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.

THE ANALYSIS OF THIN WALLED COMPOSITE LAMINATED HELICOPTER ROTOR WITH HIERARCHICAL WARPING FUNCTIONS AND FINITE ELEMENT METHOD

In the present paper, a series of hierarchical warping functions is developed to analyze the static and dynamic problems of thin walled composite laminated helicopter rotors composed of several layers with single closed cell. This method is the development and extension of the traditional constrained warping theory of thin walled metallic beams, which had been proved very successful since 1940s. The warping distribution along the perimeter of each layer is expanded into a series of successively corrective warping functions with the traditional warping function caused by free torsion or free beading as the first term, and is assumed to be piecewise linear along the thickness direction of layers. The governing equations are derived based upon the variational principle of minimum potential energy for static analysis and Rayleigh Quotient for free vibration analysis. Then the hierarchical finite element method. is introduced to form a,. numerical algorithm. Both static and natural vibration problems of sample box beams axe analyzed with the present method to show the main mechanical behavior of the thin walled composite laminated helicopter rotor.

Panel/full-span free-wake coupled method for unsteady aerodynamics of helicopter rotor blade

A full-span free-wake method is coupled with an unsteady panel method to accurately predict the unsteady aerodynamics of helicopter rotor blades in hover and forward flight. The unsteady potential-based panel method is used to consider aerodynamics of finite thickness multi-bladed rotors, and the full-span free-wake method is applied to simulating dynamics of rotor wake. These methods are tightly coupled through trailing-edge Kutta condition and by converting doublet-wake panels to full-span vortex filaments. A velocity-field integration technique is also adopted to overcome singularity problem during the interaction between the rotor wake and blades. Helicopter rotors including Caradonna–Tung, UH-60A, and AH-1G rotors, are simulated in hover and forward flight to validate the accuracy of this approach. The predicted aerodynamic loads of rotor blades agree well with available measured data and computational fluid dynamics （CFD） results, and the unsteady dynamics of rotor wake is also well simulated. Compared to CFD, the present method obtains accurate results more efficiently and is suitable to rotorcraft aeroelastic analysis.

A novel high-order scheme for numerical simulation of wake flow over helicopter rotors in hover

Accurate prediction of tip vortices is crucial for predicting the hovering performance of a helicopter rotor.A new high-order scheme(we call it WENO-K)proposed by our research group is employed to minimize numerical dissipation and extended to numerical simulation of unsteady compressible viscous flows dominated by tip vortices over hovering rotors.WENO-K is referred to as an adaptively optimized WENO scheme with Gauss-Kriging reconstruction,and its advantage is to reduce dissipation in smooth regions of flow while preserving high-resolution around discontinuities.Here WENO-K scheme is adopted to reconstruct left and right state values within the Roe Riemann solver updating the inviscid fluxes on a structured dynamic overset grid.To minimize the accuracy loss for high-order reconstruction on artificial boundaries of overset grid,a method of multilayer fringes is proposed to carry out interpolation between background grid and blade grid.Massively parallel computing considering automatic load balance on averagely partitioned overset grid is developed to reduce the wall-clock time of an unsteady simulation.Numerical results for Caradonna-Tung(C-T)rotor in hover at the conditions of subsonic and transonic tip Mach numbers show that the thrust coefficient error for the result of WENO-K scheme is no more than 3%.Compared with WENO-JS scheme,WENO-K scheme achieves about 40%improvement on accuracy of predicting rotor thrust with only 4.1%extra computational cost.More importantly,WENO-K scheme can capture more sophisticated unsteady flow structures and resolve tip vortices to a larger wake age with an increment of about 270°compared to WENO-JS scheme.

