and an outer-loop based on a proportional-derivative (PD) controller to enable trajectory tracking. Simulation results are presented to evaluate the performance of each approach.

Les nouveaux outils logiciels permettent de simuler le comportement dynamique transitoire des systèmes mécaniques multi-corps. Pourraient-ils produire une bonne prévision des efforts qui prennent place dans les systèmes de voilures tournantes des hélicoptères? Les travaux développés dans cette thèse apportent des éléments de réponse à cette question, sachant que le contexte considéré est porteur de nombreuses difficultés : flexibilités des corps, liaisons mécaniques complexes, hyperstatismes…L’enjeu est important car le cycle de développement de ces ensembles mécaniques serait fortement réduit par une bonne estimation d’efforts dès les premières étapes de la conception d’une nouvelle architecture. Dans ce cadre, un nouveau niveau de modélisation est introduit, centré sur les systèmes dynamiques à l’étude. La présentation de l’approche multi-échelle proposée implique le recours à des modèles locaux simplifiés. L’étude s’appuie sur le système de voilure tournante de l’hélicoptère H160 actuellement en phase d’industrialisation chez Airbus Helicopters et qui comprend les pales, le rotor principal et les actionneurs de commande hydrauliques. Trois développements majeurs sont détaillés : modélisations d’un sous-ensemble plateaux cycliques, des pales et du système de voilure tournante. Le nouveau cadre de modélisation ainsi constitué permet d’estimer les interefforts au niveau des liaisons mécaniques du système, et de suivre leurs évolutions au cours du temps
New software tools are used to simulate the transient dynamic behaviour of multi-body mechanical systems. Could they provide a good forecast of loads applied on helicopter rotary wing system? The work developed in this thesis provides some answers to this question, knowing that the context considered presents many difficulties: flexible bodies, complex mechanical links, hyperstatisms… The stakes are high because the development cycle of these mechanical assemblies would be greatly reduced by a correct estimation of loads during the first steps of the design of a new architecture. In this context, a new level of modeling is introduced, focusing on the dynamic systems studied. Since this level is part of a multi-scale approach, it is necessary to feed it with simplified models of the subassemblies that form the studied system, and this leads to so-called local studies. The presentation of this work is based on the rotary wing system of H160 helicopter, currently in industrialization phase at Airbus Helicopters, which includes the blades, the main rotor and the hydraulic control actuators. Three major developments are detailed: modelling of a swashplates sub-assembly, blades and the rotary wing system.The new modelling framework thus created allows the loads estimation at the level of the mechanical links of these systems and the monitoring of their evolution over time

A helicopter can be used to transport a load hanging from a suspension cable. This technique is frequently used in construction, firefighting, and disaster relief operations, among other applications. Unfortunately, the suspended load swings, which makes load positioning difficult and can degrade control of the helicopter. This dissertation investigates the use of input shaping (a command-filtering technique for reducing vibration) to mitigate the load swing problem. The investigation is conducted using two different, but complementary, approaches. One approach studies manual tracking tasks, where a human attempts to make a cursor follow an unpredictably moving target. The second approach studies horizontal repositioning maneuvers on small-scale helicopter systems, including a novel testbed that limits the helicopter and suspended load to move in a vertical plane. Both approaches are used to study how input shaping affects control of a flexible element (the suspended load) and a driven base (the helicopter). In manual tracking experiments, conventional input shapers somewhat degraded control of the driven base but greatly improved control of the flexible element. New input shapers were designed to improve load control without negatively affecting base control. A method for adjusting the vibration-limiting aggressiveness of any input shaper between unshaped and fully shaped was also developed. Next, horizontal repositioning maneuvers were performed on the helicopter testbed using a human-pilot-like feedback controller from the literature, with parameter values scaled to match the fast dynamics of the model helicopter. It was found that some input shapers reduced settling time and peak load swing when applied to Attitude Command or Translational Rate Command response types. When the load was used as a position reference instead of the helicopter, the system was unstable without input shaping, and adding input shaping to a Translational Rate Command was able to stabilize the load-positioning system. These results show the potential to improve the safety and efficiency of helicopter suspended load operations.

Multi-Body System Simulation combined with System Identification was developed as a method for determining the stability characteristics of a human powered helicopter(HPH) configurations. HPH stability remains a key component for meeting competition requirements, but has not been properly treated. Traditional helicopter dynamic analysis is not suited to the HPH due to its low rotation speeds and light weight. Multi-Body System Simulation is able to generate dynamic response data for any HPH configuration. System identification and linear stability theory are used to determine the stability characteristics from the dynamic response. This thesis focuses on the method development and doesn't present any HPH analysis results.

