Showing papers on "Helicopter rotor published in 1976"

Paired blade assembly

Abstract: A rotor blade system which is adaptd for long term reliable operation in a gas turbine engine is disclosed. Techniques incorporating composite materials into the rotor system are developed. One rotor structure shown utilizes a paired blade assembly having a core of continuous fibers running from the tip of one blade to the tip of the adjacent blade. Each of said paired blade assemblies is mchanically detachable from the engine rotor.

Numerical Calculation of Unsteady Transonic Potential Flow over Helicopter Rotor Blades

Abstract: The small disturbance potential equation appropriate to a helicopter in forward flight is derived. This equation then is solved for the flow over a nonlifting transonic rotor blade, using a completely implicit scheme that is an extension of the Murman-Cole mixed difference technique. The flow in the tip region is most unsteady in the decelerating flow region, after the blade passes the \J/ - 90°azimuthal station. The unsteadiness appears to be caused by expansion and compression waves that move slowly upstream of the blade as the relative incident flow decelerates. The influence of aspect ratio, advance ratio, and Mach number on this process is discussed.

Advanced control system for a rotor and/or a compound or rotary wing vehicle

Abstract: The control system includes a sensing mechanism by which a true pure signal from flapping bending activity resulting from an external load force on a rotor blade in a rotor system is generated and fed back to a gyroscope which then precesses to return a correcting feathering motion to the rotor blade through a swashplate. Such generation is developed by the sensing mechanism which comprises a cantilevered beam or spring system secured at its one end to a fixed hub arm about which the blade feathers and having its other end operatively connected to the gyroscope. The gyroscope is positioned outside of a high force blade feathering loop and is independently sized from the rotor to which it is coupled. The system also provides for pilot command input to produce a control feathering motion to a pair of blades independently of the sensing means.

Helicopter rotor and transmission mounting and vibration isolation system

Abstract: A helicopter rotor and transmission mounting and vibration isolation system in which the transmission is supported from the fuselage by a plurality of elastomeric mounts which are selectively positioned and focused to establish a system roll axis and a system pitch axis on opposite sides of the elastomeric mounts and which mount members are of selected stiffness to establish selected roll stiffness and selected pitch stiffness of the system, to establish the natural frequencies of the system sufficiently below the blade passage frequency to minimize the response of the fuselage to forces and moments imparted by the helicopter rotor and to provide selected torque restraint and lift restraint for the system.

Aeroelastic analysis for helicopter rotor blades with time-variable, non-linear structural twist and multiple structural redundancy: Mathematical derivation and program user's manual

Abstract: The differential equations of motion for the lateral and torsional deformations of a nonlinearly twisted rotor blade in steady flight conditions together with those additional aeroelastic features germane to composite bearingless rotors are derived. The differential equations are formulated in terms of uncoupled (zero pitch and twist) vibratory modes with exact coupling effects due to finite, time variable blade pitch and, to second order, twist. Also presented are derivations of the fully coupled inertia and aerodynamic load distributions, automatic pitch change coupling effects, structural redundancy characteristics of the composite bearingless rotor flexbeam - torque tube system in bending and torsion, and a description of the linearized equations appropriate for eigensolution analyses. Three appendixes are included presenting material appropriate to the digital computer program implementation of the analysis, program G400.

Noise produced by turbulent flow into a propeller or helicopter rotor

Abstract: than can be applied to calculate the instantaneous sound spectrum produced by the rotor at each azimuthal rotor position, and this instantaneous spectrum can be averaged over the azimuthal rotor position to find an averaged far-field sound spectrum. In taking this average, account must be taken of the different amount of retarded time that the rotor spends at each azimuthal rotor position. Further discussion of this point is given in another paper.4 A further factor taken into account in the analysis is the existence of blade-to-blad e correlation. If a given turbulent eddy is chopped by more than one rotor blade, the blade-toblade correlation leads to narrow-band noise peaked around the rotor harmonics. The far-field sound for an airfoil moving in rectilinear motion through a turbulent flow can be expressed in terms of a single wavevector component of the turbulence. The presence of blade-to-blade correlations requires that the single wavevector component be replaced by a summation over several wavevector components. This summation generally is carried out numerically for the calculations presented herein, but, if the frequency of interest is high enough, the summation can be replaced by an integral that can be evaluated in closed form; i.e., the blade-to-blade correlation becomes unimportant, and the result reduces to that for a single blade in rectilinear motion. The preceding description of the procedure for calculating the far-field sound applies to the sound produced by a spanwise segment of the rotor. This segment must have a spanwise dimension small enough so that the velocity does not vary significantly over the segment but large enough so that the loading correlation from segment to segment is not significant. This latter assumption is consistent with the highfrequency assumption mentioned previously, since high frequency corresponds to small correlation length. Thus, to find the noise contributed by the entire rotor, an integral over span must be performed.

