Showing papers on "Helicopter rotor published in 1972"

Subsonic and Transonic Potential Flow over Helicopter Rotor Blades

Abstract: Compressible potential flows over nonlifting, hovering helicopter blades are described by suitable linear and nonlinear equations of motion for subsonic and transonic cases, respectively. Analytical and numerical results are presented for the linearized subsonic three-dimensional flow in the tip region. When the tip Mach number is transonic, the flowfield is calculated using a computational method that is a formal extension to three dimensions of recently developed nonlinear two-dimensional relaxation schemes. Calculations are presented for rectangular blades with 6% thick biconvex sections. Calculations show the relative importance of tip Mach number and aspect ratio on the growth and extent of shock waves in the tip region, and indicate a significant reduction in shock strength with decreasing aspect ratio. ELICOPTER rotor blades that can operate at high tip Mach numbers without the penalties usually associated with super-critical flow would permit high-forward speeds at relatively low-advance ratios. As the advance ratio decreases, the portion of the retreating blade that is stalled would be diminished, the portion of the retreating blade that is unstalled could operate at smaller angles of attack, and the advancing blade could generate lift without also generating an excessive rolling moment. Vibratory loads would be diminished, and severe unsteady pitching moments on the lifting portion of the retreating blade would decrease if that portion could usefully operate at higher tip speeds and smaller angles of attack. The design of rotor blade geometries that can support a transonic flowfield that is nearly shock-free requires a method for computing the detailed pressure distribution on the blade for a specified rotor geometry and advance ratio. However, the unsteady aerodynamic environment of a lifting rotor that is operating at or near the critical Mach number is extremely complicated. It is easy to appreciate the difficulty of describing such flowfields when one recalls that we do not yet have an adequate model for describing the details of the flow about a lifting, hovering rotor operating in an essentially incompressible flow. Our purpose here is therefore to focus on a portion of the high-speed rotor problem that is sufficiently limited to allow for some analytical and numerical treatment, but which retains important linear and nonlinear features of the flow. We consider a two-bladed nonlifting and hovering rotor operating in the range of tip Mach numbers where compressibility effects are important. The pressure distribution near the rotor tip is studied. In this region, the flowfield is fully three-dimensional, and blade element (strip) theory is in no sense a reasonable approximation to the rotor aerodynamics. It is assumed that the flow is inviscid. The flow is steady relative to a coordinate system attached to the blades. Analytical and numerical results presented here are based on suitable approximations to the full three-dimensional potential equation for compressible flow. When the tip Mach number is subcritical, the potential equation is the usual acoustic

Anti-torque, propulsion, and directional control system

Abstract: A helicopter devoid of a conventional exteriorly mounted antitorque rotor and having at the rear end of its aft fuselage a system utilized for auxiliary propulsion and anti-torque and directional control. This combination includes an axial-flow variable-pitch fan, a diffuser, an enlarged duct immediately contiguous to the diffuser, and sets of controllable variablegeometry, series-related, articulatable vanes mounted in a boattail configuration at the rear or egress end of the aft fuselage. The fan is located immediately forward of the diffuser, with the duct extending from the diffuser to the vanes, and an air inlet through which the flow of air is induced by the fan is disposed immediately forward of the fan. The vanes operate to dually provide auxiliary thrust and anti-torque control. In operating as an anti-torque control, the flowing air in the enlarged duct is made to converge and accelerate as such air is turned by positioned articulated vanes to achieve maximum efficiency in the production of the required anti-torque force. A pair of spaced elongated slots, with automatically closing lips, extend longitudinally of the fuselage skin, along the enlarged duct and below the sweep of the main rotor, to augment the performance of the basic anti-torque system by inducing a circulation of air around the fuselage which, with the downwash from the main rotor, produces an additional anti-torque force. By varying the pitch of the fan, a large power drain from the main rotor system during critical power-off auto-rotation descents is prevented. The blade pitch also controls auxiliary propulsion thrust, independently of yaw control, while skin friction losses of air flowing in the enlarged duct are minimized.

