Showing papers on "Helicopter rotor published in 1978"

The importance of quadrupole sources in prediction of transonic tip speed propeller noise

Abstract: A theoretical analysis is presented for the harmonic noise of high speed, open rotors. Far field acoustic radiation equations based on the Ffowcs-Williams/Hawkings theory are derived for a static rotor with thin blades and zero lift. Near the plane of rotation, the dominant sources are the volume displacement and the rho U(2) quadrupole, where u is the disturbance velocity component in the direction blade motion. These sources are compared in both the time domain and the frequency domain using two dimensional airfoil theories valid in the subsonic, transonic, and supersonic speed ranges. For nonlifting parabolic arc blades, the two sources are equally important at speeds between the section critical Mach number and a Mach number of one. However, for moderately subsonic or fully supersonic flow over thin blade sections, the quadrupole term is negligible. It is concluded for thin blades that significant quadrupole noise radiation is strictly a transonic phenomenon and that it can be suppressed with blade sweep. Noise calculations are presented for two rotors, one simulating a helicopter main rotor and the other a model propeller. For the latter, agreement with test data was substantially improved by including the quadrupole source term.

Nonlinear equations of equilibrium for elastic helicopter or wind turbine blades undergoing moderate deformation.

Abstract: A set of nonlinear equations of equilibrium for an elastic wind turbine or helicopter blades are presented. These equations are derived for the case of small strains and moderate rotations (slopes). The derivation includes several assumptions which are carefully stated. For the convenience of potential users the equations are developed with respect to two different systems of coordinates, the undeformed and the deformed coordinates of the blade. Furthermore, the loads acting on the blade are given in a general form so as to make them suitable for a variety of applications. The equations obtained in the study are compared with those obtained in previous studies.

Soft inplane helicopter rotor

Abstract: A helicopter rotor hub assembly (10) in which a blade grip (30) extends inboard of a yoke (14) and transfers centrifugal forces to the yoke (14) through a spherical elastomeric bearing (38). Pitch change of the grip (30) and blade (32) is accommodated by the spherical elastomeric bearing (38) which also acts as the pivot point for lead-lag motion of the blade (32) and grip (30). A lead-lag damper (22) is located inboard of the elastomeric bearing (38) and is connected between one end of the grip (30) spaced away from the lead-lag pivot and the yoke (14) for damping the lead-lag motion between the grip (30) and yoke (14).

Surface pressure measurements at the tip of a model helicopter rotor in hover

Abstract: Surface pressures were measured near the tip of a hovering single-bladed model helicopter rotor with two tip shapes The rotor had a constant-chord, untwisted blade with a square, flat tip which could be modified to a body-of-revolution tip Pressure measurements were made on the blade surface along the chordwise direction at six radial stations outboard of the 94 percent blade radius Data for each blade tip configuration were taken at blade collective pitch angles of 0, 618 and 114 degrees at a Reynolds number of 736,000 and a Mach number of 025 both based on tip speed Chordwise pressure distributions and constant surface pressure contours are presented and discussed

Simplified rotor head fairing

Abstract: A helicopter main rotor has an improved fairing enclosing the rotor hub and blade root area for drag and vibration reduction purposes. The fairing has peripheral cutouts through which the blades project provided with rounded lips on the rim of the cutouts to streamline air flowing in and out of the fairing through the cutouts. Internal partitions divide the fairing into chambers, each of which subtends one cutout, thus preventing flow of air through the fairing and attendant drag caused thereby. The rotor pylon is provided with a well through which the rotor drive shaft projects having a rounded lip adjacent the fairing to produce streamline airflow out of the well.

Reverse velocity rotor system for rotorcraft

Abstract: A blade control method and means for helicopter rotors which increases heopter speed by mixing two/rev. control motion with one/rev. cyclic and collective motion through mechanical summation is provided. Both rotary and vertical motions are imparted to an inner control shaft and these motions are combined to impart vertical oscillatory motions to the rotor head which motions increase the thrust developed by the helicopter rotor.

Helicopter rotor blade

Abstract: A rotor blade, specially for a helicopter rotor, which blade is made of reinforced plastics. This blade comprises: a solid one-piece leading edge spar embodied by strands of glass fibre filaments extending lengthwise over the whole length of the blade; a core made of a lightweight cellular substance and placed contiguous with the spar rear face; a thick sheet-like member extending around the spar and the core and comprising at least one layer of a fabric consisting of parallel glass fibres disposed askew of the blade longitudinal axis so as to form a torque-tube; a rear filling element, and a covering extending around the torque-tube and the rear filling element. The airfoil portion of this blade is centered substantially in the front quarter of the chord of the blade cross-section and it has a very high torsional stiffness.

