Helicopter Model Research Articles

A New Robust Controller Formulation for the Full–State Feedback Position Tracking of a Small–Scaled Unmanned Model Helicopter

This work focuses on the robust attitude tracking control problem for a small– scaled unmanned helicopter where the actual system inputs, namely the elevator servo input, the aileron servo input and the rudder servo input, are used in the controller formulation. The design process is divided into two parts. Initially the problem is transformed into a second order system with an uncertain non–symmetric input gain matrix by utilizing some reasonable simplifications for the rotor model under the hovering flight conditions. Then a novel robust control methodology is utilized via a matrix decomposition method. The stability of the overall system is ensured by Lyapunov type analysis where asymptotic position tracking is ensured. Numerical simulation results are presented to demonstrate the efficiency of the proposed method.

Hybrid high-order sliding mode-based control for multivariable cross-coupling systems: Scale-laboratory helicopter system application

ABSTRACTThis paper presents a high-order sliding mode approach to design variable structure controllers for nonlinear multivariable cross-coupling systems whereby a change in either control input affects the control loops of all the subsystems dynamics. A hybrid-sliding mode control (hybrid-SMC) scheme is constructed combining a new equivalent control algorithm with a high-order discontinuous control algorithm to benefit from the advantages of both control strategies. The hybrid-SMC scheme uses weighting coefficients to weight and combine controllers. The equivalent control algorithm uses a relative degree concept through dynamic constraints imposed on the sliding variables to overcome the limitations of the conventional approaches and to provide an optimum tracking performance. A scale-laboratory helicopter model is used to sum up the main features and demonstrate the effectiveness of the developed control scheme. The proposed hybrid-SMC strategy is compared to existing sliding mode-based control approaches in terms of tracking performance, stability and control efforts. The obtained results demonstrate the validity and efficiency of the proposed hybrid-SMC scheme.

Active structural acoustic control of helicopter interior multifrequency noise using input-output-based hybrid control

This paper presents the recent advances in reduction of multifrequency noise inside helicopter cabin using an active structural acoustic control system, which is based on active gearbox struts technical approach. To attenuate the multifrequency gearbox vibrations and resulting noise, a new scheme of discrete model predictive sliding mode control has been proposed based on controlled auto-regressive moving average model. Its implementation only needs input/output data, hence a broader frequency range of controlled system is modelled and the burden on the state observer design is released. Furthermore, a new iteration form of the algorithm is designed, improving the developing efficiency and run speed. To verify the algorithm's effectiveness and self-adaptability, experiments of real-time active control are performed on a newly developed helicopter model system. The helicopter model can generate gear meshing vibration/noise similar to a real helicopter with specially designed gearbox and active struts. The algorithm's control abilities are sufficiently checked by single-input single-output and multiple-input multiple-output experiments via different feedback strategies progressively: (1) control gear meshing noise through attenuating vibrations at the key points on the transmission path, (2) directly control the gear meshing noise in the cabin using the actuators. Results confirm that the active control system is practical for cancelling multifrequency helicopter interior noise, which also weakens the frequency-modulation of the tones. For many cases, the attenuations of the measured noise exceed the level of 15dB, with maximum reduction reaching 31dB. Also, the control process is demonstrated to be smoother and faster.

Nonlinear Robust Controller for Miniature Helicopters Without Singularity

A singularity-free nonlinear robust controller is proposed for miniature helicopters to accomplish trajectory tracking missions. The helicopter model is decomposed into a cascaded structure with some unmodeled dynamics. The controller development applies the hierarchical strategy. First, a position loop controller is designed based on a saturation control scheme using differential hyperbolic tangent functions, and criteria for choosing controller parameters are established to ensure the finite-time position tracking and to obviate singularity in the command attitude extraction. Singularity may also occur during the attitude tracking when the pitch of the helicopter reaches $\pm \frac{\pi }{2}$ . Based on a barrier Lyapunov function, an attitude loop controller is put forward with a backstepping technique, and an initial condition constraint for the nonsingular attitude tracking is deduced. Simulations illustrate the applicability of the proposed controller.

Impact of turbulence and degraded visual environment on pilot workload

Offshore helicopter operations are frequently conducted in both turbulent and degraded visual environments (DVE). This investigation assesses the combined influence of turbulence and DVEs on pilot workload to identify first limits for safe operations. Flight tests using a simulation model of the research helicopter ACT/FHS (active control technology/flying helicopter simulator) flight mechanics model were conducted in the air vehicle simulator (AVES) at DLR Braunschweig. Tests were completed using four pilots, and results show the effects on pilot workload, task performance and control input activity. It was found that DVE and turbulence increase the workload and reduce task performance, but each in a different manner. Furthermore, the impact on control activity and pilot-induced oscillation tendencies are shown to have dependency upon the environmental conditions.