Hingeless Helicopter Rotor Response with Nonuniform Inflow and Elastic Blade Bending

Abstract: A generalized harmonic balance theory for the steady-state, linear, response characteristics of a hingeless rotor in forward flight is presented and evaluated with the aid of recent experimental data. Comparisons of several approximate representations for the rotor blade revealed that the simple rigid hinged blade is inadequate except at very low advance ratios, or high flap frequencies. For typical values of flap frequency the first two elastic flap bending modes are required for accurate response predictions. Simplified models of the nonuniform induced inflow were derived, using momentum arid vbrtex theory, and found to be the most important factor in improving correlation with the data. An empirical inflow model was developed from the experimental data by obtaining an inverse solution of the present theory. This inflow model was found to be in reasonable agreement with the simple inflow theories for low advance ratios but revealed large unexpected variations in the induced inflow of the rotor for advance ratios near 0.8. a B b c CT d Cm e epc El J [L] m [M] [N] n p r, R U v'i v0 w

Four blade main rotor control power coupling

Abstract: Helicopters traveling in level flight are considered to be operating in a one ''''g'''' condition; that is, the rotor is producing lift equal to the vehicle weight. In a helicopter equipped with a gimbaled, teetering rotor, which cannot transmit a rotor moment into the mast head, a control moment about the aircraft center of gravity, a requirement to command a change in aircraft attitude, is obtained by tilting the rotor and hence its thrust vector. Thus, the control moment is a function of rotor thrust and tilt angle. When such an aircraft is subjected to a sudden descending maneuver, the rotor thrust will be reduced toward a zero or negative ''''g'''' condition. Consequently, the control moment capability will be reduced to zero and the aircraft becomes uncontrollable. However, the rotor is capable of producing a moment, if cyclic pitch is introduced to the rotor through a normal helicopter control system. This moment may be transmitted across a gimbal, down the supporting rotor mast to exert a controlling moment about the helicopter center of gravity by means of an elastomeric hub spring, connecting the gimbaled rotor hub to the rotor mast. This spring attaches to a first flange as part of the rotor hub yoke, and a second flange bolted to the supporting mast.

An Analytical Evaluation of the Controllable Twist Rotor Performance and Dynamic Behavior

Abstract: : A detailed analytical evaluation is made of a new rotor system with torsionally elastic blades and dual controls-the controllable twist rotor (CTR). The controls consist of conventional pitch horn linkages at the inboard end and an aerodynamic control flap at the outboard end. The analysis involves an aerolastic loads digital computer program which was developed to account for the blade response modes and blade control modes on either single or dual control rotors. Six response modes are included: blade flapping, blade feathering, blade lagging, blade flapwise bending, blade torsion, and control flap feathering. The two control modes are included separately and incorporate control system stiffness so that control loads are calculated. The aeroelastic analysis includes nonlinear inertia distributions, nonlinear airfoil characteristics, and inertial and mechanical coupling among the modes. The analysis outputs transient responses for stability evaluation and steady-state blade load and angle-of-attack distributions, blade dynamics, and rotor performance for each trimmed flight condition.

Dynamic Nonlinear Elastic Stability of Helicopter Rotor Blades in Hover and in Forward Flight

Abstract: Equations for large coupled flap-lag motion of hingeless elastic helicopter blades are consistently derived. Only torsionally-rigid blades excited by quasi-steady aerodynamic loads are considered. The nonlinear equations of motion in the time and space variables are reduced to a system of coupled nonlinear ordinary differential equations with periodic coefficients, using Galerkin's method for the space variables. The nonlinearities present in the equations are those arising from the inclusion of moderately large deflections in the inertia and aerodynamic loading terms. The resulting system of nonlinear equations has been solved, using an asymptotic expansion procedure in multiple time scales. The stability boundaries, amplitudes of nonlinear response, and conditions for existence of limit cycles are obtained analytically. Thus, the different roles played by the forcing function, parametric excitation, and nonlinear coupling in affecting the solution can be easily identified, and the basic physical mechanism of coupled flap-lag response becomes clear. The effect of forward flight is obtained with the requirement of trimmed flight at fixed values of the thrust coefficient.