Elastomeric rotor load reaction system

Abstract: In a helicopter rotor in which the blades are supported for rotation and articulation from the hub by elastomeric bearings and including a lead-lag damper which acts to cause radially inward blade motion during rotor braking and including the improvement of utilizing members which move pivotally in response to centrifugal force and which serve blade flapping and droop stop functions both statically and dynamically and which further are programmed as a function of rotor speed to prevent or limit radial inward motion of the blade during the rotor braking operation.

Blade retention bearing for helicopter rotor

Abstract: A hub assembly for a helicopter rotor system having a monolithic thin-wall hub shell made from composite material and including centrally apertured upper and lower wall portions integrally joined at the outer periphery of the shell and spaced apart at inner regions. The upper and lower wall portions cooperate to define a plurality of bearing sockets angularly spaced relative to the axis of rotation and connecting webs extending between angularly adjacent sockets. Elastomeric blade retaining bearing assemblies supported within the sockets retain and support rotor blades for articulate movement. Each bearing assembly has a seating surface which cooperates with an associated bearing surface on the socket in which it is contained to prevent relative rotation therebetween. Mounting flanges on the helicopter rotor shaft and retaining flanges on the hub shell are arranged to permit the rotor system to be lowered on the rotor shaft when the retaining flanges are disconnected from the mounting flanges.

Blade-mounted centrifugal pendulum

Abstract: A helicopter having a multiblade main rotor lift unit includes a plurality of pendulus mass units (48,60) mounted to oscillate in the plane of the rotor. The mass units (48,60) are connected to a rotatable shaft (24) which passes through a blade grip (14) for each blade (16). Vibrations generated in the plane of the rotor system are damped by the oscillating pendulums (22) which are tuned to minimize the vibrations in the lift unit.

Elastomeric helicopter rotor load reaction system

Abstract: A helicopter rotor having one or more blades mounted for rotation with the rotor hub and each supported therefrom by an elastomeric bearing, and having a lead-lag damper connected to the blade and the hub and extending substantially parallel to the blade axis, and including the improvement of apparatus for preventing tension loading of the elastomeric bearing as the blade moves radially inwardly as a result of rotor braking.

Rotor blade support setup for rotary wing aircraft

Abstract: A composite bearingless helicopter rotor system utilizes one single and only flexible strap as at least the principal wing support, and also preferably has also a blade pitch shaft. The line of shear centers of the flexible strap, the line of shear centers of the blade pitch shaft, and the pitch axis each lie somewhere along the one straight line, or substantially so. The cross section of the flexible strap may vary but in each case it is one which is asymmetrical relative to a line through its center of gravity and parallel to the axis of rotation of a rotor, and in each case the line of shear centers of the flexible strap, the line of shear centers of the blade pitch shaft and the pitch axis all are spaced from the line of centers of gravity of the flexible strap. A number of examples of varying specific overall configurations of flex straps and blade pitch shaft will be found described and illustrated herein, of which an example is that in FIG. 15 and the text relating thereto, in which at the leading edge there is a gapped dome in the cross section of the flex strap, following which the cross section comes straight back in legs on both sides with a web straight across between in the further part and the blade pitch shaft is a solid shaft at the trailing edge between the ends of the legs.

Noise due to rotor-turbulence interaction

Abstract: A procedure for calculating the noise due to turbulent inflow to a propeller or helicopter rotor in hover is summarized. The method is based on a calculation of noise produced by an airfoil moving in rectilinear motion through turbulence. At high frequency the predicted spectrum is broadband, while at low frequency the spectrum is peaked around multiples of blade passage frequency. The results of a parametric study of the variation of the noise with rotor tip speed, blade number, chord, turbulence scale, and directivity angle are given. A comparison of the theory with preliminary experimental measurements shows good agreement.

On methods for application of harmonic control

Abstract: The paper presents data which confirm the effectiveness of higher harmonic blade pitch control in substantially reducing helicopter rotor vibratory hub loads. The data are the result of recent tests on a 2.7-m model conducted in the Langley Research Center's transonic dynamics wind tunnel. Several predictive analyses developed in support of the NASA program are shown capable of accurately predicting both amplitude and phase of the higher harmonic control input required to nullify a single 4/rev force or moment input. The use of multiple blade feathering inputs in the design of a flightworthy higher harmonic control system is discussed.

