Helicopter Vibration Reduction Research Articles

회전익항공기 진동저감장치 장착부 구조 개선 연구

회전익항공기 진동저감장치 장착부 구조 개선 연구

Optimal Sizing of Trailing Edge Flaps for Helicopter Vibration Reduction: A Response Surface Approach

On-blade active trailing edge flaps are considered for helicopter vibration reduction. A Pareto optimal multi-objective optimization approach is used to obtain the spanwise and chordwise dimensions of single and dual trailing- edge flaps. The objective of this study is to simultaneously achieve minimum hub vibration levels and minimize the flap actuation power requirement. Helicopter aeroelastic analysis is used in conjunction with an optimal controller to minimize hub vibration and flap power levels. It is found that nonlinear polynomial response surfaces based on L9 orthogonal array could adequately describe both the objectives. An intense sampling of the surrogate objective space is used to obtain optimal design points for the mutually conflicting objectives. A Pareto optimal design reduces vibration levels by about 70% and also reduces flap actuation power requirement by about 20%, in comparison to the initial design. The present study leads to an optimal design of trailing edge flaps which expend less power and can be further explored for self-powered control surfaces towards a self-powered smart blade design concept. The optimal designs are robust to uncertainty introduced by manufacturing.

Higher Harmonic Control: A Historical Perspective

Higher harmonic control (HHC) is an approach for achieving reduced helicopter vibration by controlling the vibratory rotor airloads in such a way that the fuselage excitation is minimized. This paper is a historical look at how a program aimed at helicopter vibration reduction started as an outgrowth of fixed wing flutter suppression at NASA Langley Research Center, proved the HHC concept on aeroelastically scaled wind tunnel models and went on to demonstrate viability in full-scale flight testing on the OH-6A helicopter in 1982. Following the OH-6A flight tests, the helicopter research community was stimulated to prove the effectiveness of HHC on different configurations through analysis, wind tunnel tests, and flight tests. All of these investigations have shown HHC to be effective in reducing vibration to levels not attainable with conventional vibration control methods and without any detrimental side effects. HHC development has progressed to the point that the technology provides one more option to address the ever-present vibration problem in helicopters. The literature demonstrates that helicopter ride quality equivalent to that of fixed wing aircraft is attainable with application of HHC.

Comprehensive Numerical Assessment of Rotorcraft Vibration and Noise Control Using Microflaps

Active Gurney flaps, or microflaps, are studied to determine their effectiveness in reducing noise and vibration in rotorcraft as well as improving rotor performance. The effectiveness of the microflap is examined using a comprehensive rotorcraft simulation code. The aerodynamic properties of the microflap are modeled using a nonlinear computational-fluid-dynamics-based reduced-order aerodynamic model that takes into account unsteadiness, compressibility, and time-varying freestream effects. Active control studies are conducted on a hingeless rotor configuration resembling the Messerschmitt-Bölkow-Blohm BO-105, using various spanwise microflap configurations, including single, dual, and segmented five-microflap configurations. Results indicate that the microflap is capable of substantial reductions in blade–vortex interaction noise ranging from 3 to 6 dB. Vibration reduction ranging from 70 to 90% is also demonstrated. The effect of vibration reduction on noise and vice versa is also examined, and it was found that reduction in one objective is accompanied by an increase in the other, a trend also observed when using other active control approaches. Finally, the microflap is considered for combined vibration reduction and performance enhancement at a high-speed cruise flight condition. The results clearly indicate that the microflaps are very effective for both noise and vibration reduction in helicopters, and they also have potential for rotor performance enhancement.