A fully automated framework for helicopter rotor blades design and analysis including aerodynamics,structure, and manufacturing

This study describes an integrated framework in which basic aerospace engineering aspects(performance, aerodynamics, and structure) and practical aspects(configuration visualization and manufacturing) are coupled and considered in one fully automated design optimization of rotor blades. A number of codes are developed to robustly perform estimation of helicopter configuration from sizing, performance analysis, trim analysis, to rotor blades configuration representation. These codes are then integrated with a two-dimensional airfoil analysis tool to fully design rotor blades configuration including rotor planform and airfoil shape for optimal aerodynamics in both hover and forward flights. A modular structure design methodology is developed for realistic composite rotor blades with a sophisticated cross-sectional geometry. A D-spar cross-sectional structure is chosen as a baseline. The framework is able to analyze all realistic inner configurations including thicknesses of D-spar, skin, web, number and ply angles of layers of each composite part,and materials. A number of codes and commercial software(ANSYS, Gridgen, VABS, Pre VABS,etc.) are implemented to automate the structural analysis from aerodynamic data processing to sectional properties and stress analysis. An integrated model for manufacturing cost estimation ofcomposite rotor blades developed at the Aerodynamic Analysis and Design Laboratory(AADL),Aerospace Information Engineering Department, Konkuk University is integrated into the framework to provide a rapid and dynamic feedback to configuration design. The integration of three modules has constructed a framework where the size of a helicopter, aerodynamic performance analysis, structure analysis, and manufacturing cost estimation could be quickly investigated. All aspects of a rotor blade including planform, airfoil shape, and inner structure are considered in a multidisciplinary design optimization without an exception of critical configuration.

EXPERIMENTAL INVESTIGATION OF AERODYNAMICINTERACTION EFFECT OF ROTOR WAKE ONFUSELAGE OF HELICOPTER

The interaction effect of rotor wake on fuselage of helicopter was investigated with experimental method. The results from experiment have proved that for the drag of fuselage the effect of rotor airflow is closely in connection with both the flight speed and the collective pitch of blades, while for the thrust and pitch moment of fuselage the collective pitch angle of blades plays more important role. A simple and effective computing method about aerodynamic interaction can be derived from the measured data. In order to implement the experiment, a fuselage model, a special sensor, the measurement and data acquisition and processing system were designed and manufactured according to the special requirements of this research project, thereby a good base was built up for carrying out experiments successfully with high quality.

High-fidelity simulation of blade vortex interaction of helicopter rotor based upon TENO scheme

Numerical simulation methods for unsteady vortex field of helicopter rotor with high resolution and low dissipation TENO8-AA primitive variables reconstruction schemes are established based on moving-embedded grid and Navier-Stokes equations.Firstly,the Targeted Essentially Non-Oscillatory(TENO)scheme are developed by employing ENO-like candidate stencil selection strategy,and the candidate stencil is adopted with optimal weight in smooth region while it is discarded completely in discontinuous region,which reduces the dissipation and dispersion errors and approaches better spectral properties.Then,the aerodynamic characteristics of Helishape-7A model rotor in Blade Vortex Interaction(BVI)state and the flowfield of Lynx rotor in hover are simulated,which validates that the blade tip vortex trajectory with larger wake age and more details of vortex can be captured by TENO8-AA scheme with only a quarter of grid points and half time comparing to WENO-JS scheme.Moreover,the simulation accuracy of thrust coefficient is improved by up to 36%.Finally,the analyses for BVI and aeroacoustic characteristics of Operational Loads Survey(OLS)rotor are conducted,and the different forms of interaction mechanism are explored,such as oblique and parallel interactions.The results indicate that TENO scheme not only ensures the resolution of simulation in discontinuous region,but also minimizes the numerical dissipation in smooth region dominated by blade tip vortex.Therefore,the acoustic pressure peak prediction error of rotor in BVI state is significantly reduced to 5.6%and 0.8%at two microphone locations,respectively.

