Showing papers on "Helicopter rotor published in 1999"

Aeroelastic Analysis of the NASA/ARMY/MIT Active Twist Rotor

Abstract: Aeroelastic modeling procedures used in the design of a piezoelectric controllable twist helicopter rotor wind tunnel model are described. Two aeroelastic analysis methods developed for active twist rotor studies, and used in the design of the model blade, are described in this paper. The first procedure uses a simple flap-torsion dynamic representation of the active twist blade, and is intended for rapid and efficient control law and design optimization studies. The second technique employs a commercially available comprehensive rotor analysis package, and is used for more detailed analytical studies. Analytical predictions of hovering flight twist actuation frequency responses are presented for both techniques. Forward flight fixed system nP vibration suppression capabilities of the model active twist rotor system are also presented. Frequency responses predicted using both analytical procedures agree qualitativelty for all design cases considered, with best correlation for cases where uniform blade properties are assumed.

Feedback attenuation and adaptive cancellation of blade vortex interaction on a helicopter blade element

Abstract: Blade vortex interaction (BVI) noise has been recognized as the primary determinant of the helicopter's far field acoustic signature. Given the limitations of design in eliminating this dynamic phenomenon, there exists a need for control. We believe that this paper is the first model-based effort to attempt the same. We present the application, first of feedback control strategies, and then of adaptive cancellation on Hariharan and Leishman's linear aerodynamic model of a trailing edge flap. Lift fluctuations caused by vortices are taken as output disturbance. The contribution of the vortices to lift is obtained from Leishman's (1996) indicial model for gusts. The use of an active structure for actuation is assumed, and the actuator is approximated as a lag element. To design an adaptive cancellation scheme that is applicable not only to BVI but also to general problems with periodic disturbances, we start with the classical sensitivity method, and arrive at an adaptive scheme whose stability we discuss via averaging. Sacks et al. (1996) arrived at the same result by introducing a phase advance into a pseudogradient scheme.

Improved design and performance of the SMA torque tube for the DARPA Smart Wing program

Abstract: In recent Smart Wing wind tunnel tests at NASA Langley, we demonstrated over 5° of span-wise wing twist at M=O.205. This was a considerable improvement over the 1.25° of twist demonstrated during the initial tunnel test. Key to the improvements were two developments. First a different torque loading path in the structure, which resulted in torque beingdirectly reacted from root to wing tip. Secondly, a new SMA actuator was developed, with a measured blocking torque of3500 in-lb. The second round of tunnel tests not only demonstrated increased wing twist; we also were able tocommand a variety of twist angles and were able to show that the wing could maintain a predetermined twist for over anhour with a stability of 0.05°. Power consumption was recorded, with maximum power of 200W during twisting, and apower demand of2OW for maintaining wing twist. 1. iNTRODUCTION Shape memory alloy (SMA) torque tube development was focused on adaptive control of spanwise wing twist, specificallyto improve aircraft landing and take-off capabilities rather than maneuver performance. Today's inherent limitation ofbulk SMA material to respond quickly (a few Hz or higher), without highly aggressive active thermal management, limitsits use as a primary control surface actuator for aircraft applications. The real near term benefit of the technology, lies inlow rate wing twist optimization. Prior to this program, research and development work had focused on small torqueapplications for full scale designs (e.g. helicopter rotor tab). The program accomplishment -

Helicopter rotor blade with a movable flap

Abstract: A helicopter rotor blade includes a main airfoil body ( 12 ) and a movable control flap ( 3 ) incorporated in the trailing edge profile of the airfoil body. The flap ( 3 ) is movably connected to the main airfoil body ( 12 ) by a flexibly bendable junction element ( 4 ), and is actuated by a piezoelectric actuator unit ( 5 ) via a push/pull rod ( 6 ) and a lever arm ( 7 ), whereby the control flap is deflected relative to the main airfoil body. The junction element ( 4 ) is preferably a continuous integral fiber-reinforced composite component having a flexible bending portion ( 42 ) with a reduced thickness in comparison to the adjoining portions, whereby the reinforcing fibers extend continuously through the joint in the direction of the connection between the main airfoil body and the flap.

