Adaptive predictor-based control for a helicopter system with input delays:Design and experiments

In this paper,we consider a 2-degrees-of-freedom(DOF)helicopter system subject to long input delays and uncertain system parameters.To address the challenges including unknown system parameters and input delays in control design,we develop an adaptive predictor-feedback control law to achieve trajectory tracking.Stability of the closed-loop system is further established,where the tracking errors are shown to converge towards zero.Through simulation and experiments on the helicopter system,we illustrate that tracking of a desired trajectory is achieved with the proposed control scheme.

Numerical Simulation of Aerodynamic Interaction Effects in Coaxial Compound Helicopters

The so-called coaxial compound helicopter features two rigid coaxial rotors,and possesses high-speed capabilities.Nevertheless,the small separation of the coaxial rotors causes severe aerodynamic interactions,which require careful analysis.In the present work,the aerodynamic interaction between the various helicopter components is investigated by means of a numerical method considering both hover and forward flight conditions.While a sliding mesh method is used to deal with the rotating coaxial rotors,the Reynolds-Averaged Navier-Stokes(RANS)equations are solved for the flow field.The Caradonna&Tung(CT)rotor and Harrington-2 coaxial rotor are considered to validate the numerical method.The results show that the aerodynamic interaction of the two rigid coaxial rotors significantly influences hover’s induced velocity and pressure distribution.In addition,the average thrust of an isolated coaxial rotor is smaller than that of the corresponding isolated single rotor.Compared with the isolated coaxial rotor,the existence of the fuselage results in an increment in the thrust of the rotors.Furthermore,these interactions between the components of the considered coaxial compound helicopter decay with an increase in the advance ratio.

Neural-Network-Based Adaptive Finite-Time Control for a Two-Degree-of-Freedom Helicopter System With an Event-Triggering Mechanism

Helicopter systems present numerous benefits over fixed-wing aircraft in several fields of application.Developing control schemes for improving the tracking accuracy of such systems is crucial.This paper proposes a neural-network(NN)-based adaptive finite-time control for a two-degree-of-freedom helicopter system.In particular,a radial basis function NN is adopted to solve uncertainty in the helicopter system.Furthermore,an event-triggering mechanism(ETM)with a switching threshold is proposed to alleviate the communication burden on the system.By proposing an adaptive parameter,a bounded estimation,and a smooth function approach,the effect of network measurement errors is effectively compensated for while simultaneously avoiding the Zeno phenomenon.Additionally,the developed adaptive finite-time control technique based on an NN guarantees finitetime convergence of the tracking error,thus enhancing the control accuracy of the system.In addition,the Lyapunov direct method demonstrates that the closed-loop system is semiglobally finite-time stable.Finally,simulation and experimental results show the effectiveness of the control strategy.

Numerical Study on Low-Reynolds Compressible Flows around Mars Helicopter Rotor Blade Airfoil

High-speed rotor rotation under the low-density condition creates a special low-Reynolds compressible flow around the rotor blade airfoil where the compressibility effect on the laminar separated shear layer occurs. However, the compressibility effect and shock wave generation associated with the increase in the Mach number (M) and the trend change due to their interference have not been clarified. The purpose is to clear the compressibility effect and its impact of shock wave generation on the flow field and aerodynamics. Therefore, we perform a two-dimensional unsteady calculation by Computational fluid dynamics (CFD) analysis using the CLF5605 airfoil used in the Mars helicopter Ingenuity, which succeeded in its first flight on Mars. The calculation conditions are set to the Reynolds number (Re) at 75% rotor span in hovering (Re = 15,400), and the Mach number was varied from incompressible (M = 0.2) to transonic (M = 1.2). The compressible fluid dynamics solver FaSTAR developed by the Japan aerospace exploration agency (JAXA) is used, and calculations are performed under multiple conditions in which the Mach number and angle of attack (α) are swept. The results show that a flow field is similar to that in the Earth’s atmosphere above M = 1.0, such as bow shock at the leading edge, whereas multiple λ-type shock waves are observed over the separated shear layer above α = 3° at M = 0.80. However, no significant difference is found in the Cp distribution around the airfoil between M = 0.6 and M = 0.8. From the results, it is found that multiple λ-type shock waves have no significant effect on the airfoil surface pressure distribution, the separated shear layer effect is dominant in the surface pressure change and aerodynamic characteristics.