Transforming Civil Helicopters into Personal Aerial Vehicles: Modeling, Control, and Validation

This paper presents the implementation of robust control strategies to augment an identified state-space model of a civil light helicopter. The aim of this study is to augment the helicopter model to achieve response types and handling qualities of a new category of aircraft called personal aerial vehicles, which can be flown even by inexperienced pilots. Two control methods are considered to augment the helicopter model, and synthesis. Differences, advantages, and limitations of the implemented control architectures are highlighted with respect to the personal aerial vehicle reference dynamics in terms of robust stability, nominal performance, and handling qualities. Furthermore, results are presented of an experiment performed with the Max Planck Institute CyberMotion Simulator. The aim of the experiment is to assess the discrepancies between the two augmented systems and the personal aerial vehicle reference model. The experiment consists of piloted closed-loop control tasks performed by participants without any prior flight experience. The results show that the two implemented augmented systems allow inexperienced pilots to achieve workload and performance levels comparable to those defined for personal aerial vehicles.

Experimental investigations on active control of multifrequency helicopter vibrations using discrete model predictive sliding mode control

The presented experimental results illustrate the recent advances in the reduction of multifrequency vibrations of helicopter fuselage using an active control of structural response system. Recently, to cancel the multifrequency helicopter vibrations, a hybrid control approach has been proposed combining the filtered-x least mean square algorithm with a discrete model predictive sliding mode controller. To verify its effectiveness and self-adaptability, a set of active control experiments of structural response are conducted on a free–free elastic beam, which simulates a helicopter in flight. Considering that the helicopter vibrations in practical applications are much more complex, the further experiments of real-time active control are performed using a model helicopter test system. Higher discrete frequency components, which are actually of concern, are selected as the control objectives during the tests. The algorithm’s control effects are sufficiently checked by single-input single-output and multiple-input multiple-output tests under different excitation conditions. For many cases the attenuation of measured response exceed level of 20 dB, with maximum reduction reaching 34.1 dB. These two sets of tests confirm that the active control system is practical for canceling the multifrequency helicopter vibrations.

Main Rotor Trimming Using CFD Method in Forward Flight

In this paper a CFD trimming method was proposed and compared with wind tunnel experiment and blade element method. NASA’s generic ROBIN helicopter model was adopted for transient simulations to get the final main rotor trimming condition. Totally three steps were applied to CFD method, one being no cyclic pitch motion, one being pure longitudinal cyclic pitch motion and the other one being pure lateral cyclic pitch motion. At the same time, a simple linear equation system between the roll and pitching moment was adopted to get the final longitudinal and lateral cyclic pitch angles for the main rotor through these three steps. An overset grid approach was used where the volume around each blade was modeled in an individual overset grid region. The rotor rotation was resolved with 3 degree per time-step. Turbulence was modeled with the well known SST K-omega model with 2nd order convection. The helicopter was in straight forward flight with an advance ratio of μ=0.151. Three sources of stick angles, which are also calling rotor trimming angles, were shown and compared with each other. And the corresponding results were also plotted with type of history plot in CFD condition. In the simulations the results became quasi periodic after about 1.5 rotations and 4 rotor rotations were simulated for each case. The pitch moment coefficient and roll moment coefficient were all trimmed to zero by CFD trimming method while moments were not be removed thoroughly in other two sources condition. So, it is concluded that CFD trimming method has an ability to trim helicopter main rotor in forward flight.

Recursive fuzzy instrumental variable based evolving neuro-fuzzy identification for non-stationary dynamic system in a noisy environment

In this paper an online identification algorithm for instrumental variable based evolving neuro fuzzy modeling applied to dynamic systems in noisy environment, is proposed. The adopted methodology is based an online neuro-fuzzy inference system with Takagi–Sugeno evolving structure, which employs an adaptive distance norm based on the maximum likelihood criterion with instrumental variable recursive parameter estimation. The application and performance analysis of the proposed algorithm is based on black box modeling of a 2DOF Helicopter with errors in the variables.

Inverse dynamic modeling and analysis of a coaxial helicopter’s swashplate mechanism

In this paper, the application of the principle of virtual work to the inverse dynamic modelling of a coaxial helicopter’s swashplate mechanism is discussed. The swashplate mechanism is always designed as a configuration of multi closed-loop kinematic chains which dynamics is difficult to be analyzed, especially for coaxial layout. Firstly, the whole mechanism is decomposed into two submodules, the lower one and the middle-upper one, with driving torques of actuators and aerodynamic hinge moments as the input to each submodule respectively. After derivation of Jacobian matrices for the main components, the equations of equilibrium about the two submodules are constructed based on the principle of virtual work and then combined to achieve the complete inverse dynamic model. Furthermore, an integrated actuating system model for coaxial helicopter is established by combining the inverse dynamic model with mechanism kinematics, rotor dynamics, actuator dynamics and command allocation. Based on that model, an efficient approach to load analysis for actuators of helicopter in flight is created. Finally, a series of simulation for load analysis of actuators is performed.