그래프 분석과 호감도 함수를 이용한 헬리콥터 개발형상의 진동저감 사례

Purpose: This paper presents graphic methods and desirability function approach to determine best vibration reducing configuration for Surion helicopter. Many flight tests were executed and nine vibration levels in cockpit, cabin, and engine room were measured in each test and analyzed to find optimal configuration. Methods: Graphic analysis methods such as matrix, scatter, and box plots are used to identify better vibration-reducing flight test conditions. As an integrated measure of the performance of 9 vibration levels desirability function approach is adopted. Results: Three vibration reducing configurations are found to be proper and one configuration is recommended. Conclusion: It is expected to be helpful to adopt graphic and desirability function methods presented in this paper in developing new products or systems like helicopters. For efficient and effective flight testing of helicopters, it will be necessary to have consistently homogeneous environment for flight testing and applYing design of experiments techniques and analyzing test data. Key words: Helicopter Vibration Reduction, Product Development, Graphic Analysis, Multiple Responses, Desirability Function

Aeroelastic control of non-rotating and rotating wings using the dynamic stiffness modulation principle via piezoelectric actuators

Carleton University’s Rotorcraft Research Group is working on the development of an active rotor control system that incorporates a mechanism for helicopter blade pitch dynamic stiffness modulation at the root, the Active Pitch Link. This system overcomes stroke limitations of smart material and attains superior performance for helicopter rotor-induced vibration reduction. The system was tested at the whirl tower facility and this article reports the achievements obtained with a dynamically similar hinged rotor blade model. Up to 100% reduction in the transmitted loads occurred at the target 2/rev frequency when the blade was excited by a transversal jet to mimic the asymmetric flow of the helicopter rotor in forward flight. An open-loop control algorithm optimized to a target higher-harmonic frequency of the rotor also minimized the impact on the rotor fundamental cyclic control frequency at 1/rev. In another experiment at University of São Paulo, semi-passive control techniques using shunted piezoelectric materials were investigated for the aeroelastic control of fixed wings. Flutter oscillations of a typical section were controlled out over a range of airflow speeds. Finally, the similarity between both control techniques is discussed and recognized that they are based on a dynamic stiffness modulation control principle.

Black-box MIMO model identification for structural vibration reduction in helicopters

This work addresses the identification of a control-oriented multivariable model of the structural response of a helicopter, with reference to the Agusta A109 helicopter. The estimated model will be used to design an active vibration control system to reduce the vibrations transmitted to the cabin by the main gearbox. The considered system consists of a set of piezoelectric actuators mounted at the interface between the gearbox and the cabin and accelerometers located on the cabin roof. Identification data are generated by independent excitation of each individual actuator. Model identification is carried out using a subspace identification algorithm.

Robust harmonic control: an application to structural vibration reduction in helicopters

This work addresses the problem of designing robust multivariable controllers for the attenuation of the structural response of a helicopter, with reference to the Agusta A109 helicopter. The considered configuration consists of a set of piezoelectric actuators and accelerometers located on the fuselage of the helicopter. The robust harmonic control approach is adopted, based on an data-driven mathematical model of the system.

Robust design of multiple trailing edge flaps for helicopter vibration reduction: A multi-objective bat algorithm approach

The objective of this study is to determine an optimal trailing edge flap configuration and flap location to achieve minimum hub vibration levels and flap actuation power simultaneously. An aeroelastic analysis of a soft in-plane four-bladed rotor is performed in conjunction with optimal control. A second-order polynomial response surface based on an orthogonal array (OA) with 3-level design describes both the objectives adequately. Two new orthogonal arrays called MGB2P-OA and MGB4P-OA are proposed to generate nonlinear response surfaces with all interaction terms for two and four parameters, respectively. A multi-objective bat algorithm (MOBA) approach is used to obtain the optimal design point for the mutually conflicting objectives. MOBA is a recently developed nature-inspired metaheuristic optimization algorithm that is based on the echolocation behaviour of bats. It is found that MOBA inspired Pareto optimal trailing edge flap design reduces vibration levels by 73% and flap actuation power by 27% in comparison with the baseline design.

A feasibility study of a post-buckled beam for actuating helicopter trailing edge flap

Active trailing edge flaps (TEFs) are one of the most promising devices for helicopter vibration reduction. Smart actuators such as the piezoelectric stack actuators (PEAs) are used for TEF actuation. PEAs possess high energy density and have large force in dynamic condition but are limited to small displacements. In this investigation, we study a linear to rotary motion amplification mechanism (AM-2) based on a pinned-pinned post-buckled beam to actuate trailing edge flaps. A linear motion amplification mechanism is developed and coupled with AM-2 to amplify angular flap deflections. Experiments are conducted on bench top-test setup, and maximum flap angle deflections of the order of 12A degrees are achieved in the static case. An aeroelastic analysis is performed and 91 % reduction in helicopter vibration is obtained with multiharmonic control inputs.