Disturbance Observer-Based Safe Tracking Control for Unmanned Helicopters With Partial State Constraints and Disturbances

In this paper, a disturbance observer-based safe tracking control scheme is proposed for a medium-scale unmanned helicopter with rotor flapping dynamics in the presence of partial state constraints and unknown external disturbances. A safety protection algorithm is proposed to keep the constrained states within the given safe-set. A second-order disturbance observer technique is utilized to estimate the external disturbances. It is shown that the desired tracking performance of the controlled unmanned helicopter can be achieved with the application of the backstepping approach, dynamic surface control technique, and Lyapunov method. Finally, the availability of the proposed control scheme has been shown by simulation results.

CHARACTERISTICS OF FLOW IN A PARTIALLY DYNAMIC HELICOPTER INLET

An experimental study of the flow in a helicopter inlet with front output shaft and partial flow dynamic head is conducted in low speed wind tunnel. The flow characters of the inlet in the range of the yaw angle from 0~135°are presented in this paper. The static pressure distributions along the duct, distortions of the flow field at the outlet section and total pressure recovery coefficients are measured and analyzed. The results show that this type of inlet has high total pressure recovery coefficients at a wide range of yaw angle. The regions of local flow separation and distortion are closely related to the yaw angle. It′s also found that the outlet section has the best characteristics at sideslip, and sharply deteriorated characteristics at the yawed flight with a yaw angle of more than 90°

MULTI-BODY AEROELASTIC STABILITY ANALYSIS OF TILTROTOR AIRCRAFT IN HELICOPTER MODE

The muhi-body analysis of the aeroelastic stability of the tiltrotor aircraft is presented. Muhi-body dynamic differential equations are combined with the equations of the unsteady dynamic inflow model to establish the complete unsteadily aeroelastic coupling analytical model of the tiltrotor. The stability of the tiltrotor in the helicopter mode is analyzed aiming at a semi span soft-inplane tihrotor model with an elastic wing. Parametric effects of the lag stiffness of blades and the flight speed are analyzed. Numerical simulations demonstrate that the multibody analytical model can analyze the aeroelastic stability of the tiltrotor aircraft in the helicopter mode.

Dynamics Analysis of Close-coupling Multiple Helicopters System

The particularity and practicality of harmony operations of close-coupling multiple helicopters indicate that the researches on it are urgent and necessary, Using the model that describes two hovering helicopters carrying one heavy load, an inertia coordinate system and body coordinate systems of each sub-system are established. A nonlinear force model is established too. The equilibrium computation results can be regarded as the reference control inputs of the flight control system under hovering or low-speed flight condition. After the establishment of a translation kinematics model and a posture kinematics model, a coupling dynamics model of the multiple helicopter system is set up. The results can also be regarded as the base to analyze stabilization and design a controller for a close-coupling multiple helicopters harmony operation system.

DIRECT SELF-REPAIRING CONTROL FOR HELICOPTER VIA QUANTUM CONTROL AND ADAPTIVE COMPENSATOR

A direct self-repairing control approach is proposed for helicopter via quantum control techniques and adaptive compensator when some complex faults occur. For a linear varying-parameter helicopter control system, the model reference adaptive control law is designed and an adaptive compensator is used for improving its self-re- pairing capability. To enhance anti-interference capability of helicopter, quantum control feedforward is added be- tween fault and disturbance. Simulation results illustrate the effectiveness and feasibility of the approach.