The thesis presents a framework for integrating models, simulation, and experimental data to diagnose incipient failure modes and prognosticate the remaining useful life of critical components, with an application to the main transmission of a helicopter. Although the helicopter example is used to illustrate the methodology presented, by appropriately adapting modules, the architecture can be applied to a variety of similar engineering systems. Models of the kind referenced are commonly referred to in the literature as physical or physics-based models. Such models utilize a mathematical description of some of the natural laws that govern system behaviors. The methodology presented considers separately the aspects of diagnosis and prognosis of engineering systems, but a similar generic framework is proposed for both. The methodology is tested and validated through comparison of results to data from experiments carried out on helicopters in operation and a test cell employing a prototypical helicopter gearbox. Two kinds of experiments have been used. The first one retrieved vibration data from several healthy and faulted aircraft transmissions in operation. The second is a seeded-fault damage-progression test providing gearbox vibration data and ground truth data of increasing crack lengths. For both kinds of experiments, vibration data were collected through a number of accelerometers mounted on the frame of the transmission gearbox. The applied architecture consists of modules with such key elements as the modeling of vibration signatures, extraction of descriptive vibratory features, finite element analysis of a gearbox component, and characterization of fracture progression. Contributions of the thesis include: (1) generic model-based fault diagnosis and failure prognosis methodologies, readily applicable to a dynamic large-scale mechanical system; (2) the characterization of the vibration signals of a class of complex rotary systems through model-based techniques; (3) a reverse engineering approach for fault identification using simulated vibration data; (4) the utilization of models of a faulted planetary gear transmission to classify descriptive system parameters either as fault-sensitive or fault-insensitive; and (5) guidelines for the integration of the model-based diagnosis and prognosis architectures into prognostic algorithms aimed at determining the remaining useful life of failing components.

On-board gimbal systems for camera stabilization in helicopters are typically based on linear models. Such models, however, are inaccurate due to system nonlinearities and complexities. As an alternative approach, artificial neural networks can provide a more accurate model of the gimbal system based on their non-linear mapping and generalization capabilities. This thesis investigates the applications of artificial neural networks to model the inertial characteristics (on the azimuth axis) of the inner gimbal in a gyro-stabilized multi-gimbal system. The neural network is trained with time-domain data obtained from gyro rate sensors of an actual camera system. The network performance is evaluated and compared with measured data and a traditional linear model. Computer simulation results show the neural network model fits well with the measured data and significantly outperforms a traditional model.

Simulation of maneuvers with multibody models of rotorcraft vehicles is an important research area due to its complexity. During the maneuvering flight, some important design limitations are encountered such as maximum loads and maximum turning rates near the proximity of the flight envelope. This increases the demand on high fidelity models in order to define appropriate controls to steer the model close to the desired trajectory while staying inside the boundaries. A framework based on the hierarchical decomposition of the problem is used for this study. The system should be capable of generating the track by itself based on the given criteria and also capable of piloting the model of the vehicle along this track. The generated track must be compatible with the dynamic characteristics of the vehicle. Defining the constraints for the maneuver is of crucial importance when the vehicle is operating close to its performance boundaries. In order to make the problem computationally feasible, two models of the same vehicle are used where the reduced model captures the coarse level flight dynamics, while the fine scale comprehensive model represents the plant. The problem is defined by introducing planning layer and control layer strategies. The planning layer stands for solving the optimal control problem for a specific maneuver of a reduced vehicle model. The control layer takes the resulting optimal trajectory as an optimal reference path, then tracks it by using a non-linear model predictive formulation and accordingly steers the multibody model. Reduced models for the planning and tracking layers are adapted by using neural network approach online to optimize the predictive capabilities of planner and tracker. Optimal neural network architecture is obtained to augment the reduced model in the best way. The methodology of adaptive learning rate is experimented with different strategies. Some useful training modes and algorithms are proposed for these type of applications. It is observed that the neural network increased the predictive capabilities of the reduced model in a robust way. The proposed framework is demonstrated on a maneuvering problem by studying an obstacle avoidance example with violent pull-up and pull-down.