Blade stress calculations -- mode deflection vs. force integration

Abstract: Two methods for calculating blade stresses from aeroelastic responses, mode deflection and force integration are described and compared with respect to accuracy and ease of operational usage. Analytic results are presented for a realistic bearingless helicopter rotor configuration which is capable of experiencing concentrated shears and moments at discrete spanwise points on the blade. It is generally concluded that, when properly implemented, the force integration method yields more reliably accurate stress calculations than the mode deflection method.

Vibration absorber for helicopter

Abstract: A vibration absorber for a helicopter rotor includes a hub portion, a mass portion encircling the hub portion and a plurality of resilient arms extending in a spaced-apart spiral pattern between the hub and the mass. In operation, the hub portion of the vibration absorber is mounted on a helicopter rotor so as to be rotated with the rotor about an axis of rotation of the rotor, the resilient arms ensuring that the mass is capable of equal resilient deflection in any direction within its plane of rotation whereby the vibration absorber simultaneously reduces in-plane vibration forces of different frequencies.

Helicopter rotor and blade connection

Abstract: This invention relates to helicopter rotors in which the rotor blades are supported by flexure members and resilient means to permit flap and lead/lag movements. Hitherto such support means have been subjected to centrifugal forces thus complicating their design and requiring compromises such that optimum properties in flap and lead/lag planes have been unattainable. In this invention, each blade is attached by an individual spindle 19 supported in an elastomeric bearing 21 secured in a hollow rotor hub 12. Arms 26, flexible in a flapping plane, extend from the rotor hub 12 and a bearing 46 at the outer end of each spindle is attached to its associated arms through resilient means 30 to permit lead/lag movement of the spindle 19 relative the arms 26. In operation, centrifugal loads are transmitted to the hub 12 so that the arms 26 and resilient means 30 can be designed to provide optimum properties in the flap and lead/lag planes. Furthermore, bending of the spindles 19 due to such blade movements significantly reduces corresponding deflections of the elastomeric bearing 21.

Flexible connector for helicopter rotor blade - has fork member with elastomer bearing and folding pivot bolts

Abstract: The rotor of a helicopter has rotor blade holders with integral fibre composite material. The blade holders have each a fork ensuring axial thrust load and allowing folding of the helicopter blade. The blade bearing does not enclose the blade root and the advantage of double folding is ensured. The blade root (40) is connected by means of a compressive loaded connection transverse bolt (43) and by elastomer bearing (44) with the fork (41). The fork root has plug with connector loaded in compression. The blade is pivoted on two pivot bolts (46). On removal of one bolt the blade can be folded on the remaining bolt.

An improved computational procedure for determining helicopter rotor blade natural modes

Abstract: An existing computer program, used for predicting the natural frequencies and mode shapes of helicopter rotor blades, was refined to improve program accuracy and versatility. The program is based on the Holzer-Myklestad approach adapted for rotating beams. Coupled vertical (out-of-plane), horizontal (in-plane), and torsional mode characteristics were determined for a variety of hub and blade configurations. The resulting program is documented by presenting the recursion equations and techniques for determining natural frequencies and mode shapes, input data requirements, and descriptions of various program outputs. The accuracy of the program is demonstrated by comparing computed results with exact solutions to classical problems and experimental data.

An experimental investigation of hingeless helicopter rotor-body stability in hover

Abstract: Model tests of a 1.62 m diameter rotor were performed to investigate the aeromechanical stability of coupled rotor-body systems in hover. Experimental measurements were made of modal frequencies and damping over a wide range of rotor speeds. Good data were obtained for the frequencies of the rotor lead-lag regressing mode. The quality of the damping measurements of the body modes was poor due to nonlinear damping in the gimbal ball bearings. Simulated vacuum testing was performed using substitute blades of tantalum that reduced the effective lock number to 0.2% of the model scale value while keeping the blade inertia constant. The experimental data were compared with theoretical predictions, and the correlation was in general very good.