Steady-State Dynamic Behavior of a Flexible Rotor With Auxiliary Support From a Clearance Bearing

Abstract: This paper investigates the steady-state responses of a rotor system supported by auxiliary bearings in which there is a clearance between the rotor and the inner race of the bearing. A simulation model based upon the rotor of a production jet engine is developed and its steady-state behavior is explored over a wide range of operating conditions for various parametric configurations. Specifically, the influence of rotor imbalance, clearance, support stiffness and damping is studied. Bifurcation diagrams are used as a tool to examine the dynamic behavior of this system as a function of the aforementioned parameters. The harmonic balance method is also employed for synchronous response cases. The observed dynamical responses is discussed and some insights into the behavior of such systems are presented.

Optimum speed rotor

Abstract: A variable speed helicopter rotor system and method for operating such a system are provided which allow the helicopter rotor to be operated at an optimal angular velocity in revolutions per minute (RPM) minimizing the power required to turn the rotor and thereby resulting in helicopter performance efficiency improvements, reduction in noise, and improvements in rotor, helicopter transmission and engine life. The system and method provide for an increase in helicopter endurance and. The system and method also provide a substantial improvement in helicopter performance during take-off, hover and maneuver.

Interaction Dynamics Between a Flexible Rotor and an Auxiliary Clearance Bearing

Abstract: This study investigates the application of synchronous interaction dynamics methodology to the design of auxiliary bearing systems. The technique is applied to a flexible rotor system and comparisons are made between the behavior predicted by this analysis method and the observed simulation response characteristics. Of particular interest is the influence of coupled shaft/bearing vibration modes on rotordynamical behavior. Experimental studies are also perFormed to validate the simulation results and provide insight into the expected behavior of such a system.

Modeling of smart composite box beams with nonlinear induced strain

Abstract: A new smart composite box beam model is developed to investigate the behavior of helicopter rotor blades built around the active box beam. Piezoelectric actuators and sensors are surface bonded on the walls of the composite box beam. The new theory, based on a refined higher order displacement field of a plate with eccentricity, is a three-dimensional model which approximates the elasticity solution so that the box beam cross-sectional properties are not reduced to one-dimensional beam parameters. Both in-plane and out-of-plane warpings are included automatically in the formulation. The formulations also include nonlinear induced strain effects of piezoelectric actuators. The procedure is implemented using finite element method. The developed theory is used to model the load carrying member of helicopter rotor blades with moderately thick-walled sections. Static analysis of the smart box beam under varying degrees of actuation has been performed. Very good overall agreement is observed with available experimental data for thin-walled sections without embedded actuators. The results show that piezoelectric actuation significantly reduces the deflection along the box beam span and therefore can be used to control the magnitude of rotor blade vibrations. The nonlinear actuation effect is found to be closely related to the material stiffness of the primary structure.

Experimental and Theoretical Correlation of Helicopter Rotor Blade-Droop Stop Impacts

Abstract: The transient response of a nonrotating articulated rotor blade undergoing a droop stop impact is examined. The rotor blade is modeled using the finite element method, and the droop stop is simulated using a conditional rotational spring. No aerodynamic effects are modeled. Three methods of time integrating the equations of motion were studied: 1) a direct integration of the full finite element space equations of motion; 2) a modal space integration using only hinged modes; and 3) a modal space integration using either hinged or cantilevered modes, depending on blade/droop stop contact. Given a range of initial flap hinge angles, drop tests of a one-eighth Fronde-scaled articulated model rotor blade were conducted at zero rotational speed. The transient tip deflection, flap hinge angle, and strain were measured, and they displayed good correlation with all three analytic methods. Modal parameter identification tests were performed on the model blade to determine its natural frequencies and damping ratios for both hinged and cantilevered conditions. The measured structural damping was shown to significantly improve correlation between the experimental and analytic results. Computational efficiency for the problem under consideration was not of serious concern. However, in a comprehensive aeroelastic analysis, it was found that a modal space integration using either hinged or cantilevered modes, depending on blade/droop stop contact, reduced computational time by two orders of magnitude.

