Helicopter Vibratory Loads Alleviation through Combined Action of Trailing-Edge Flap and Variable-Stiffness Devices (original) (raw)
Related papers
The aim of the present paper is the examination of control systems for the alleviation of vibrating loads arising at the rotor hub of helicopters in forward flight. These are obtained by combining the effects of actuated trailing edge flaps with the action of controllable-stiffness devices located at the pitch link and roots of the blades. Control laws are obtained through an optimal control procedure yielding the best compromise between control effectiveness and control effort. The numerical investigation concerns the analysis of performance and robustness of the control techniques examined through application to a four bladed helicopter rotor in level flight. The identification of the most efficient control configuration is also attempted.
helicopter vibratory loads and vibrations reduction using
This paper studies vibratory hub loads and cabin vibrations reduction using higher harmonic control. The dynamic analysis of the rotor is performed using a comprehensive analysis tool coupled with a gradient based optimization algorithm to evaluate the higher harmonic cyclic input which is expected to minimize vibratory hub loads. The hub force minimization problem is converted into a cabin vibration reduction problem using a comprehensive aeroservoelastic helicopter model which includes a linear time invariant aeroelastic rotor model, rigid and elastic fuselage modes, servo actuators and sensors for vibration measurement. The optimal higher harmonic control inputs are then applied and the accelerations are measured at the selected sensor locations and additionally aeroservoelastic stability analysis is performed. For the same aeroservoelastic model, a cabin vibration reduction procedure is formulated considering direct acceleration reduction at sensor locations. All studies are conducted on a light utility helicopter. Vibratory hub loads and sensor accelerations are presented and the differences in the objectives of minimum force at hub and minimum vibration at sensors are discussed.
Detailed Design of an Active Rotor Blade for Reducing Helicopter Vibratory Loads
2011
An active trailing-edge flap blade named as Seoul National University Flap (SNUF) blade is designed for reducing helicopter vibratory loads and the relevant aeroacoustic noise. Unlike the conventional rotor control, which is restricted to 1/rev frequency, an active control device like the present trailing-edge flap is capable of actuating each individual blade at higher harmonic frequencies i.e., higher harmonic control (HHC) of rotor. The proposed blade is a small scale blade and rotates at higher RPM. The flap actuation components are located inside the blade and additional structures are included for reinforcement. Initially, the blade cross-section design is determined. The aerodynamic loads are predicted using a comprehensive rotorcraft analysis code. The structural integrity of the active blade is verified using a stress-strain recovery analysis.
Helicopter vibration reduction with trailing edge flaps
American Helicopter Society Northeast …, 1995
Main rotor blades with plain trailing edge aps are investigated as a potential means of vibration reduction using a comprehensive rotorcraft analysis. The analysis is modi ed to incorporate an unsteady aerodynamic model including ap e ects. Predicted blade and pitch link loads show adequate correlation with experimental data. An initial open loop study shows that, for a typical articulated four bladed rotor, signi cant reductions in xed system 4/rev vertical shears and hub moments are possible with open loop ap de ections at 3/rev and 4/rev. The reductions are most closely associated with changes in the response of the blade third atwise bending mode and torsional de ections. A reduction in blade torsional sti ness reduced the e ectiveness of the ap in reducing vibration. Notation C T Thrust coe cient C L Pro le lift coe cient C L qs Quasisteady value of C L M Trailing edge ap hinge moment coe cient M Y Blade sectional apwise bending moment, positive f o r u p wards blade tip de ections. L Vector of sectional aerodynamic loads, fP M M g T x Vector of local blade de ections, fv w g T R Rotor radius S Non-dimensional time, dS = ( V (t)=b)dt U T Local velocity parallel to blade undeformed axis, normalized to R U P Local velocity perpendicular to blade undeformed axis, normalized to R, positive is velocity from above.