Modeling, Trimming and Simulation of a Full Scale Helicopter

The complex configuration of helicopter guarantees that the vehicle modeling, trim and simulation are significantly more difficult than fixed-wing aircrafts. In this paper, general expressions for aerodynamic forces and moments, acting on helicopter due to its main and tail rotors at any flight conditions, are derived by using momentum and blade element theories. These complex expressions are inserted in the rigid body equations of motion, derived from Newton second law, to build a generic nonlinear mathematical model for single main and tail rotors helicopters; in order to obtain their responses to arbitrary control inputs. This model can be used in pilot training, control system design, and studying the helicopter stability characteristics. Trimming problem is solved at general flight conditions; arbitrary turn rate, flight path and side slip angles. The power required to fly helicopter at forward flight with several flight path angles is determined. The flight path angle required for helicopter autorotation condition is calculated at any forward speed. The mathematical model is solved by numerical integration (Runge-Kutta method) in the simulation code. The resulting trim conditions are verified by supplying the trim control inputs to the simulation code and verifying that the helicopter is flying in steady-state.

Modeling, Trimming and Simulation of a Full Scale Helicopter

Modeling, Trimming and Simulation of a Full Scale Helicopter

Development of a smoothed particle hydrodynamics method and its application to aircraft ditching simulations

The present study addresses the development and validation of a smoothed particle hydrodynamics (SPH) method, particularly to examine its feasibility and capability in hydrodynamics and dynamics of aircraft during ditching. The developed method solves the weakly compressible Navier–Stokes equations coupled with six-degree of freedom dynamics to achieve an accurate prediction of the interaction between the aircraft and the fluid. In this SPH method, a dummy particle wall-boundary condition is automatically implemented to meet the requirement of application on geometrically complex engineering problems. An efficient particle search strategy merging the ideal of Cell-linked list with Vertlet list is proposed to speed up the neighbor particles search process. The present SPH method uses an OpenMP memory-shared parallelization in conjunction with Z-curve reordering to accelerate the computation. Validations have been performed on several classic hydrodynamic problems, where good agreements were achieved via comparing with documented experimental results. The developed SPH method is applied to predict the ditching event of a complex helicopter model. Results demonstrate the ditching process, indicating that the method can be potentially used in aircraft ditching applications.

Variable-speed tail rotors for helicopters with variable-speed main rotors

ABSTRACTVariable tail rotor speed is investigated as a method for reducing tail rotor power, and improving helicopter performance. A helicopter model able to predict the main rotor and tail rotor powers is presented, and the flight test data of the UH-60A helicopter is used for validation. The predictions of the main and tail rotor powers are generally in good agreement with flight tests, which justifies the use of the present method in analysing main and tail rotors. Reducing the main rotor speed can result in lower main rotor power at certain flight conditions. However, it increases the main rotor torque and the corresponding required tail rotor thrust to trim, which then decreases the yaw control margin of the tail rotor. In hover, the tail rotor may not be able to provide enough thrust to counter the main rotor torque, if it is slowed to follow the main rotor speed. The main rotor speed corresponding to the minimum main rotor power increases, if the change of tail rotor power in hover is considered. As a helicopter translates to cruise, the induced power decreases, and the profile power increases, with the profile power dominating the tail rotor. Reducing the tail rotor speed in cruise reduces the profile power to give a 37% reduction in total tail rotor power and a 1.4% reduction to total helicopter power. In high-speed flight, varying the tail rotor speed is ineffective for power reduction. The power reduction obtained by the variable tail rotor speed is reduced for increased helicopter weight.

Experimental investigation of a helicopter rotor with Gurney flaps

ABSTRACTThe present work describes an experimental activity carried out to investigate the performance of Gurney flaps on a helicopter rotor model in hovering. The four blades of the articulated rotor model were equipped with Gurney flaps positioned at 95% of the aerofoil chord, spanning 14% of the rotor radius. The global aerodynamic loads and torque were measured for three Gurney flap configurations characterised by different heights. The global measurements showed an apparent benefit produced by Gurney flaps in terms of rotor performance with respect to the clean blade configuration. Particle image velocimetry surveys were also performed on the blade section at 65% of the rotor radius with and without the Gurney flaps. The local velocity data was used to complete the characterisation of the blade aerodynamic performance through the evaluation of the sectional aerodynamic loads using the the control volume approach.