System dynamics modelling has been used for around 40 years to address complex, systemic, dynamic problems, those often described as wicked. But, system dynamics modelling is not an exact science and arguments about the most suitable techniques to use in which circumstances, continues. The nature of these wicked problems is investigated through a series of case studies where poor situational awareness among stakeholders was identified. This was found to be an underlying cause for management failure, suggesting need for better ways of recognising and managing wicked problem situations. Human cognition is considered both as a limitation and enabler to decision-making in wicked problem environments. Naturalistic and deliberate decision-making are reviewed. The thesis identifies the need for integration of qualitative and quantitative techniques. Case study results and a review of the literature led to identification of a set of principles of method to be applied in an integrated framework, the aim being to develop an improved way of addressing wicked problems. These principles were applied to a series of cases in an action research setting. However, organisational and political barriers were encountered. This limited the exploitation and investigation of cases to varying degrees. In response to a need identified in the literature review and the case studies, a tool is designed to facilitate analysis of multi-factorial, non-linear causality. This unique tool and its use to assist in problem conceptualisation, and as an aid to testing alternate strategies, are demonstrated. Further investigation is needed in relation to the veracity of combining causal influences using this tool and system dynamics, broadly. System dynamics modelling was found to have utility needed to support analysis of wicked problems. However, failure in a particular modelling project occurred when it was found necessary to rely on human judgement in estimating values to be input into the models. This was found to be problematic and unacceptably risky for sponsors of the modelling effort. Finally, this work has also identified that further study is required into: the use of human judgement in decision-making and the validity of system dynamics models that rely on the quantification of human judgement.

L’une des principales sources d’inconfort dans un hélicoptère sont les vibrations transmises par le rotor à la structure de l’appareil. En vol d’avancement, des efforts aérodynamiques cycliques sont subis par l’ensemble des pales en tête rotor et génèrent de très fortes vibrations basse fréquence (aux alentours des 17Hz) transmises aux passagers via la boîte de transmission principale puis le fuselage lui-même. Afin de garantir le confort des membres d’équipage et des passagers, de nombreux systèmes antivibratoires ont été conçus. Ces systèmes sont généralement passifs car la majorité de l’énergie vibratoire transmise à la structure se situe à une fréquence unique ωc correspondant à bΩ avec b le nombre de pales et Ω la fréquence de rotation du rotor. Cependant, les appareils modernes évoluent et le régime rotor jusqu’alors fixe durant toutes les phases de vol varie à présent pour des préoccupations de performances et de consommation (variation de l’ordre de +/-10% autour de bΩ). Cette nouvelle contrainte dans la conception des hélicoptères rend pertinente la technologie des systèmes antivibratoires actifs, pouvant s’adapter à la sollicitation en termes d’amplitude et fréquence. Lors de ces travaux de thèse, la suspension passive SARIB de Airbus Helicopters basée sur le principe du DAVI (Dynamic Antiresonant Vibration Isolator) est modifiée afin d’être rendue active par ajout d’une partie actuation/commande. La théorie des lois et algorithmes de contrôle utilisés dans ces travaux, est présentée en détail afin de poser solidement les bases du contrôle actif du prototype de suspension conceptualisé ici à savoir le contrôle FXLMS (adaptatif) et le contrôle optimal LQG. Afin de simuler le fonctionnement du système, un modèle tridimensionnel de la suspension active est construit, couplé à la structure souple de l’hélicoptère (NH90). Sur ce modèle sont alors appliquées les différentes lois de commande introduites auparavant et leurs performances comparées dans différents cas de chargement en tête rotor et surtout pour différentes fréquences de sollicitation. De même, pour chaque algorithme, différentes localisations des capteurs d’erreur sont étudiées afin de converger vers une configuration optimale. Les simulations démontrent que l’algorithme FXLMS feedforward est très bien adapté au contrôle des perturbations harmoniques et permet de réduire très significativement le niveau vibratoire du plancher cabine, sans réinjection parasite dans le reste de la structure. Une comparaison de l’efficacité du SARIB actif avec les systèmes d’absorbeurs en cabine est ensuite effectuée pour démontrer la pertinence d’utiliser le principe du DAVI comme base d’un système actif. Les travaux de cette thèse traitent également des essais réalisés en laboratoire sur le prototype échelle 1 de la suspension SARIB active avec contrôle FXLMS
One of the main causes of discomfort in helicopters are the vibrations transmitted from the rotor to the structure. In forward flight, the blades are submitted to cyclic aerodynamic loads which generate low frequency (around 17Hz) but high energy mechanical vibrations. These vibrations are transmitted from the rotor to the main gearbox, then to the structure and finally to the crew and passengers. In order to maintain acceptable comfort for crew members and passengers, a lot of antivibration devices have been developed since the last 30 years. These systems are generally passive because most of the mechanical energy transmitted to the structure is at only one frequency ωc which is equal to the product bΩ with b the number of blades and Ω the rotor rotational speed. However, modern helicopters evolve and the rotor rpm, which has always been considered as fixed during flight is now a function of time, depending on the flight phases in order to increase performances and reduce energy consumption (variation bandwidth of Ω +/- 10%). This new constraint on the design of helicopters makes the active antivibration technology completely relevant with its capacity to adapt in terms of amplitude and frequency to the perturbation. During this thesis, the passive suspension called SARIB from Airbus Helicopters, based on the DAVI principle (Dynamic Antiresonant Vibration Isolator) is modified in order to implement active components and command (actuation). The theory of the control algorithms used in this thesis is presented in detail in order to define the theoretical tools of the active DAVI control which are : FXLMS control (adaptive control) and LQG (optimal control). To simulate the complete system, a 3D multibody model of the active suspension has been set up, coupled to a the flexible structure of a NH90 (Airbus Helicopters). On this model are applied the different control algorithms presented before and their performances are compared for different loads with variable frequency on the rotor hub. In the same way, different locations for the error sensors in the structure are studied to find the optimal control configuration. The simulations show that the FXLMS algorithm is well suited for the control of harmonic perturbations and reduce significantly the dynamic acceleration level on the cabin floor, without parasite reinjection on other parts of the structure. A comparison of the active SARIB with classical cabin vibration absorbers is also made in terms of efficiency in order to show the advantages of using the DAVI system as a base for an active antivibration device. Finally, this thesis also presents the experiments realized in the dynamics laboratory of Airbus Helicopters on a 1:1 scale prototype of the active SARIB suspension with FXLMS control. The results demonstrate the efficiency of the active suspension architecture and control algorithms