RESEARCH ON SCHEME OF HIGH-SPEED ROTOR／WING TRANSITION HELICOPTER RD15

To analyze the existing schemes of high-speed rotorcrafts and some new technologies, a new conceptual sketch of the high-speed rotor/wing transition helicopter RD15 is proposed. The overall layout of the RD15 is given out and the transition process from the helicopter mode to the airplane mode is designed. The lift system consists of a circular disk-wing with four retractable blades. The technology of individual blade control is adopted for flight control in hover and low speed flight. The tail is a vectored thrust duct propeller. It can provide the anti-torque in hover, and offer the multi-directional controls and propulsion drive for the airplane mode flight. The aerodynamic characteristics and key technologies in the transition process for this layout, including the nose up angle of disk-wing, the length of the blade, rotation speed, pitch angle and other parameters, are theoretically ana lyzed and experimentally tested. Calculation and experiments show that the shift process of the lift, the power and controls are smooth, and the designed scheme is feasible.

Design, Control and Analysis of Low Cost Archetype Dual Rotor Helicopter for Educational Institution

This paper presents the design and development of low cost archetype dual rotor helicopter (LCADRH) for academic research in an educational institution. The LCADRH is installed with optical pitch encoder and yaw encoder which measure elevation and side to side motion of helicopter. The objective of the project is to design and integrate the helicopter with data acquisition board and sensors to provide hardware features, software support capability for its rapid real time measurement and control. The low cost designed LCADRH facilitates the academic research for students in the institution and is able to provide hands on training to understand the concept of nonlinearity, system modelled and unmodelled dynamics and uncertainty, modelling, simulation and control by doing practical experiments. The mathematical model of the LCADRH is derived using grey box modelling method. The control of LCADRH is challenging due to its nonlinearity and effect of strong coupling between aerodynamic forces and torques generated by the both pitch and yaw actuators. In closed loop position control of LCADRH, pitch and yaw axis motion is regulated using linear quadratic controller (LQR). Encouraging results are obtained both in simulation and hardware.

Analytical Model of Elastomeric Lag Damper Kinematic Coupling and Its Effect on Helicopter Air Resonance in Hover

For an elastomeric lag damper with nonlinear properties and kinematic couplings, its differential equation about equilibrium position was derived in a rotating frame. The equation was then transformed into a non-rotating frame by multi-blade transformation and incorporated into the rotor/airframe differential equations for eigen analysis. The effects of damper steady displacement and kinematic couplings on helicopter air resonance in hover were analyzed. The results demonstrate that the elastomeric damper can increase helicopter dynamic stability; however, its available damping will decrease as its steady displacement increased. For the notional rotor system, the damper steady displacement will decrease when kinematic couplings are introduced, and hence the regressive lag modal damping can be increased.

Autonomous Landing of Small Unmanned Aerial Rotorcraft Based on Monocular Vision in GPS-denied Area

Focusing on the low-precision attitude of a current small unmanned aerial rotorcraft at the landing stage, the present paper proposes a new attitude control method for the GPS-denied scenario based on the monocular vision. Primarily, a robust landmark detection technique is developed which leverages the well-documented merits of supporting vector machines (SVMs) to enable landmark detection. Then an algorithm of nonlinear optimization based on Newton iteration method for the attitude and position of camera is put forward to reduce the projection error and get an optimized solution. By introducing the wavelet analysis into the adaptive Kalman filter, the high frequency noise of vision is filtered out successfully. At last, automatic landing tests are performed to verify the method's feasibility and effectiveness. © 2014 Chinese Association of Automation.

Research on Aerodynamic Characteristics of Hovering Rotor Based on Leading Edge Droop

In view of the reduction of hovering efficiency near high tension when a helicopter rotor hovers,a numerical simulation method of lifting rotor hovering aerodynamic characteristics based on leading edge droop is established in this paper. It is dominated by Reynolds average N-S equation in integral form. Firstly,VR-12 airfoil is taken as the research object,and the influence of leading edge droop angle on the aerodynamic characteristics of two-dimensional airfoil is studied. Secondly,the modified 7 A rotor is taken as the research object,and the effects of different leading edge droop angles at the position of blade r/R=0.75—1 on the aerodynamic characteristics in hover are explored. It is found that the leading edge droop can significantly improve the aerodynamic characteristics of two-dimensional airfoil and three-dimensional hovering rotor near high angle of attack,and can effectively inhibit the generation of stall vortex.