Effect of the Simultaneous Variation in Blade Root Chord Length and Blade Taper on Helicopter Flight Control Effort

In this study, the effect of simultaneous variation in blade root chord length and blade taper on the control effort of helicopter flight control system (i.e., FCS) of a helicopter is investigated. Therefore, helicopter models (i.e., complex, control-oriented, and physics-based models) including the main physics and essential dynamics are used. The effect of simultaneous variation in the blade root chord length and blade taper (i.e., in both chordwise and lengthwise directions dependently) on the control effort of an FCS of a helicopter and also on the closed-loop responses is studied. Comparisons in terms of the control effort and peak values with and without variations in the blade root chord and blade taper changes are carried out. For helicopter FCS variance-constrained controllers, specific output variance-constrained controllers are beneficial.

Research on the Vibration Characteristics of a Compounded Periodic Strut Used for Helicopter Cabin Noise Reduction

A special periodic strut is developed to reduce the helicopter cabin noise in this paper. The strut exhibits unique dynamic characteristics which can isolate the gearbox vibrations from transferring to the fuselage and radiating noise. Modeling, simulation, and experimental research are carried out to explore its characteristics and performance. A theoretical model of the strut is firstly established, with particular emphasis on correlating the passband and stop band behaviors with the damping and a series of boundary conditions. Then, through simulations, it is shown that both the damping and boundary conditions have significant influences on the stop band, including the beginning and end frequencies and attenuation effects. Based on these analytical simulations, experimental analyses are conducted with the newly developed strut. Inspiring performances are validated under different conditions. Considering that the helicopter vibration in practical applications is much more complex, further experimental investigations are carried out on a helicopter model, which generates prominent gear mesh tones similar to a real helicopter. The experimental results show that the compounded periodic strut can significantly attenuate the vibrations transmitted to the fuselage. Compared with the plain strut, attenuations in excess of 40 dB are measured in the frequency range from 300 to 2000 Hz.

Assessment of Helicopter Pilot-in-the-Loop Models

The aim of this paper is the evaluation of several pilot models found in the literature, suited for helicopter pilot-assisted and pilot-induced oscillations analyses. Three main topics are discussed: (i) sensitivity of rotorcraft-pilot couplings simulations on the application of the different pilot models available in the literature; (ii) effect of vehicle modeling on active pilot modeling; (iii) effects of interactions between active and passive pilot models. The focus is on hovering flight, where a specific adverse rotorcraft-pilot coupling phenomenon, the vertical bounce, may occur. Pilot models are coupled with a comprehensive aeroservoelastic model of a mid-weight helicopter. The numerical investigations are performed in frequency domain, in terms of eigenanalysis and frequency response analysis.

バックステッピング法と最適サーボ系の組み合わせによるシングルロータヘリの高速旋回飛行制御系の設計とその実験的検証

Aiming at the realization of agile autonomous flight for the extension of the flying range of single-rotor helicopters in the limited flight time, the flight control system design and the control performance verification for the high-speed turn flight are conducted. To achieve the reference following control for large attitude angles and the rotor thrust, the control system is composed of the main-rotor thrust and body torque centralized controller based on the MIMO (Multiple-Input, Multiple-Output) helicopter model including the blade motion dynamics, rotor speed proportional-integral controller, the attitude controller designed via the backstepping method based on the quaternion attitude model, and the guidance controller based on translational motion model. Simulation results show that the hovering and 10-m/s straight cruise flight followed by over 45-degree bank turn is feasible by using the proposed control system. In addition, flight tests have been conducted using the experimental small unmanned electric helicopter equipped with the flight control computer, and the high speed turn agile autonomous flight was successful in the real environment.

Adaptive fuzzy tracking control for non-affine nonlinear yaw channel of unmanned aerial vehicle helicopter

The article deals with the problem of the yaw control of the unmanned aerial vehicle helicopter which is non-affine nonlinear by use of a novel projection-based adaptive fuzzy control approach. First, principle model and control design model of the yaw channel of the unmanned aerial vehicle helicopter are described. Then, a dynamic approximation technique is introduced to approach the non-affine model of the unmanned aerial vehicle helicopter into an affine model with variable parameters, which is applied to facilitate the design of nonlinear control scheme. Next, in the proposed controller, fuzzy controller is designed to deal with the unknown uncertainties and disturbances. Meanwhile, the projection-based adaptive law applied in fuzzy controller bounds the parameter estimation function and can also guarantee the robustness of the designed control scheme against uncertain disturbances. Moreover, the convergence and stability of the designed controller are proved by Lyapunov stability theory. Finally, the simulation results of the yaw channel of an unmanned aerial vehicle helicopter are performed to illustrate that the designed controller has good tracking performance, stability, and robustness under the condition of the time-vary uncertain disturbances.