Military helicopter operations require precise maneuverability characteristics for performance to be determined for the entire helicopter flight envelope. Historically, these maneuverability analyses are combinatorial in nature and involve human-interaction, which hinders their integration into conceptual design. A model formulation that includes the necessary quantitative measures and captures the impact of changing requirements real-time is presented. The formulation is shown to offer a more conservative estimate of maneuverability than traditional energy-based formulations through quantitative analysis of a typical pop-up maneuver. Although the control system design is not directly integrated, two control constraint measures are deemed essential in this work: control deflection rate and trajectory divergence rate. Both of these measures are general enough to be applied to any control architecture, while at the same time enable quantitative trades that relate overall vehicle maneuverability to control system requirements. The dimensionality issues stemming from the immense maneuver space are mitigated through systematic development of a maneuver taxonomy that enables the operational envelope to be decomposed into a minimal set of fundamental maneuvers. The taxonomy approach is applied to a helicopter canonical example that requires maneuverability and design to be assessed simultaneously. The end result is a methodology that enables the impact of design choices on maneuverability to be assessed for the entire helicopter operational envelope, while enabling constraints from control system design to be assessed real-time.

Thesis (M. S.)--Aerospace Engineering, Georgia Institute of Technology, 2005.
Bottasso, Carlo, Committee Chair ; Hodges, Dewey, Committee Member ; Bauchau, Olivier, Committee Member. Includes bibliographical references.