Cross beam rotor

Abstract: A helicopter rotor having flexible blades mounted to a drive shaft by means of hub arms, wherein opposing blade members are interconnected by a common spar passing across the rotor axis. The spar members are supported from the hub arms by spherical bearing members. The universal freedom of these bearing members provides torsional freedom for blade pitching motions without restricting blade flapping or in-plane bending.

Helicopter rotor lead-lag damper

Abstract: A helicopter rotor lead-lag hydraulic fluid damper has an internal anti-torque shaft preventing rotary motion of the dynamic fluid retaining seal. An independent fluid reservoir under centrifugal force supplies lubrication to the low pressure side of the seal. At rest, the fluid returns to the lower reservoir chamber by gravity.

A Method for Predicting Helicopter Hub Drag

Abstract: : A procedure has been developed to determine the contribution of the rotor hub to the total helicopter drag. The method developed uses a three- dimensional potential flow analysis to determine the flow environment in which the hub operates combined with empirical data in order to predict the drag of the hub and its associated interference drag. Predictions using the method are in good agreement with test data for unfaired and faired rotor hubs. A review of available rotor hub drag test data was conducted in order to identify the factors affecting helicopter rotor hub drag. The data base established was used in the development of the hub drag prediction method and also to define a systematic wind tunnel test program to refine and verify the drag prediction method and to investigate in detail the parameters affecting the drag contribution of the rotor hub.

Adjusting the static moment of a rotor blade

Abstract: In finishing the construction of a rotary wing, particularly helicopter rotor blade, the blade is weighed to determine the position of its static moment. The measured value is compared with a predetermined value and, if necessary, the static moment is adjusted by applying a high specific gravity filler material to the surface of the blade. The filler material has a specific gravity in a range of 6 to 10 p/cm 3 and can be formed of a mixture of tungsten powder and epoxy resin. The weighing operation is performed on a three scale apparatus with one scale supporting the end of the blade attached to the rotor and the other two scales supporting the opposite end of the blade.

Integrated helicopter flight control

Abstract: A helicopter rotor control mechanism is provided in which cyclic and collive rotor control functions are performed by a single integrated control having three degrees of freedom and operable by either or both hands of the pilot as desired. The integrated control comprises a floor mounted control column rotatable about a fore and aft collective pitch control axis and a pair of hand grips mounted for rotation about two independent cyclic control axes (by one or both hands of the operator) in a control head assembly mounted on the upper end of the floor mounted control column. The floor column and control head effect displacement of collective pitch, cyclic pitch and cyclic roll output rods for ultimate control of the helicopter rotor.

Model helicopter throttle governor/collective pitch control apparatus

Abstract: A throttle governor/collective pitch control apparatus for radio controlled model helicopters for proportionally controlling either the model helicopter rotor speed or the rotor collective pitch, including sensing and timing means determining the rotor speed and any changes thereof, comparison means for comparing a subsequently sensed rotor speed with a first sensed rotor speed and developing an error signal, and control means responsive to the error signal to provide a control signal proportional to any changes in the rotor speed. In the throttle governor mode any variations in rotor speed result in a proportional control signal to vary the model helicopter throttle so as to maintain constant rotor speed. In the auto-collective mode changes in rotor speed result in a proportional control signal coupled to the collective pitch servo for proportionally varying the model helicopter collective pitch.

A Note on Blade Loss Dynamics of Rotor-Bearing Systems

Abstract: The equations for a single-mass rotor are examined for maximum response following an impact load created by a blade loss. The close examination of the undamped system gives insight and understanding to the dynamic response of the rotor system during an actual blade loss. The time transient response orbits for the single-mass system are presented for operation below, coincident with and above the given natural frequency. The orbits clearly indicate the ratio of maximum response relative to steady-state (i.e., overshoot ratio). The results of the influence of various damping ratios on the transient response peak amplitude are presented in dimensionless form. The blade loss transient response of a three bearing, eighteen-degree-of-freedom rotor system is presented to illustrate the validity of the basic concepts developed in the single-mass rotor analysis.