Algorithm to Integrate the Ffowcs Williams -Hawkings Equation on Supersonic Rotating Domain

Abstract: The numerical solution of the Ffowcs Williams -Hawkings equation (Ffowcs Williams, J. E., and Hawkings, D. L., " Sound Generation by Turbulence and Surfaces in Arbitrary Motion," Philosophical Transactions of the Royal Society , Vol. A264, No. 1151, 1969, pp. 321 -342) on a rotating supersonic domain is discussed. Based on the emission-surface algorithm, the adopted solver performs the integration on the so-called acoustic domain to avoid the Doppler singularity in the integral kernels. The presence of multiple emission times for the supersonic source points and the particular time evolution of the integration domain force the use of a particular data-e tting procedure on both the geometrical and integral quantities. The algorithm may be used in thenumerical prediction of the quadrupole source term for helicopter rotors operating at a high transonic regime and in the aeroacoustic analysis of the modern propeller blades, rotating at supersonic tip speed. INCE the end of the 1970s, when the importance of nonlin- ear terms was highlighted in the numerical estimation of noise from a high tip speed rotating blade, a change in the development and the application of theoretical and computational methodologies for the aeroacoustic analysis of helicopter rotors has occurred. Be- cause of the requirement for a three-dimensional integration and the presence of the Doppler singularity in the integral kernels, the adoption of the Ffowcs Williams -Hawkings(FW-H) 1 equation for the numerical prediction of the high-speed impulsive (HSI) noise hasalwaysbeenconsidereddife cultandcomputationallyexpensive. Over the past 15 years, interest of the aeroacoustic community has progressively moved toward alternative solution forms, such as the Kirchhoffapproach 2 andthecomputationalaeroacousticsmethods, 3 relegating the acousticanalogy approach to the role of alinear prob- lemsolver.Actually,thecomputationofthenonlineartermsisavery dife cult task. The delocalization of the shock waves off the blade tip arising at Mtip ¸ 0:88 forces the extention of the computing domain beyond the sonic cylinder, where the Doppler factor pre- vents the usual FW -H solvers from achieving a reliable prediction of noise. From a theoretical point of view, the problem may be sim- ply bypassed through the use of the emission-surface algorithms, where the integrals are determined on the blade retarded cone gu- ration, and the Doppler singularity does not appear in the integral kernels.Recently,anewmethodhasbeenproposedandsuccessfully implemented, 4 where the calculations proceed forward with respect to time, thus avoiding the solution of the retarded-time equation and the effects of the Doppler singularity. The emission-surface ap- proach rarely has been used: The occurrence of multiple emission timesinthesupersonicregioncausesunconnectedpatchestoappear, which temporarily link together into a single domain, following a time evolution, which is very dife cult to numerically model. Very interesting results concerning the integration on the acous- tic surface have been published by Wells 5;6 and Wells and Han. 7 In these papers the need for determining, at each time step, a new computational grid with a clustering of the integration points along some particular critical radii was recognized, and very smooth sig- natureswereobtainedinthedeterminationoftheFW -Hlinearterms from a rectangular rotor blade and a propfan-type blade rotating at supersonic tip speed. Wells provides a detailed discussion of the mathematical aspects of the problem and points out the important

A Theoretical Study of the Effect of Pinion Offset on the Dynamics of Hypoid Geared Rotor System

Abstract: A new 3-dimensional coupled rotational-translational vibratory model of a high-speed loaded hypoid geared rotor system has been formulated to analyze the dynamic effect of pinion offset. This model includes the effective shaft and bearing flexibilities, and gear mesh induced dynamic couplings of the lateral, axial, torsional and rotational motions. The proposed formulation is also capable of simulating drive and coast operating cases. Its effective mesh point and line of action are assumed stationary for a given steady-state condition, and they are defined by the theoretical pitch point vector and corresponding normal vector at the point of contact respectively. The proposed analytical model is applied here to compute the modal response functions of a typical automotive drivetrain design for a selected range of pinion offset. The calculations revealed interesting frequency-dependent effects of pinion offset on the generation of dynamic mesh force and bearing reaction loads.