Lau, Tak Kit.
Thesis (M.Phil.)--Chinese University of Hong Kong, 2009.
Includes bibliographical references (leaves 168-176).
Abstract also in Chinese.
Chapter 1 --- Introduction --- p.18
Chapter 1.1 --- Motivation and Problem Statement --- p.18
Chapter 1.2 --- Literature Review --- p.19
Chapter 1.2.1 --- Avionics Design --- p.19
Chapter 1.2.2 --- Controller Design --- p.21
Chapter 1.2.3 --- Dynamics Analysis --- p.23
Chapter 1.2.4 --- State Estimation for GNSS Outage --- p.24
Chapter 1.3 --- Outline --- p.25
Chapter 2 --- Actuation Dynamics --- p.26
Chapter 2.1 --- Mcchanism Of The Rotor --- p.27
Chapter 2.2 --- Mcchanism Of Swashplate And Rotor --- p.28
Chapter 2.3 --- Numerical Analysis Of Cyclic Pitch Angle --- p.31
Chapter 2.4 --- Helicopter Dynamics --- p.33
Chapter 2.4.1 --- Aerodynamic Forces And Moments --- p.35
Chapter 2.4.2 --- Aerodynamic Drag --- p.36
Chapter 2.4.3 --- Incremental Lift --- p.36
Chapter 2.4.4 --- Tail Rotor Thrust And Moment --- p.36
Chapter 2.4.5 --- Deadweight And Moment --- p.37
Chapter 2.5 --- The Conventional Inadequacy Of Adding A 90° Phase-Lag --- p.38
Chapter 2.6 --- The Gyroscopic Effect In Helicopter Dynamics --- p.39
Chapter 2.6.1 --- How Precession Works --- p.43
Chapter 2.6.2 --- The Analytical Form --- p.45
Chapter 2.6.3 --- Numerical Analysis Of The Gyroscopic Effect --- p.48
Chapter 3 --- State Estimation For GNSS Outage --- p.52
Chapter 3.1 --- GNSS Error And UAV Failure --- p.52
Chapter 3.2 --- "Kalman Filter, And The Extended Kalman Filter" --- p.53
Chapter 3.3 --- Unscented Kalman Filter --- p.54
Chapter 3.4 --- Process And Measurement Model --- p.55
Chapter 3.4.1 --- The IMU Driven Model And Sensor Error --- p.57
Chapter 3.5 --- Modifications To The Model And UKF Algorithm --- p.62
Chapter 3.5.1 --- Acceleration White Noise Bias (AWNB) --- p.62
Chapter 3.5.2 --- Acceleration Scale (AS) --- p.64
Chapter 3.5.3 --- Prioritized Propagation Of States (PPS) --- p.64
Chapter 3.5.4 --- Performance Of The Proposed Enhancements --- p.66
Chapter 3.5.5 --- Tripled Percentage Reduction Of Position RMSE When Using PPS With AWNB --- p.73
Chapter 4 --- Autopilot For Attitude Stabilization --- p.84
Chapter 4.1 --- Oil Test Bondi --- p.85
Chapter 4.2 --- On Unconstrained Flight --- p.88
Chapter 4.2.1 --- Tracking Reference Problem --- p.88
Chapter 4.2.2 --- An Alternative To PID Attitude Control --- p.91
Chapter 4.2.3 --- The Proposed Hierarchical PD Controller --- p.91
Chapter 4.2.4 --- Stability Analysis --- p.92
Chapter 4.2.5 --- Hierarchy Of The Varying Tracking Reference --- p.95
Chapter 4.2.6 --- Asymptotical Stability And Robustness --- p.99
Chapter 4.2.7 --- Experiment and Performance Of The Proposed Controller --- p.101
Chapter 5 --- Avionics And Test Bench Design --- p.105
Chapter 5.1 --- Avionics Design --- p.105
Chapter 5.1.1 --- Design Essentials --- p.107
Chapter 5.1.2 --- Synchronization Of Commands --- p.107
Chapter 5.1.3 --- Normalization Of Servomechanism Commands --- p.110
Chapter 5.2 --- Test Bench Design --- p.110
Chapter 5.2.1 --- The Idea --- p.111
Chapter 5.2.2 --- Concern --- p.111
Chapter 5.2.3 --- Test Bench Design Options --- p.112
Chapter 5.2.4 --- Building The Test Bench --- p.113
Chapter 5.2.5 --- Disturbance In IMU Data --- p.113
Chapter 5.2.6 --- The Solution To IMU Saturation --- p.115
Chapter 6 --- Conclusion --- p.118
Chapter 6.1 --- Actuation Dynamics --- p.118
Chapter 6.2 --- State Estimation for GNSS Outage --- p.119
Chapter 6.3 --- Hierarchical PD Controller --- p.121
Chapter A --- Appendix - Derivation From Recursive Least Square Estimation To Kalman Filter --- p.122
Chapter A.1 --- Recursive Least Square --- p.122
Chapter A.1.1 --- Alternate Estimator form for RLS --- p.134
Chapter A.1.2 --- Propagation of States and Covariance --- p.137
Chapter A.1.3 --- Kalman Filter --- p.139
Chapter B --- Appendix - Actuation by Gyroscopic Effect --- p.144
Chapter B.1 --- Expression of The Induced Moment Due to Gyroscopic Effect In The Total External Moment --- p.150
Chapter B.2 --- An Illustrated Example --- p.153
Chapter B.3 --- Another Derivation By Using A Different Orientation Definition --- p.156
Chapter B.4 --- Dimensions of the helicopter for experiments --- p.166
Bibliography --- p.167