Force control of semi‐active valve lag dampers for vibration reduction in helicopters

This study considers the design of a closed-loop force-tracking system for a semi-active damper, designed to be used to reduce in-plane vibrations caused by helicopter rotor blades during steady-state forward flight conditions. The study describes the development of the control law and includes details of (i) how the initial mathematical model of the system is adapted for controller design; (ii) how a non-linear dynamic inversion (NDI) control law is modified into a form suitable for implementation; and (iii) how the free parameters in the NDI controller can be optimised for various different operational modes. The success of the approach is demonstrated through both force-tracking simulations and also more comprehensive tests in which the controller is incorporated into a large-scale vibration simulation of the AgustaWestland 101 helicopter. The results show that the NDI-based controller can provide a satisfactory level of performance and hence greatly assist in the reduction of unwanted vibrations.

Online Adaptive Fuzzy Neural Identification of a Piezoelectric Tube Actuator System

A coupled actuator-flap-circuit system model and its online identification are presented. The coupled system consists of a piezoelectric tube actuator, a trailing-edge flap, and a series R-L-C circuit. The properties of the coupled system are examined using a Mach-scaled rotor simulation on hovering state. According to the high nonlinear hysteretic characteristics of the coupled system, the generalized dynamic fuzzy neural networks (GD-FNN) implementing Takagi-Sugeno-Kang (TSK) fuzzy systems based on extended ellipsoidal radial basis function (EBF) neural network is used to identify the coupled system. The structures and parameters are adaptive adjusted during the learning process, and don’t need too much expert experiences. Simulation studies show that the piezoelectric tube actuator has high authority with a broad frequency bandwidth, satisfies the requirements for helicopter vibration reduction; GD-FNN has a high learning speed, the final networks have a parsimonious network structure and generalize well, possessing broad application prospects in the helicopter vibration reduction.

Effect of Piezoceramic Actuator Hysteresis on Helicopter Vibration and Noise Reduction

The objective of this studywas to assess the importance of actuator hysteresis on vibration and noise reduction in a rotorcraft equippedwith an actively controlledflap (ACF) system.Ahysteresismodel, based ona generalized formof the classical Preisach model, was developed using experimental data. The model displays good agreement with experimental data for inputs representative of those used for vibration and noise reduction in full-scale rotors. The hysteresis model has been incorporated into the AVINOR code to determine the effect of actuator hysteresis on vibration and noise reduction in both open and closed-loop modes of operation. It was found that the hysteresis level of a representative piezoceramic actuator does not produce a significant performance degradation of theACF system for vibration reduction under blade-vortex interaction (BVI) and dynamic stall conditions. For open-loop BVI noise control, the presence of hysteresis produces a significant effect on themaximumnoise reduction obtained using 2 per revolution (rev) and 5=rev harmonic inputs. For closed-loop BVI noise control, hysteresis introduces a substantial difference inflap deflection time histories; however, the overall noise reduction performance of theACF system is not affected because the controller compensates for actuator hysteresis.

Surrogate based design optimisation of composite aerofoil cross-section for helicopter vibration reduction

AbstractDesign optimisation of a helicopter rotor blade is performed. The objective is to reduce helicopter vibration and constraints are put on frequencies and aeroelastic stability. The ply angles of the D-spar and skin of the composite rotor blade with NACA 0015 aerofoil section are considered as design variables. Polynomial response surfaces and space filling experimental designs are used to generate surrogate models of the objective function with respect to cross-section properties. The stacking sequence corresponding to the optimal cross-section is found using a real-coded genetic algorithm. Ply angle discretisation of 1°, 15°, 30° and 45° are used. The mean value of the objective function is used to find the optimal blade designs and the resulting designs are tested for variance. The optimal designs show a vibration reduction of 26% to 33% from the baseline design. A substantial reduction in vibration and an aeroelastically stable blade is obtained even after accounting for composite material uncertainty.