High torque actuation system for an active rotor control system

Abstract: An actuation system for pivoting a flap on a helicopter rotor blade to reduce the interaction of the blade with the preceding blade vortex. The actuation system includes a fluid supply which is connected to first and second fluid supply lines. The fluid supply lines convey flows of pressurized fluid from the fluid supply to an actuator. The actuator includes a housing mounted within the rotor blade and having a channel formed in it. A butterfly shaft is pivotally mounted within the channel and has laterally extending arms which separate the channel into four lobes. A first port connects the first fluid supply line with two diametrically opposed lobes in the channel. A second port connects the second fluid supply line with the other two diametrically opposed lobes in the channel. A torque coupling is attached to the butterfly shaft and engaged with the flap such that rotation of the torque coupling produces concomitant rotation of the flap. The pressurization of the first fluid supply line causes the torque coupling to rotate in a first direction. The pressurization of the second supply fluid line causes the torque coupling to rotate in the opposite direction.

Helicopter with control by displacement of its center gravity

Abstract: In the case of a helicopter comprising a rotor system ( 3 ) with rotor blades ( 6 and 6 ′) held on a fuselage ( 2 ) and at least one rotor shaft ( 5 and 5 ′) and furthermore comprising a drive system ( 4 ) for causing rotation of the rotor system ( 3 ), the rotor system ( 3 ) together with the drive system ( 4 ) is able to be slid in the longitudinal direction of the fuselage ( 2 ) and to be pivoted around a pivot axis extending along the fuselage ( 2 ).

Development and validation of a refined piezostack-actuated trailing-edge flap actuator for a helicopter rotor

Abstract: A methodology to develop piezostack-based actuator for trailing edge flap on a full-scale helicopter rotor blade is presented. By a systematic down-selection approach through the experimental validation on piezostack characteristics, one among numerous candidates was chosen for current study, which has higher stiffness as well as good strain/block force capability. The prototype actuator was designed and fabricated using two pieces of selected piezostack incorporated with a double-lever amplification mechanism that extends a level-fulcrum amplification concept. To validate the feasibility of the designed actuator, the prototype actuator was tested in vacuum chamber that simulates the centrifugal force environment. It was shown that the actuator operated with no major degradation in performance up to 600g of centrifugal loading. The present actuator has a capability of oscillating the trailing edge flap to the required deflection/frequency to actively control rotor vibration in the rotating environment. Then, a further design refinement was considered to reduce the drawbacks, which finally results in an improved actuator with peak actuation force of 80 lbs at 38 mils stroke.© (1999) COPYRIGHT SPIE--The International Society for Optical Engineering. Downloading of the abstract is permitted for personal use only.

Adaptive Neurocontrol of Simulated Rotor Vibrations Using Trailing Edge Flaps

Abstract: Smart structure activated trailing edge flaps are capable of actively altering the aerodynamic loads on rotor blades. Coupled with a suitable feedback control law, such actuators could potentially be used to counter the vibrations induced by periodic aerodynamic loading on the blades, without the bandwidth constraints and with a potential of lower weight penalties incurred by servo actuation methods. This paper explores new, robust individual blade control (IBC) methodologies for vibration suppression using a piezoactuated trailing edge flap. The controllers employ a single hidden layer neural network, learning in real time, to adaptively cancel the effects of periodic aerodynamic loads on the blades, greatly attenuating the resulting vibrations. Both collocated and noncollocated sensorlactuator pairs are considered. Proofs of the stability and convergence of the proposed neurocontrol strategies are provided, and numerical simulation results for a one-eighth Froude scale blade model are given which demonstrate that the controller can nearly eliminate the blade vibration arising from a wide variety of unknown, periodic disturbance sources.

Improvement of aeroelastic stability of hingeless helicopter rotor blade by passive piezoelectric damping

Abstract: To augment weakly damped lag mode stability of a hingeless helicopter rotor blade in hover, piezoelectric shunt with a resistor and an inductor circuits for passive damping has been studied. A shunted piezoceramics bonded to a flexure of rotor blade converts mechanical strain energy to electrical charge energy which is dissipated through the resistor in the R-L series shunt circuit. Because the fundamental lag mode frequency of a soft-in-plane hingeless helicopter rotor blade is generally about 0.7/rev, the design frequency of the blade system with flexure sets to be so. Experimentally, the measured lag mode frequency is 0.7227/rev under the short circuit condition. Therefore the suppression mode of this passive damping vibration absorber is adjusted to 0.7227/rev. As a result of damping enhancement using passive control, the passive damper which consists of a piezoelectric material and shunt circuits has a stabilizing effect on inherently weakly damped lag mode of the rotor blades, at the optimum tuning and resistor condition.