The present work focuses on the two areas of investigation: system identification of helicopter and design of controller for the helicopter. Helicopter system identification, the first subject of investigation in this thesis, can be described as the extraction of system characteristics/dynamics from measured flight test data. Wind tunnel experimental data suffers from scale effects and model deficiencies. The increasing need for accurate models for the design of high bandwidth control system for helicopters has initiated a renewed interest in and a more active use of system identification. Besides, system identification is likely to become mandatory in the future for model validation of ground based helicopter simulators. Such simulators require accurate models in order to be accepted by pilots and regulatory authorities like Federal Aviation Regulation for realistic complementary helicopter mission training. Two approaches are widely used for system identification, namely, black box and gray box approach. In the black-box approach, the relationship between input-output data is approximated using nonparametric methods such as neural networks and in such a case, internal details of the system and model structure may not be known. In the gray box approach, parameters are estimated after defining the model structure. In this thesis, both black box and gray box approaches are investigated. In the black box approach, in this thesis, a comparative study and analysis of different Recurrent Neural Networks(RNN) for the identification of helicopter dynamics using flight data is investigated. Three different RNN architectures namely, Nonlinear Auto Regressive eXogenous input(NARX) model, neural network with internal memory known as Memory Neuron Networks(MNN)and Recurrent MultiLayer perceptron (RMLP) networks are used to identify dynamics of the helicopter at various flight conditions. Based on the results, the practical utility, advantages and limitations of the three models are critically appraised and it is found that the NARX model is most suitable for the identification of helicopter dynamics. In the gray box approach, helicopter model parameters are estimated after defining the model structure. The identification process becomes more difficult as the number of degrees-of-freedom and model parameters increase. To avoid the drawbacks of conventional methods, neural network based techniques, called the delta method is investigated in this thesis. This method does not require initial estimates of the parameters and the parameters can be directly extracted from the flight data. The Radial Basis Function Network(RBFN)is used for the purpose of estimation of parameters. It is shown that RBFN is able to satisfactorily estimate stability and control derivatives using the delta method. The second area of investigation addressed in this thesis is the control of helicopter in flight. Helicopter requires use of a control system to achieve satisfactory flight. Designing a classical controller involves developing a nonlinear model of the helicopter and extracting linearized state space matrices from the nonlinear model at various flight conditions. After examining the stability characteristics of the helicopter, the desired response is obtained using a feedback control system. The scheduling of controller gains over the entire envelope is used to obtain the desired response. In the present work, a helicopter having a soft inplane four bladed hingeless main rotor and a four-bladed tail rotor with conventional mechanical controls is considered. For this helicopter, a mathematical model and also a model based on neural network (using flight data) has been developed. As a precursor, a feed back controller, the Stability Augmentation System(SAS), is designed using linear quadratic regulator control with full state feedback and LQR with out put feedback approaches. SAS is designed to meet the handling qualities specification known as Aeronautical Design Standard ADS-33E-PRF. The control gains have been tuned with respect to forward speed and gain scheduling has been arrived at. The SAS in the longitudinal axis meets the requirement of the Level1 handling quality specifications in hover and low speed as well as for forward speed flight conditions. The SAS in the lateral axis meets the requirement of the Level2 handling quality specifications in both hover and low speed as well as for forward speed flight conditions. Such conventional design of control has served useful purposes, however, it requires considerable flight testing which is time consuming, to demonstrate and tune these control law gains. In modern helicopters, the stringent requirements and non-linear maneuvers make the controller design further complicated. Hence, new design tools have to be explored to control such helicopters. Among the many approaches in adaptive control, neural networks present a potential alternative for modeling and control of nonlinear dynamical systems due to their approximating capabilities and inherent adaptive features. Furthermore, from a practical perspective, the massive parallelism and fast adaptability of neural network implementations provide more incentive for further investigation in problems involving dynamical systems with unknown non-linearity. Therefore, adaptive control approach based on neural networks is proposed in this thesis. A neural network based Feedback Error Neural adaptive Controller(FENC) is designed for a helicopter. The proposed controller scheme is based on feedback error learning strategy in which the outer loop neural controller enhances the inner loop conventional controller by compensating for unknown non-linearity and parameter un-certainties. Nonlinear Auto Regressive eXogenous input(NARX)neural network architecture is used to approximate the control law and the controller network parameters are adapted using updated rules Lyapunov synthesis. An offline (finite time interval)and on-line adaptation strategy is used to approximate system uncertainties. The results are validated using simulation studies on helicopter undergoing an agile maneuver. The study shows that the neuro-controller meets the requirements of ADS-33 handling quality specifications. Even though the tracking error is less in FENC scheme, the control effort required to follow the command is very high. To overcome these problems, a Direct Adaptive Neural Control(DANC)scheme to track the rate command signal is presented. The neural controller is designed to track rate command signal generated using the reference model. For the simulation study, a linearized helicopter model at different straight and level flight conditions is considered. A neural network with a linear filter architecture trained using back propagation through time is used to approximate the control law. The controller network parameters are adapted using updated rules Lyapunov synthesis. The off-line trained (for finite time interval)network provides the necessary stability and tracking performance. The on-line learning is used to adapt the network under varying flight conditions. The on-line learning ability is demonstrated through parameter uncertainties. The performance of the proposed direct adaptive neural controller is compared with feedback error learning neural controller. The performance of the controller has been validated at various flight conditions. The theoretical results are validated using simulation studies based on a nonlinear six degree-of-freedom helicopter undergoing an agile maneuver. Realistic gust and sensor noise are added to the system to study the disturbance rejection properties of the neural controllers. To investigate the on-line learning ability of the proposed neural controller, different fault scenarios representing large model error and control surface loss are considered. The performances of the proposed DANC scheme is compared with the FENC scheme. The study shows that the neuro-controller meets the requirements of ADS-33 handling quality specifications.