Helicopter rotor blade

Abstract: Torsional oscillations of helicopter rotor blades are reduced with stall, otherwise leading to such torsional oscillations, precluded consistent with allowable high angle of attack, by sucking air from the boundary layer on the upper surface of the blade near the leading edge thereof at a radial location of 80-90% maximum radius thereof and pumping the sucked boundary layer air, utilizing the inherent centrifugal pumping capability of the blade by establishing an exhaust of the centrifugal pumping passage within the blade which allows adequate pumping for the torsional oscillation reduction purpose consistent with the required rate flow of air.

Investigation of a bearingless helicopter rotor concept having a composite primary structure

Abstract: Experimental and analytical investigations were conducted to evaluate a bearingless helicopter rotor concept (CBR) made possible through the use of the specialized nonisotropic properties of composite materials. The investigation was focused on four principal areas which were expected to answer important questions regarding the feasibility of this concept. First, an examination of material properties was made to establish moduli, ultimate strength, and fatigue characteristics of unidirectional graphite/epoxy, the composite material selected for this application. The results confirmed the high bending modulus and strengths and low shear modulus expected of this material, and demonstrated fatigue properties in torsion which make this material ideally suited for the CBR application. Second, a dynamically scaled model was fabricated and tested in the low speed wind tunnel to explore the aeroelastic characteristics of the CBR and to explore various concepts relative to the method of blade pitch control. Two basic control configurations were tested, one in which pitch flap coupling could occur and another which eliminated all coupling. It was found that both systems could be operated successfully at simulated speeds of 180 knots; however, the configuration with coupling present revealed a potential for undesirable aeroelastic response. The uncoupled configuration behaved generally as a conventional hingeless rotor and was stable for all conditions tested.

Theoretical Study of Multicyclic Control of a Controllable Twist Rotor

Abstract: Analytical studies were performed to ascertain the feasibility of reducing helicopter rotor induced 4/rev vibratory forces by means of multicyclic flap control input on a dual control, four bladed rotor system. The dual control consisted of a primary inboard pitch horn blade control and a secondary outboard flap control. Flap control was put in at frequencies greater than the rotor rotational speed.

Aeroelastic Stability Of Complete Rotors With Application To A Teetering Rotor In Forward Flight

Abstract: The derivation of a set of nonlinear coupled flap-lag-torsion equations of motion for moderately large deflections of an elastic, two bladed teetering helicopter rotor in forward flight is concisely outlined. The following degrees of freedom are included in the mathematical model: rigid body flapping, rigid body lead lag, elastic bending in flap and lead-lag blade root torsion and shaft torsion. Quasi-steady aerodynamic loads are considered and the effects of reversed flow are included. The aeroelastic stability of the complete rotor is investigated using a linearized system of equations of motion. The equilibrium position about which the equations are linearized is obtained by considering the trim state of the helicopter, in true or simulated forward flight conditions. The sensitivity of the aeroelastic stability boundaries to interblade structural and mechanical coupling is illustrated by comparing the complete rotor stability boundaries with those obtained from a single blade analysis for a number of hover and forward flight cases.

Flight-Test Experience Of A Helicopter Encountering An Airplane Trailing Vortex

Abstract: This paper presents results of a flight-test experiment of a UH-1H helicopter encountering the vortex wake of a C-54 airplane. The helicopter was instrumented to record the pilot control inputs, determine the upset experience, and measure critical loads within the rotor system. During the flight-test program 132 penetrations of the vortex wake were made by the helicopter at separation distances from 3/8 to 6-1/2 nautical miles. Test results indicated that the helicopter upsets and the vortex induced blade loads experienced were minimal and well within safe limits. The upsets were very mild when compared to a typical response of a small airplane to the vortex wake of the C-54 airplane.

Helicopter rotor blades

Abstract: A helicopter rotor blade is constructed at least in part of composite material of the type in which fibres (such as glass or carbon fibres) are embedded in a matrix. The stiffness in the flapping plane, the lagging plane, or both are modified at one or more positions along the blade to provide advantageous resonance frequencies. Modification of the stiffness is achieved by diverting a substantial number of fibres inwardly towards the relevant spanwise neutral bending plane.