Aeroelastic Modeling of Trailing-Edge-Flap Helicopter Rotors Including Actuator Dynamics
Journal of Aircraft, 2004
The effect of actuator dynamics on a helicopter rotor with trailing-edge flaps for vibration control is investigated. Trailing-edge flap, actuator, and elastic rotor blade equations of motion are formulated using Hamilton's variational principle. The coupled nonlinear, periodic equations are solved using finite elements in space and time. The baseline correlation study is based on wind-tunnel test data for a typical five-bladed bearingless rotor system. Good agreement is seen for the blade flap bending, chord bending, and torsion moments. It is shown that actuator dynamics cannot be neglected for a trailing-edge flap system with torsionally soft actuators. The parametric study performed using both coupled flap/actuator model and prescribed flap motion model indicated that the placement of trailing-edge flaps at 78% radius resulted in minimum flap input for this rotor. The vibration reduction level and trend are close between the predictions of both models at different forward speeds. Control inputs predicted by the coupled model show less sensitivity to the forward speed than that of prescribed model. Superscripts = d/dx = d/dψ = d 2 /dψ 2
Continuous Trailing-Edge Flaps for Primary Flight Control of a Helicopter Main Rotor
2014
The use of continuous trailing-edge flaps (CTEFs) for primary flight control of a helicopter main rotor is studied. A practical, optimized bimorph design with Macro-Fiber Composite actuators is developed for CTEF control, and a coupled structures and computational fluid dynamics methodology is used to study the fundamental behavior of an airfoil with CTEFs. These results are used within a comprehensive rotorcraft analysis model to study the control authority requirements of the CTEFs when utilized for primary flight control of a utility class helicopter. A study of the effect of blade root pitch index (RPI) on CTEF control authority is conducted, and the impact of structural and aerodynamic model complexity on the comprehensive analysis results is presented. The results show that primary flight control using CTEFs is promising; however, a more viable option may include the control of blade RPI, as well.
Contributions to the dynamics of helicopters with active rotor controls
This dissertation presents an aeromechanical closed loop stability and response analysis of a hingeless rotor helicopter with a Higher Harmonic Control (HHC) system for vibration reduction. The analysis includes the rigid body dynamics of the helicopter and blade flexibility. The gain matrix is assumed to be fixed and computed off-line. The discrete elements of the HHC control loop are rigorously modeled, including the presence of two different time scales in the loop. By also formulating the coupled rotor-fuselage dynamics in discrete form, the entire coupled helicopter-HHC system could be rigorously modeled as a discrete system. The effect of the periodicity of the equations of motion is rigorously taken into account by converting the system into an equivalent system with constant coefficients and identical stability properties using a time lifting technique. The most important conclusion of the present study is that the discrete elements in the HHC loop must be modeled in any HHC analysis. Not doing so is unconservative. For the helicopter configuration and HHC structure used in this study, an approximate continuous modeling of the HHC system indicates that the closed loop, coupled helicopter-HHC sys-tem remains stable for optimal feedback control configurations which the more rigorous discrete analysis shows can result in closed loop instabilities. The HHC gains must be reduced to account for the loss of gain margin brought about by the discrete elements. Other conclusions of the study are: (i) the HHC is effective in quickly reducing vibrations, at least at its design condition, although the time constants associated with the closed loop transient response indicate closed loop bandwidth to be 1 rad/sec on average, thus overlapping with FCS or pilot bandwidths, and raising the issue of potential interactions; (ii) a linearized model of helicopter dynamics is adequate for HHC design, as long as the periodicity of the system is correctly taken into account, i.e., periodicity is more important than nonlinearity, at least for the mathematical model used in this study; and (iii) when discrete and continuous systems are both stable, and quasisteady conditions can be guaranteed, the predicted HHC control harmonics are in good agreement. Quantitative assessment of the results needs to be tempered with the natural limitations of the nonlinear analytical helicopter model at hand to accurately predict vibration.
Trailing-Edge Flaps for Rotor Performance Enhancement and Vibration Reduction
Journal of the American Helicopter Society, 2013
The comprehensive analysis University of Maryland Advanced Rotor Code (UMARC) was used to quantify the capabilities of trailing-edge flaps (TEFs) for helicopter vibration reduction and performance improvement. The rotor performance in hover was improved with a combination of torsionally softer blades and positive TEF deflections. Suitable combinations of lower harmonic TEF inputs were shown capable of reducing the rotor power requirement by about 4-5% at an advance ratio of μ = 0.4. The TEF was shown to be capable of suppressing vibratory loads at a range of forward speeds, using half peak-to-peak deflections of about 5 • -10 • . Softening the blades in torsion resulted in larger flap actuation requirements for vibration reduction. A combination of 1, 2, 3, 4, and 5/rev TEF inputs resulted in a power reduction of 1.5%, while also reducing certain vibratory loads by more than 50% in high-speed forward flight.
Journal of the American Helicopter Society, 2008
A computational study of helicopter vibration and rotor shaft power reduction is conducted using actively-controlled 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 ACF approach for power as well as simultaneous 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 c c Flap chord C T Rotor thrust coefficient D Matrix defined to be T