碩士
國立成功大學
航空太空工程學系碩博士班
93
From 1970’s, the development of Integrated Circuit (IC) has grown into more important in hardware circuit, because of its stable capability. Without a doubt, the reasons are its smallness, preciseness, high speed, and cheapness. Due to these advantages, we use DSP chip which is manufactured by Texas Instrument to implement this dynamic simulation system in this thesis. The complicated helicopter control loops are simulated by MATLAB & Simulink, translated to C code by real-time workshop (RTW), and finally downloaded to DSP .What we have expect is making the helicopter auto-pilot system lighter and smaller through this research. 　The contribution in this thesis is combination of system design on a loop : helicopter nonlinear model, LMI flight controller, ship motion, and helicopter shipboard operational environment. We utilize DSP chip for controller of this system to implement the designing of hardware-in-the-loop.

The helicopter is a Multiple-Input Multiple-Output (MIMO) system with highly coupled characteristics, which increases the complexity of the system dynamics. In addition, the system dynamics of the helicopter are unstable, referring to its tendency to deviate from an equilibrium when disturbed. Despite the complexity in its modelling and control, the benefit of using a helicopter for unmanned, autonomous applications can be tremendous. One particular application that motivates this research is the use of an unmanned small-scale helicopter in an autonomous survey mission over an area struck by disaster, such as an earthquake. The work presented in this thesis provides a framework for utilizing a platform system for research and development of small-scale helicopter systems. A platform system enables testing and analysis to be performed indoor in a controlled environment. This can provide a more convenient mean for helicopter research since the system is not affected by environmental elements, such as wind, rain or snow condition. However, the presence of the platform linkages poses challenges for analysis and controller design as it alters the helicopter system flight dynamics. Through a six degree-of-freedom (6 DOF) platform model derived in this research, the criteria for matching the trim conditions between the platform system and a stand alone helicopter have been identified. With the matched trim conditions, linearization is applied to perform analysis on the effects that the platform has on the system dynamics. The results of the analysis provide insights into both the limitations and benefits of utilizing the platform system for helicopter research. Finally, a Virtual Joint Control scheme is proposed as an unified control strategy for both the platform and the stand alone helicopter systems. Having a consistent control scheme between the two systems allows for comparisons between simulation and experimental results for the two systems to be made more readily. Furthermore, the Virtual Joint Control scheme represents a novel flight control strategy for stand alone helicopter systems.

碩士
國防大學
資源管理及決策研究所
105
The principle of our military strategy follows the concept of “Strong Defense Posture, Effective Deterrence” and strengthens the capability of air-ground operations, Special Forces, logistic support, etc. Since the Army Aviation and Special Forces Command’s helicopters have high mobility and can work with the air-ground operations; therefore, it could implement a variety of tasks. As a result, the Ministry of National Defense uses the helicopter troops as a part of restructuring of the armed forces. This research focus on the type of Army Aviation helicopters, as this type of helicopters has been serving for more than twenty years accumulating operation hours. The Army Aviation units are facing the problems of spare parts shortage and delivery delay. Consequently, the spare parts inventory is hard to meet the maintenance requirement, which let the aircraft be grounded or even be put into storage. We apply system dynamics to find out the cause-effect relation of critical variables and build a system dynamic analysis model. As a result, this model can be used to simulate different policies, and see the differences of the availability, the number of storage, the spare parts demand and supply, the man-hour demand and the delay time of aircraft removal from storage. The policies, include budget satisfaction rate, safety stock, one for one exchange rate, domestic repair ratio, and storage period. Relevant policy recommendations will also be discussed in this paper.

A dissertation submitted to the Faculty of Engineering and the Built Environment, University of the Witwatersrand, Johannesburg, in partial fulfilment of the requirements for the degree of Master of Science in Engineering. Johannesburg, April 2016
A Rotary Wing Unmanned Aerial Vehicle (RW UAV) as a platform and its payload consisting of sophisticated sensors would be costly items. Hence, a RW UAV in the 500 kg class designed to fulfil a number of missions would represent a considerable capital outlay for any customer. Therefore, in the event of an engine failure, a means should be provided to get the craft safely back on the ground without incurring damage or causing danger to the surrounding area. The aim of the study was to design a controller for autorotative landing of a RW UAV in the event of engine failure. In order to design a controller for autorotative landing, an acceleration model was used obtained from a study by Stanford University. FLTSIM helicopter flight simulation package yielded necessary RW UAV response data for the autorotation regimes. The response data was utilized in identifying the unknown parameters in the acceleration model. A Differential Dynamic Programming (DDP) control algorithm was designed to compute the main and tail rotor collective pitch and the longitudinal and lateral cyclic pitch control inputs to safely land the craft. The results obtained were compared to the FLTSIM flight simulation response data. It was noted that the mathematical model could not accurately model the pitch dynamics. The main rotor dynamics were modelled satisfactorily and which are important in autorotation because without power from the engine, the energy in main rotor is critical in a successful execution of an autorotative landing. Stanford University designed a controller for RC helicopter, XCell Tempest, which was deemed successful. However, the DDP controller was designed for autonomous autorotative landing of RW UAV weighing 560 kg, following engine failure. The DDP controller has the ability to control the RW UAV in an autorotation landing but the study should be taken further to improve certain aspects such as the pitch dynamics and which can possibly be achieved through online parameter estimation.
MT 2017