능동 비틀림 제어기법을 이용한 복합재료 로터 블레이드의 진동 억제

능동 비틀림 제어기법을 이용한 복합재료 로터 블레이드의 헬리콥터 진동억제에 대한 수치연구를 수행하였다. 허브에 작용하는 진동하중 억제를 위해 복합재료 블레이드의 탄성 연계와 함께 압전 소재의 전단변형 메커니즘을 이용하였다. 로터 블레이드는 표면에 압전 작동기를 부착한 박벽 상자형 단면을 갖는 복합재료 보로 모델링하였다. 회전익에 대한 지배 운동방정식은 Hamilton 원리를 이용하여 구성하였고, 공력하중은 자유후류모델을 포함하는 비정상 공력 이론을 이용하여 구했다. 다양한 탄성연계 적층과 능동 작동기를 부착한 복합재료 블레이드에 대해 허브진동 하중 특성을 고찰하였다. 수치해석 결과 최적 제어 알고리듬을 적용하여 <TEX>$N_b$</TEX>/rev 진동하중을 대폭 줄일 수 있음을 보였다. In this study, an assessment is made for the helicopter vibration reduction of composite rotor blades using an active twist control concept. The piezoceramic shear actuation mechanism along with elastic couplings of composite blades is used for vibration reduction. The rotor blades are modeled as composite box-beams with actuator layers bonded on the outer surfaces of the thin-walled section. The governing equations of motion for helicopter blades are obtained using Hamilton's principle. A time domain unsteady aerodynamic theory with free wake model is used to obtain the airloads. Various rotor configurations with different elastic couplings with appropriate actuator placement are used to investigate the hub vibration characteristics. Numerical results show that a substantial reduction of <TEX>$N_b$</TEX>/rev hub vibration can be achieved using the optimal control algorithm.

Helicopter Vibration Reduction throughout the Entire Flight Envelope Using Surrogate-Based Optimization

Helicopter Vibration Reduction throughout the Entire Flight Envelope Using Surrogate-Based Optimization

Single-crystal-material-based induced-shear actuation for vibration reduction of helicopterswith composite rotor system

In this study, an assessment is made for the helicopter vibration reduction of compositerotor blades using an active twist control concept. Special focus is given to the feasibility ofimplementing the benefits of the shear actuation mechanism along with elastic couplings ofcomposite blades for achieving maximum vibration reduction. The governingequations of motion for composite rotor blades with surface bonded piezoceramicactuators are obtained using Hamilton’s principle. The equations are then solvedfor dynamic response using finite element discretization in the spatial and timedomains. A time domain unsteady aerodynamic theory with free wake modelis used to obtain the airloads. A newly developed single-crystal piezoceramicmaterial is introduced as an actuator material to exploit its superior shear actuationauthority. Seven rotor blades with different elastic couplings representing stiffnessproperties similar to stiff-in-plane rotor blades are used to investigate the hubvibration characteristics. The rotor blades are modeled as a box beam with actuatorlayers bonded on the outer surface of the top and bottom of the box section.Numerical results show that a notable vibration reduction can be achieved forall the combinations of composite rotor blades. This investigation also bringsout the effect of different elastic couplings on various vibration-reduction-relatedparameters which could be useful for the optimal design of composite helicopter blades.

Performance sensitivity of helicopter global and local optimal harmonic vibration controller

A comparison of local and global controllers is conducted for helicopter vibration reduction using actively controlled single/dual trailing-edge flaps. It is found that for high-speed flight conditions, both local and global controllers perform equally well. However, for low-speed flight, local controller gives a marginally better optimal solution compared to the global controller but at the cost of much greater computation time. An objective function containing the hub vibratory loads is minimized using the global controller. The control effort is weighted such that each flap is used to complete authority. A reduction of 73% and 75% in hub vibration is obtained at high-speed flight using single and dual flap configurations, respectively. At low-speed flight, single and dual flap configurations gave 52% and 63% reduction in vibration, respectively. Numerical studies using an aeroelastic simulation show that the vibration control system performs well up to ±5 degrees error in flap phasing from optimal phase. The control law is insensitive to small perturbations (up to ±0.3 deg) in the optimal flap harmonics and is fairly insensitive to stiffness and mass changes of the rotor blade as well as noise in the measured hub load data. Dual flap configuration was found to be more insensitive to errors in optimal flap control harmonics than single flap configuration.