DESIGN AND IMPLEMENTATION OF GUST RESPONSEALLEVIATION CONTROL SYSTEM FOR HELICOPTERS

Gust response alleviation is very important for helicopters which have strong coupling and vibration. Gust disturbance not only influences the ride quality and the precision of the weapon delivery, but also affects to the structural fatigue load and the strength. The method of an optimal control law to suppress the gust disturbance for helicopters is presented. The optimization requires the minimization of the vertical overload at the pilot′s seat, the attitude variation and the control energy consumption under the gust disturbance. Based on the original control system, the new system can be easily realized by adding a vertical speed feedback passage. In order to develop the real-time operational flight control system, the optimized control law is written in C language. The hybrid simulations prove that the performance of gust response alleviation and the efficiency of digitalization are satisfactory.

ESPRIT-Based Feature Extraction of Helicopter Acoustic Signal

Aim To extract harmonic frequencies of helicopter acoustic signal as features for hel icopter identification. Methods Estimation of signal parameters via rotational invariance techniques(ESPRIT) was selected to extract harmonic frequencies from really measured helicopter acoustic signal and an algorithm based on the SVD TLS was used. Results ESPRIT correctly extracted harmonic frequencies of helicopter using the data of limited length under the variousflight conditions. Conclusion ESPRIT is an effective method of extracting harmonic frequencies and using harmonic frequencies of helicopter acoustic signal to recognize helicopter is feasible.

APPLICATION OF FUZZY INFERENCE IN IDENTIFICATION OF HELICOPTER MODEL

Helicopter mathematical model mainly depends on design helicopter control system, flight simulator, and real time control simulation system. But it is difficult to establish a helicopter flight dynamics mathematical model that has features such as rapidness, reliability and precision, because there is no unique and precise expression to some sophisticated phenomenon of helicopter. In this paper a fuzzy helicopter flight model is constructed based on the flight experimental data. The fuzzy model, which is identified by fuzzy inference, has characteristics of computed rapidness and high precision. In order to guarantee the precision of the identified fuzzy model, a new method is adopted to handle the conflict fuzzy rules. Additionally, using fuzzy clustering technology can effectively reduce the number of rules of fuzzy model, namely, the order of the fuzzy model. The simulation results indicate that the method of this paper is effective and feasible.

ADAPTIVE NEURAL NETWORK ATTITUDE CONTROL FOR UNMANNED HELICOPTER

Adaptive flight control technology, feedback linearization, model inversion theory are reviewed and the error dynamic characteristics are analyzed, and an adaptive on-line neural network attitude control system is presented. The model inversion is under the hover condition. And the adaptive control law based on the neural network is designed to guarantee the boundedness of tracking error and control signals. Simulation results demonstrate that the nonlinear neural network augmented model inversion can self-adapt to the uncertainty and modeling errors of unmanned helicopters. Results are compared while the parameters of PD controller and robustness items are changed.

EXPERIMENTAL STUDY ON NO TAIL ROTOR (NOTAR) HELICOPTER

The circulation contr ol tail boom is a new technology used as an anti-torque system for the no tail rotor helicopter. In order to study the effect of the different parameters on th e circulation control tail boom, two models with the different circulation contr ol tail booms are designed and constructed. The experiments are introduced in th is paper by measuring the pressure distribution on the models in wind tunnel and under main rotor, respectively. By analyzing and discussing the experimental re sults, the effects of the momentum coefficient, the slot′s geometry parameters and the freestream velocity on the aerodynamic forces of the circulation control tail boom have been expounded, and the reasonable explanations about the test r esult have been given in theory.

AERODYNAMIC ANALYSIS OF HELICOPTER SHROUDED TAIL ROTOR BY MOMENTUM THEORY

In this paper, in order to clarify the gains of the shroud in the shrouded tail rotor system, a thrust division factor q , which represents the ratio of the shroud thrust to the total thrust of the shrouded tail rotor, is introduced. With the help of q , the slipstream theory for the static and axial flow states of the shrouded tail rotor are fully derived. Based on the sliptream theory, the variations of the thrust, power and disk area against q for different cases are emphatically analysed and the comparisons between a shrouded tail rotor and an isolated one are made. It is shown that, although the shroud can provide as much as 50% of the total thrust of shrouded tail rotor for the static state, the thrust gains of the shroud rapidly decrease for the axial flow state, which depends on the flow velocity ratio.