碩士
淡江大學
機械工程研究所
83
The induced wake has a great effect on the stability of a helicopter rotor blades in hover and forward hlight. The research work in this paper is to study the interactions among induced wake and rotor blades. 　　This paper contains three parts:(1) the stability of induced wake and its mode shapes analysis. (2) the coupled system analysis of induced wake and rigid blades. (3) the coupled system analysis of induced wake and elastic blades. The Peters' Generalized Dynamic Inflow Theory with its accuracy and excellent characteristic in coupling is chosen as induced wakde theory. The merit of aeroelastic system coupled with rigid blades and elastic blades lies in it's a three-dimensional model and can be ulitized to predict the stability of a helicopter in different foward flight conditions. This aeroelastic system is etablished in a non-rotating system with which it containts periodic coefficients; thus, Floquet Theory is chosen to perform an eigen-analysis. 　　The results of this paper provide us an eigen-analysis among induced wake coupled with rigid blades and with elastic blades in aeroelastic system. In accordance with the results of this eigen-analysis, we realize that induced wake on rigid and elastic blades flapping damper somehow has an effect in hover and forward flight; especially, the effect in forward flight is much stronger. Above all, the research work in this paper can help to realize a helicopter flight conditions and also to motivate its futher research and design.

碩士
國防大學
資源管理及決策研究所
106
Along with the advances in technology and changing in the form of war, the reformation of the R.O.C. Armed Forces should focus on keeping cultivating talents, establishing asymmetric warfare capabilities, and innovating military theories. For the perspective of national security and military defense, safeguarding the security of sea areas is an important issue for Taiwan, a country which is surrounded by sea. The anti-submarine helicopters of our Naval Anti-Submarine Aviation Wing have become an important force in naval anti-submarine operations due to their high mobility and strong monitoring capability. Maintaining the equipment to make helicopters successfully carry out anti-submarine missions has also become an issue that is worthy of attention and discussion. In this study, we use naval anti-submarine helicopters as research object, apply system dynamics to find out the key variables that affect the helicopter maintenance mode and build a system dynamics model. We use this model to analyze and simulate in different policies, and explore the impact of logistic maintenance policy adjustments on the helicopter availability and maintenance costs. Relevant policy recommendations will also be discussed in this paper.

碩士
國防大學
資源管理及決策研究所
107
This study takes a certain type of attack helicopter in Army Aviation as the research object and uses system dynamics to explore the key factors affecting the helicopter's availability of maintenance and supply policy. This study builds a dynamic analysis model and divides the model into two parts: one is a maintenance and supply policy analysis model for peacetime; the other is a combat simulation model assuming regional conflict in 2035 and taking the available attack helicopters from peacetime model as input. There are several factors taken into consideration, such as annual task hours, spare parts inventory with safety level, the budget satisfaction rate of military procurement and pre-operation time. All these factors have effects on helicopter availability and cost-effectiveness in peacetime. The wartime simulation model takes the helicopter availability from peacetime model as input, and analyze the combat effectiveness (attrition and killed rate) for both sides with Lanchester’s Law in anti-landing operation. This study builds a system dynamics model combining peacetime and wartime considerations, which could be used as a reference for making the regulation of maintenance and supply policy.

碩士
淡江大學
航空太空工程學系
87
Optimum design of two different design variables, chord length and twist angle, through an unsteady aerodynamic system will be considered in this study. Besides, this paper also discuss the coupling effect between chord length and twist angle, and apply wake dynamics, areodynamics and optimality criterion theory to obtain the optimum configuration of rotor blades. The purpose of this study is to obtain a helicopter blades' chord length and twist angle which to minimize the power output and also maintain the lift force in a misson. Because design variable doubles and couples so that the problem becomes complicated, an improve move limit with Bezier curve technique will be implemented in optimal program to overcome these effect. The BELL UH-1H helicopter rotor blades will be redesign by optimum design program exactly and steadily. The result of new design rotor blade will compare with the original rectangular rotor blade in numerical examples.