Pitch control system for helicopter rotor blades

Abstract: A system is disclosed for controlling the pitch angle of helicopter rotor blades to compensate for differences in lift generated by advancing and retreating blades during translational flight or due to gust loading of individual blades. Each rotor blade is free to move about its pitch axis to achieve an adjustable balance between a pitching moment exerted by the blade and centrifugal force exerted by weights rotating with the blade system. Collective and cyclic pitch commands from the pilot are transmitted through a conventional swash plate to a linkage which adjusts the relationship of the pitching-moment and centrifugal forces which produces a balanced condition. The system maintains near-constant blade lift as the blade rotates, and minimizes blade distortion and asymmetric coning, thereby significantly reducing vibration arising from a shifting of the center of mass of the rotor-blade system away from the center of rotation.

Helicopter rotor system related vibration amplitude detecting system

Abstract: What follows is a description of a system for detecting rotor system related vibration in a helicopter. The vibration detecting system comprises three essential units: a sensing unit or device; a control unit; and a model module. Actually, any number of sensing units can be utilized depending upon the number of locations at which the vibration is to be monitored. The sensing unit generates a signal indicative of the rotor system related vibration and delivers this signal to the control unit which operates on the generated signal and produces an output signal within a selected frequency and amplitude range. The output signal is then displayed, for example, on a scale meter. The model module is adaptable for compatibility with the dynamic rotor system of a given helicopter model and is utilized as such to set the limits of the selected frequency and amplitude range of the control unit. The model module is readily connected to and removed from the housing of the control unit.

Analogue computer simulation of a rotor system containing a transverse crack

Abstract: This paper describes the analogue computer simulation of a complete rotor and bearing system. The physical equations of motion of the rotor with its shaft are derived in rotating coordinates and those of the bearings in stationary coordinates. The system is modelled to have four degrees of freedom, taking into consideration the flexibility, damping, and cross- coupling of the hydrodynamic bearings. The model can be reduced to two degrees of freedom by assuming that the rotor is rigidly supported. The simulation of a transverse fatigue crack in the rotor shaft is pre sented and its effect on the dynamic response of the system is briefly examined.

Hingeless rotor system

Abstract: The present hingeless rotor system for helicopters has carrier spars or b each of which is subdivided into three functional sections substantially forming the blade neck connecting the lift producing portion of the respective rotor blade to the rotor mast The carrier spar sections form a single structural blade neck member extending for about two tenths of the rotor radius from the rotor mast to the lift producing blade portions About one half of the length of the blade neck adjacent to the rotor mast form the connecting zone proper The spar sections comprise bands forming respective loops in the connecting zone proper Two of the band loops are connected one above the other and a third band loop is displaced in the direction of rotation of the rotor from the first mentioned band loops The third band loop is releasably connected to the rotor mast to allow a folding of the respective rotor blade The spar sections are interconnected by shear resistant bridging webs having a T- or X cross-section

Full-scale testing of an Ogee tip rotor

Abstract: Full scale tests were utilized to investigate the effect of the ogee tip on helicopter rotor acoustics, performance, and loads. Two facilities were used: the Langley whirl tower and a UH-1H helicopter. The text matrix for hover on the whirl tower involved thrust values from 0 to 44 480 N (10,000 lb) at several tip Mach numbers for both standard and Ogee rotors. The full scale testing on the UH-1H encompassed the major portion of the flight envelope for that aircraft. Both near field acoustic measurements and far field flyover data were obtained for both the ogee and standard rotors. Data analysis of the whirl tower test shows that the ogee tip does significantly diffuse the tip vortex while providing some improvement in hover performance at low and moderate thrust coefficients. Flight testing of both rotors indicates that the strong impulsive noise signature of the standard rotor can be reduced with the ogee rotor. Analysis of the spectra indicates a reduction in energy in the 250 Hz and 1000 Hz range for the ogee rotor. Forward flight performance was significantly improved with the ogee configuration for a large number of flight conditions. Further, rotor control loads were reduced through use of this advanced tip rotor.

Helicopter rotor head mounting assembly

Abstract: A helicopter rotor head mounting assembly includes a rotor drive shaft and rotor hub splined to the shaft for rotation therewith. The rotor hub includes a central bore having an internal tapered portion which is forced downwardly upon a mating tapered portion of the rotor drive shaft. An annular split cone is mounted between the rotor shaft and the hub at a bottom portion of the hub bore and the split cone is forced upwardly into firm engagement with the hub and drive shaft by means of a plurality of individual loading means which are carried by an annular plate affixed to the bottom of the hub. The loading means are individually adjustable and apply loading force directly onto the split cone to thereby assure a secure, non-vibrational rotation of the hub with the rotor shaft.