Helicopter rotor blade weighting

Abstract: Rotor blade weighting apparatus comprising in combination at least one elongated rod carried within a rotor blade, and multiple weights received and retained on the rod, the number of weights adjustable for reducing vibration during rotor rotation.

Cosite interference between wire antennas on helicopter structures and rotor modulation effects: FDTD versus measurements

Abstract: As telecommunication systems become more complex and more antennas are placed on the same structure (e.g., helicopter airframe) the problem of interference becomes significant for the performance of the systems. The finite-difference time-domain (FDTD) method is used to calculate coupling between wire elements; e.g., monopoles and loops mounted on ground planes and helicopter airframes. Also, rotor modulation effects on coupling are investigated. All the numerical results obtained by FDTD are validated by comparison with measurements.

Preliminary testing of a mach-scaled active rotor blade with a trailing edge servo-flap

Abstract: The results of preliminary tests on an active helicopter rotor blade are presented. The blade, a Mach-scaled model of a CH-47D helicopter blade, has a discrete piezoelectric actuator embedded within the spar that controlsa trailing edge flap via a pushrod. Ultimately, the blade will be tested on a helicopter rotor hover stand at MIT. In this paper, we describe the tests performed prior to hover testing. First, the actuator was tested on the bench to determine its control authority and frequency response. Second, the actuator was tested on a shake table to simulate the out-of-plane accelerations that would be encountered in a full-scale helicopter in forward flight. Third, the actuator was embedded in the model blade, and its response to low-frequency sinusoidal actuation was obtained and compared to the bench test results. Finally, the frequency response of the actuator in the blade was determined using swept sine excitation. All test results indicate that the actuator should produce the desired level of control authority in the model-scale rotor.

Parallel implementation of an upwind Euler solver for hovering rotor flows

Abstract: Abstract An Euler method for computing compressible hovering rotor flows is described. The equations are solved using an upwind finite-volume method in a blade-fixed rotating co-ordinate system, so that hover is a steady problem. Transfinite interpolation, along with a periodic transformation, is used to generate grids for the periodic domain. Computation of these flows to an acceptable accuracy requires fine grids, and a long integration time for the wake to develop, resulting in excessive run times on a single processor. Hence, the method is developed as a multiblock code in a parallel environment, and various aspects of data passing and communication between processors have been considered. It is shown that a considerable increase in performance is available from the use of non-blocking and asynchronous communication. It is also demonstrated that increased performance may be available by balancing the residual levels rather than the number of cells on each processor.

Aeroelastic response of a coupled rotor/fuselage system in hovering and forward flight

Abstract: The aeroelastic response analysis of a coupled rotor/fuselage system is approached by iterative solution of the blade aeroelastic response in the non-inertial reference frame fixed on the hub, and the periodic response of the fuselage in the inertial reference frame. A model of the coupled system hinged with the flap and lag hinges, the pitching bearing which may not coincide with the hinges, and the sweeping-blade configuration is established. The moderate-deflection beam theory and the two-dimensional quasi-steady aerodynamic model are employed to model the aeroelastic blade, all the kinetic and inertial factors are taken into account in a unified manner. A five-nodes, 15-DOFs pre-twisted nonuniform beam element is developed for the discretization of the blade, three rigid-body-motion DOFs are introduced for the motion of the hinges and the bearing. The Hamilton's principle is employed to evaluate the equation of motion of the blade. The derived nonlinear ordinary differential equations with time-dependent periodic coefficients are solved by a modified quasi-linearization method, which is developed for the higher DOF periodic system. The resulting periodic forces and moments exerted on the fuselage by all the blades are evaluated every time, when the converged nonlinear periodic response of the blade is obtained under the consideration of the equilibrium of the blades. The fuselage structure is simplified to be a beam structure, the governing equation is established in the inertial reference frame and a two-nodes beam element is used to discretize the flexible fuselage. The periodic response of the fuselage is solved by a simple shooting method. The iteration of the rotor/fuselage response is continued, until the aeroelastic responses of the blade and the fuselage converge simultaneously. Both the hovering and the forward flight states can be considered. The results of a computed numerical example by the developed program are presented to verify in practice the economy of the modeling as well as the reliability and efficiency of the corresponding solving methods.