Torsionally compliant helicopter rotor blade with improved stability and performance characteristics

Abstract: A torsionally compliant helicopter rotor blade with maximum blade torsional flexibility at the blade tip portion and so shaped at the blade root portion that the locus of shear centers and the locus of centers of lift are substantially coincident, and so shaped at the tip portion so that the locus of shear centers is selectively forward of the locus of centers of lift to establish an aerodynamic restoring moment in response to blade tip torsional excursions, thereby improving blade torsional stability.

A critical examination of the flap-lag dynamics of helicopter rotor blades in hover and in forward flight

Abstract: A critical examination of flap-lag stability of a centrally hinged, spring-restrained rigid blade in both hover and forward flight is presented. Several differences in the equations of motion for blade flap-lag stability in the existing literature are identified. A rigorous and systematic development of these equations for a rigid articulated blade in forward flight shows the existence of some linear aerodynamic coupling terms associated with blade steady-state flapping and lagging in the perturbation equations. The differences identified are shown to be associated with the order in which the flap and lag transformations are taken in developing the equations of motion. The implications of these differences on stability are examined, and it is shown that the pitch-lag coupling terms associated with a flap-lag hinge transformation sequence have a marked influence on flap-lag stability. Some qualitative considerations on the role of the assumed transformation sequence in the development of the flap-lag equations for a hingeless elastic blade are also given. On the basis of these considerations, it is shown that aerodynamic coupling terms associated with blade steady-state flapping and lagging similar to those found for the rigid blade will also appear in the equations for the elastic blade.

Correlation of full-scale helicopter rotor performance in air with model-scale Freon data

Abstract: An investigation was conducted in a transonic dynamics tunnel to measure the performance of a 1/5 scale model helicopter rotor in a Freon atmosphere. Comparisons were made between these data and full scale data obtained in air. Both the model and full scale tests were conducted at advance ratios between 0.30 and 0.40 and advancing tip Mach numbers between 0.79 and 0.95. Results show that correlation of model scale rotor performance data obtained in Freon with full scale rotor performance data in air is good with regard to data trends. Mach number effects were found to be essentially the same for the model rotor performance data obtained in Freon and the full scale rotor performance data obtained in air. It was determined that Reynolds number effects may be of the same magnitude or smaller than rotor solidity effects or blade elastic modeling in rotor aerodynamic performance testing.

Soil stabilizer machine with detachable hydraulic motor structure driving rotatably mounted cutting elements

Abstract: A self-propelled soil stabilizer machine employs a heavy-duty single horizontal rotor for pulverizing and mixing soil it passes over. The rotor is driven by hydraulic motors which are mounted at the rotor ends and operated by a hydraulic pump (engine driven) which is hydraulically coupled to a hydraulic traction pump (also engine driven) which propels the machine. Each hydraulic motor is mounted for rapid removal and replacement thereof. Toward this goal, each end of the rotor is hollow, whereby the drive shaft of each hydraulic motor can be releasably accepted in one of the ends of the rotor in a spline connection therebetween. Overloads on the rotor system are automatically sensed by a closed-loop hydrostatic system and result in a direction in the speed of machine travel until the load on the rotor diminishes. A hydraulic rotary servo-valve and mechanical feed back system automatically controls the rotor to maintain it at a preset depth and provides a visual read-out of depth.

Thermoelectric device for helicopter rotor blades - has heating element net works along leading edges and protective skin

Abstract: The helicopter blade de-icer consists of a heating network which runs the length and both sides of the leading edge of each blade. An external anti-abrasive and anti-corrosive skin consists stainless steel strips covering the leading edge, their adjoining edges forming transverse joints. The heating network is divided into several elements electrically connected to one another by bridges crossing joints only in the vicinity of the leading edge. The bridges connecting the elements are made of slightly undulating metal cables.

The effect of varying freestream velocity on dynamic stall characteristics

Abstract: A low speed wind tunnel equipped with an axial gust generator to simulate the aerodynamic environment of a helicopter rotor was used to study the dynamic stall of a pitching blade. The objective of this investigation was to find out to what extent harmonic velocity perturbations in the freestream affect dynamic stall. The study involved making measurements of the aerodynamic moment on a two-dimensional, pitching blade model in both constant and pulsating airstreams. Using an operational analog computer to perform on-line data reduction, plots of moment versus angle of attack and work done by the moment were obtained. The data taken in the varying freestream were then compared to constant freestream data, and to the results of two analytical methods. These comparisons showed that the velocity perturbations had a significant effect on the pitching moment which could not be consistently predicted by the analytical methods, but had no drastic effect on the blade stability.