ON INFLUENCE OF KINEMATICS TO EQUIVALENT LINEAR DAMPING OF HELICOPTER BLADE HYDRAULIC DAMPER

An analytical model of hydraulic damper was presented in forward flight accounting for pitch/flap/lag kinematic coupling and its nonlinear force-velocity curve. The fourth order Runge-Kutta was applied to calculate the damper axial velocity in time domain. Fourier series based moving block analysis was applied to calculate equivalent linear damping in terms of transient responses of damper axial velocity. Results indicate that equivalent linear damping will be significantly reduced if pitch/flap/lag kinematic coupling introduced for notional model and flight conditions.

Low Dissipation Simulation for Vortex Flowfield of Rotor in Hover Based upon Discontinuous Galerkin Method

The discontinuous Galerkin(DG) method is established and innovatively conducted on accurately simulating the evolution of blade-tip vortex and the aerodynamic characteristics of helicopter rotor. Firstly,the Reynolds-Averaged Navier-Stokes(RANS)equations in rotating reference frame are employed,and the embedded grid system is developed with the finite volume method(FVM)and the DG method conducted on the blade grid and background grid respectively. Besides,the Harten-Lax-Van Leer contact(HLLC)scheme with high-resolution and low-dissipation is employed for spatial discretization,and the explicit third-order Runge-Kutta scheme is used to accomplish the temporal discretization. Secondly,the aerodynamic characteristics and the evolution of blade-tip vortex for Caradonna-Tung rotor are simulated by the established CFD method,and the numerical results are in good agreement with experimental data,which well validates the accuracy of the DG method and shows the advantages of DG method on capturing the detailed blade-tip vortex compared with the FVM method. Finally,the evolution of tip vortex at different blade tip Mach numbers and collective pitches is discussed.

Data-driven active vibration control for helicopter with trailing-edge flaps using adaptive dynamic programming

The helicopter Trailing-Edge Flaps(TEFs)technology is one of the recent hot topics in morphing wing research.By employing controlled deflection,TEFs can effectively reduce the vibration level of helicopters.Thus,designing specific vibration reduction control methods for the helicopters equipped with trailing-edge flaps is of significant practical value.This paper studies the optimal control problem for helicopter-vibration systems with TEFs under the framework of adaptive dynamic programming combined with Reinforcement Learning(RL).Time-delay and disturbances,caused by complexity of helicopter dynamics,inevitably deteriorate the control performance of vibration reduction.To solve this problem,a zero-sum game formulation with a linear quadratic form for reducing vibration of helicopter systems is presented with a virtual predictor.In this context,an off-policy reinforcement learning algorithm is developed to determine the optimal control policy.The algorithm utilizes only vertical vibration load data to achieve a policy that reduces vibration,attains Nash equilibrium,and addresses disturbances while compensating for time-delay without knowledge of the dynamics of the helicopter system.The effectiveness of the proposed method is demonstrated in a virtual platform.

Identification method for helicopter flight dynamics modeling with rotor degrees of freedom

A comprehensive method based on system identification theory for helicopter flight dynamics modeling with rotor degrees of freedom is developed. A fully parameterized rotor flapping equation for identification purpose is derived without using any theoretical model, so the confidence of the identified model is increased, and then the 6 degrees of freedom rigid body model is extended to 9 degrees of freedom high-order model. Bode sensitivity function is derived to increase the accuracy of frequency spectra calculation which influences the accuracy of model parameter identification. Then a frequency domain identification algorithm is established. Acceleration technique is developed furthermore to increase calculation efficiency, and the total identification time is reduced by more than 50% using this technique. A comprehensive two-step method is established for helicopter high-order flight dynamics model identification which increases the numerical stability of model identification compared with single step algorithm. Application of the developed method to identify the flight dynamics model of BO 105 helicopter based on flight test data is implemented. A comparative study between the high-order model and rigid body model is performed at last. The results show that the developed method can be used for helicopter high-order flight dynamics model identification with high accuracy as well as efficiency, and the advantage of identified high-order model is very obvious compared with low-order model.