Helicopter blade positioning mechanism for foldable rotor blades

Abstract: An improved blade positioning mechanism for folding a helicopter blade attached to a pitch control housing permits a controlled folding of the main rotor blade for increased safety and decreased risk of damage to the helicopter or other equipment. The invention also folds the main rotor blade while the blade remains attached to the pitch control housing, thus eliminating the need to rebalance the blade. In an embodiment of the invention blade positioning mechanism comprises a pitch control housing connection, a rotor blade clamp, and a clamp positioner. The pitch control housing connection is adapted to temporarily attach to the pitch control housing of the helicopter and pivot with two degrees of freedom relative to the pitch control housing. The rotor blade clamp is adapted to temporarily attach to the rotor blade of the helicopter. The clamp positioner is attached to the rotor blade clamp and the pitch control housing connection. The positioner is adapted to pivot with two degrees of freedom relative to the rotor blade clamp. Further, once all but one of the blade retention pins that attach the blade to the pitch control housing has been removed, the positioner is adapted to position the rotor blade clamp a distance from the pitch control housing connection, whereby the rotor blade pivots about the remaining pin.

The phenomenon of helicopter rotor blade sailing

Abstract: Blade sailing is an aeroelastic phenomenon affecting helicopter rotors when rotating at low speeds in high wind conditions. This is a potentially dangerous blade motion and the excessive flapwise tip deflections generated endanger the airframe, the flight crew and any personnel working close to the aircraft. This phenomenon is particularly applicable to naval helicopters or those operating off exposed sights such as oil rigs. The research covered an experimental investigation into the effect of an abeam wind flow over a simulated flight deck of a Rover Class Royal Fleet Auxiliary vessel. Blade flexibility and rotor hub mechanical features were introduced into the theory and the resulting method applied to the Westland Lynx and Sea King aircraft. The semi-rigid rotor of the Lynx is relatively well controlled, but the rotor hub construction of the articulated rotor of the Sea King and the interaction with the flexing blades allow blade tip deflections to be generated of an order to strike the fusela...

Simulation and identification of helicopter rotor dynamics using a general-purpose multibody code

Abstract: With the purpose of evaluating different model identification strategies and of assisting in the validation of active control designs, a simulation program for the dynamics of the main rotor of the AGUSTA A109c helicopter has been developed (based on a general-purpose multibody code). The program has been applied to the testing of various system identification algorithms, both for time-invariant and time-periodic models of the response of the rotor to perturbations in the control inputs, in simulated experiments.

Thickness Noise of Helicopter Rotors at High Tip Speeds

Abstract: A new formulation of helicopter rotor thickness, noise for hover and forward flight, is discussed. The parameters required for this formulation are rotor motion, planform and airfoil thickness distribution. A computer program has been developed to calculate the pressure signature due to blade thickness for a helicopter in arbitrary motion. Comparison with high-speed helicopter tests shows good agreement with calculations when the observer is in or near the horizontal plane in which the rotor disc lies. Characteristics of thickness noise are illustrated by numerical examples indicating strongly that the high-speed blade slap may be due primarily to the thickness effect. The methods of Deming and Arnoldi are discussed as the special cases of this technique.

Development of a piezoelectric actuator for trailing edge flap control of rotor blades

Abstract: Piezoelectric actuator technology has now reached a level where macro-positioning applications in the context of smart structures can be considered. One application with high payoffs is vibration reduction, noise reduction, and performance improvements in helicopters. Integration of piezoelectric actuators in the rotor blade is attractive, since it attacks the problem at the source. The present paper covers the development of a piezoelectric actuator for trailing edge flap control on a 34-foot diameter helicopter main rotor. The design of an actuator using bi-axial stack columns, and its bench, shake, and spin testing are described. A series of enhancements lead to an improved version that, together with use of latest stack technology, meets the requirements. Next steps in this DARPA sponsored program are development of the actuator and full scale rotor system for wind tunnel testing in the NASA Ames 40x80 foot wind tunnel and flight testing on the MD Explorer.