A NEW AEROMECHANICAL STABILITY ANALYSIS METHODOLOGY FOR COUPLED ROTOR/FUSELAGE SYSTEM OF HELICOPTERS

The aeromechanical st ability for the coupled rotor/fuselage system of helicopters in forward flight i s investigated. The periodic time-varying equations of motion are developed thr ough building a new 24DOF coupled rigid/elastic blended element based on the fle xible multibody system theory in this paper. It accounts for the effects of prec one, sweep, and the moderately large elastic deflections on the blade and elasti city of shaft and fuselage of the helicopter. The dynamic coupling between the r igid motion of blades about the flap, lag and pitch hinges of articulated rotor and moderately large elastic deflections are included. There is no restriction o n the rotation amplitudes of flap, lag and pitch in the formulation. The stabili ty of periodic solution is studied using the Floquet theory. The transition matr ix is calculated by the Newmark integration method. The aeromechanical stability of a new helicopter is studied. The results show that it is stable in the given forward flight. But the instability arises with the decrease of the bending and torsion stiffness of the shaft.

COMPARISON BETWEEN TWO ANALYTICAL MODELS OF HELICOPTER GROUND RESONANCE

The 2-dimensional equivalent model used in helicopter ground resonance analysis is convenient for its simplicity and clarity. The equivalent body mass, the stiffness and the damping are derived from the modal charaCteristics. In this paper, a 3-dimensional space model is constructed to analyse the ground resonance.There is little difference between the calculated results of the 2-dimensional equivalent model and those of the 3-dimensional space model. Hence, the 2-dimensional model is verified for the practical application. Generally, the linear lead-lag damping of the equivalent rotor blade at the modal frequency is derived from the lead-lag damper characteristics by a procedure of iteration. The transformation of the imaginary part of the complex modulus of the viscoelastic damper into a linear velocity damping is complicated in analysis. A method to use the complex modulus directly in the ground resonance analysis is derived and verified.

Simulation Study on the Feasibility of Gun-Launched Missile Against Attack Helicopters

The feasibility of providing the tank a limited anti helicopter ability with gun launched missile is studied. A type of simulation model of gun launched missile against attack helicopters is established. The simulation and the parameter optimization of missile control system under various circumstances are done. The gun launched missile can directly hit the helicopters in the typical tracks, all the missdistances are less than 1?m and the maximum overload is less than available overload. Gun launched missile is a feasible choice for tanks against attack helicopters.

DIAGNOSIS OF DAMPING FAULTS IN HELICOPTER ROTOR HUB BASED ON FUSELAGE VIBRATIONS

Damping faults in a helicopter rotor hub are diagnosed by using vibration signals from the fuselage. Faults include the defective lag damper and raspings in its flap and feathering hinges. Experiments on the diagnosis of three faults are carried out on a rotor test rig with the chosen fault each time. Fuselage vibration signals from specified locations are measured and analyzed by the fast Fourier transform in the frequency domain. It is demonstrated that fuselage vibration frequency spectra induced by three faults are different from each other. The probabilistic neural network （PNN） is adopted to detect three faults. Results show that it is feasible to diagnose three faults only using fuselage vibration data.

Modeling and Backstepping-based Nonlinear Control Strategy for a 6 DOF Quadrotor Helicopter

In this article, a nonlinear model of an underactuated six degrees of freedom （6 DOF） quadrotor helicopter is derived on the basis of the Newton-Euler formalism. The derivation comprises determining equations of the motion of the quadrotor in three dimensions and approximating the actuation forces through the modeling of aerodynamic coefficients and electric motor dynamics. The derived model composed of translational and rotational subsystems is dynamically unstable, so a sequential nonlinear control strategy is used. The control strategy includes feedback linearization coupled with a PD controller for the translational subsystem and a backstepping-based PID nonlinear controller for the rotational subsystem of the quadrotor. The performances of the nonlinear control method are evaluated by nonlinear simulation and the results demonstrate the effectiveness of the proposed control strategy for the quadrotor helicopter in quasi-stationary flights.