Design and Manufacturing of a Model-scale Active Twist Rotor Prototype Blade

The design and manufacturing of an active twist rotor blade for vibration reduction in helicopters are presented. The rotor blade is integrally twisted by direct strain actuation through embedded piezoelectric fiber composite actuators distributed along the span of the blade. Highlights of the analysis formulation used to design this type of active blade are presented. The requirements for the prototype blade, along with the final design results are also presented. Detailed aspects of its manufacturing are described. Experimental structural characteristics of the prototype blade compare well with design goals, and bench actuation tests characterize its basic actuation performance. The design and manufacturing processes permit the realization of an active blade that satisfies a given set of design requirements. This is used to later develop a fully active rotor blade system.

Rotor Performance Enhancement and Vibration Reduction in Presence of Dynamic Stall Using Actively Controlled Flaps

A computational study of helicopter vibration and rotor shaft power reduction is conducted using activelycontrolled trailing-edge flaps (ACFs), implemented in both single and dual flap configurations. Simultaneous vibration reduction and performance enhancement is demonstrated under level flight condition at high advance ratios, where dynamic stall effects are significant. Power reduction is achieved using the adaptive Higher Harmonic Control (HHC) algorithm in closed loop, with 2-5/rev flap control harmonics. This approach is compared with an off-line, nonlinear optimizer available in MATLAB, and favorable comparisons are obtained. A parametric study of flap spanwise location is also conducted to determine its optimal location for power reduction. The effectiveness of the ACF approach for power as well as simultaneous vibration and power reduction is also compared with conventional individual blade control (IBC) approach. Rotor power reduction and simultaneous reduction of vibration and power are shown to be larger at higher rotor thrust and advance ratio. Finally, the effect of active flap on dynamic stall is examined to determine the mechanisms of rotor power reduction. The simulation results clearly demonstrate the potential of the ACF system for power reduction as well as simultaneous vibration and power reduction. Nomenclature c Blade chord cc Flap chord CT Rotor thrust coefficient D Matrix defined as TTQT+R fb(.) Blade equations of motion ft(.) Trim equations FHX4,FHY4, FHZ4 4/rev hub shears, nondimensionalized by MbΩR MHX4,MHY4, MHZ4 4/rev hub moments, nondimensionalized by MbΩR J Quadratic-form objective function to be minimized Presented at the American Helicopter Society 62nd Annual Forum, Phoenix, AZ, June 9-11, 2006. 1 Mb Blade mass mc Flap mass per unit length, nondimensionalized by Mb/R MHz1 Yawing moment about rotor hub Nb Number of rotor blades PR Rotor shaft power, nondimensionalized by MbΩR qt Vector of trim variables qti Vector of trim variables at i-th control step Q Weighting matrix for objectives to be minimized R Weighting matrix on control input R Rotor blade radius Rt Trim residuals vector T Sensitivity, transfer matrix between control inputs and objective function uk Control input vector, kth control step uk,opt Optimum value of control input vector xc Spanwise location of center of control surface zk Objective vector, kth control step αR Rotor shaft angle of attack δ f Flap deflection angle ∆Cd,flap Additional drag due to flap deflection δNc,δNs N/rev cosine and sine amplitude of δ βp Blade precone angle γ Lock number μ Helicopter advance ratio ωF , ωL, ωT Blade flap, lead-lag and torsional natural frequencies Ω Rotor angular speed ψ Rotor azimuth angle φR Lateral roll angle θ0,θ1c,θ1s Collective and cyclic pitch components θ0t Tail rotor collective pitch θtw Built-in twist angle σ Rotor solidity Introduction and Background Specifications for noise and vibration levels in rotorcraft are continuously increasing in stringency, thus motivating research related to active noise and vibration reduction. Desirable vibration levels have been identified to be below 0.05g to provide passengers with “jet smooth” ride (Ref. 1). A number of active control techniques have emerged for effective vibration reduction (Refs. 1,2), as illustrated schematically in Fig. 1. These approaches generally fall into one of two categories: (a) active control approaches aimed at reducing vibrations in the rotor before they propagate into the fuselage, and (b) active control approaches