Investigation of rotor control system loads

This paper concentrates on the aeroelasticity analysis of rotor blade and rotor control systems. A new multi-body dynamics model is established to predict both rotor pitch link loads and swashplate servo loads. Two helicopter rotors of UH-60A and SA349/2, both operating in two critical flight conditions, high-speed flight and high-thrust flight, are studied. The analysis shows good agreements with the flight test data and the calculation results using CAMRAD II. The mechanisms of rotor control loads are then analyzed in details based on the present predictions and the flight test data. In high-speed conditions, the pitch link loads are dominated by the integral of blade pitching moments, which are generated by cyclic pitch control. In high-thrust conditions, the positive pitching loads in the advancing side are caused by high collective pitch angle, and dynamic stall in the retreating side excites high-frequency responses. The swashplate servo loads are predominated by the rotor pitch link loads, and the inertia of the swashplate has significant effects on high-frequency harmonics of the servo loads.

Structurally coupled characteristics of rotor blade geometrically exact formulation

In rotor dynamics,blades are normally modelled as a slender beam,in which elastic deformations are coupled with each other.To identify these coupling effects,new rigid-flexible structural model for helicopter rotor system is proposed in this paper.Finite rotations of the whole blade(on flapwise,lagwise,and torsional)are described as three global rigid degrees of freedom.The nonlinear deformation geometrics of the beam is built on geometrically exact beam theory.New expressions for blade strain energy,kinetic energy,and virtual work of various kinds of external forces are derived as functions of finite rotations and elastic deformations.To quantify the coupling characteristics,following the definition of coupling factor in electromagnetics,a new coupling factor between two modal components on each mode is introduced in modal analysis.Simulations show that the new structural model is highly capable of solving static and dynamic problems in rotor system and the maximum deformation that moderate deformation beam theory can predict might be 15%of beam length.After the new coupling factor is applied to study structurally coupled characteristics of rotor blade,it can be concluded that closeness of natural frequencies likely indicates considerable coupling between corresponding DOFs in structure.

FINITE ELEMENT ANALYSIS OF THE FLOW INDUCED BY ROTATING BLADES IN AN INCOMPRESSIBLE VISCOUS FLUID

Aerodynamic loads on a multi-bladed helicopter rotor in hovering flight were calculated by solving the three-dimensional incompressible Navier-Stokes equations. The rotor wake effects were accounted by the correction of local geometric angle of attack according to a free-wake modeling in addition to an empirical modification for the tip flow effect. The validity and efficiency of the present method were verified by the comparisons between numerical results and experimental data.

Numerical Simulation of the Tip Aerodynamics and Acoustics Test

The application of an efficient flow control system on helicopter rotor blades may lead to improved aerodynamic performance. Recently, our invention of Rod Vortex Generators(RVGs) has been analyzed for helicopter rotor blades in hover with success. As a step forward, the study has been extended to forward flight conditions. For this reason, a validation of the numerical modelling for a reference helicopter rotor(without flow control) is needed. The article presents a study of the flow-field of the AH-1G helicopter rotor in low-, medium- and high-speed forward flight. The CFD code FLOWer from DLR has proven to be a suitable tool for the aerodynamic analysis of the two-bladed rotor without any artificial wake modelling. It solves the URANS equations with LEA(Linear Explicit Algebraic stress) k-ω model using the chimera overlapping grids technique. Validation of the numerical model uses comparison with the detailed flight test data gathered by Cross J. L. and Watts M. E. during the Tip Aerodynamics and Acoustics Test(TAAT) conducted at NASA in 1981. Satisfactory agreements for all speed regimes and a presence of significant flow separation in high-speed forward flight suggest a possible benefit from the future implementation of RVGs. The numerical results based on the URANS approach are presented not only for a popular, low-speed case commonly used in rotorcraft community for CFD codes validation but preferably for medium- and high-speed test conditions that have not been published to date.