Helicopter rotors comprising a reinforced plastics hub

Abstract: A helicopter rotor has a reinforced plastics hub. Arms extend radially outwardly from the hub each arm carrying through, the intermediary of a sleeve, a blade. The arms are reinforced with fibres.

Measurements of Velocity Components in the Wake of a Full-Scale Helicopter Rotor in Hover

Abstract: : The report presents three-component wake velocity measurements made with a split-film total vector anemometer. The measurements were made in the wake of a full-scale OH-13E helicopter rotor which was mounted on a 60-foot rotor test tower at Mississippi State University. Time-averaged velocity distributions along wake radii at various distances below the rotor disk were measured for two conditions of disk loading and three combinations of blade pitch and rotor speed. Instantaneous velocity measurements were made across the helical vortex trails to investigate the effects of blade pitch and rotor speed on vortex structure, core size, transport velocity, and distribution of axial and tangential velocity components within the vortices.

Concentric tube spring rotor mount

Abstract: The helicopter main rotor is mounted on a rotating support mast by feathering bearings and flapping bearings to provide pitch change control and to tilt the rotation plane of the rotor Rotor tilting is provided by the flapping bearings coupling a blade yoke to the mast The flapping bearings are elastomeric elements fitted onto pivot shafts on a trunnion mounted on the mast These elastomeric elements also mount in bearing blocks bolted to the rotor yoke To increase the control power about the pitch and roll axis of the aircraft, particularly at rotor loadings of less than one ''''G,'''' a concentric tube torsion spring interconnects pivot shafts of the trunnion to respective bearing blocks Each of the concentric tube torsion springs includes an inner tube mounted within an outer tube and joined together at the outboard ends The inboard end of the inner tube connects to the pivot shaft and the inboard end of the outer tube connects to the bearing block

Vtol aircraft with cruciform rotor wing

Abstract: An aircraft having a combined rotary and fixed wing providing aerodynamic support in vertical take-off and landing and in high speed cruising flight. The wing is cruciform in configuration, with four similar arms extending radially from a center body, the outer portions of the arms being pivotal and controllable in the manner of helicopter rotor blades in the rotating mode of the wing. In fixed wing mode the wing is locked with two arms longitudinal to the aircraft and the other two extending laterally, one or both of the lateral arm portions being used for roll control in forward flight. The wing arms may have air brakes to slow rotation and stop the wing in proper alignment. A common turbojet power source provides propulsion for wing rotation and cruising flight.

Helicopter rotor balancing method and system

Abstract: The method of balancing a rotor having multiple blades and defining an axis of rotation, and wherein structure proximate the rotor is subject to vibratory motion due to dynamic unbalance of the rotating rotor, the method employing a vibration pickup and a stroboscope, the method including: A. ATTACHING THE PICKUP TO SAID STRUCTURE, OPERATING THE PICKUP TO PRODUCE A VIBRATION SIGNAL WHICH IS A FUNCTION OF SAID OSCILLATORY MOTION, AND PROCESSING SAID SIGNAL TO PRODUCE A CORRESPONDING OSCILLATORY OUTPUT SIGNAL, B. TRIGGERING THE STROBOSCOPE IN SYNCHRONISM WITH SAID OUTPUT SIGNAL AND DIRECTING THE STROBOSCOPE AT THE ROTOR TO PRODUCE FLASHES REPEATEDLY ILLUMINATING A TARGET ROTOR BLADE AT A CHARACTERISTIC ANGULARITY WITH RESPECT TO SAID AXIS, AND C. VARYING THE WEIGHTING OF THE ROTOR AS A FUNCTION OF THE MAGNITUDE OF SAID SIGNAL AND SAID CHARACTERISTIC ANGULARITY TO ACHIEVE SUBSTANTIAL BALANCE.

Compound bearing for connecting a helicopter blade to a helicopter rotor

Abstract: A compound bearing for connecting a helicopter blade to a helicopter rotor, such as connecting the blade end-shaft to the rotor connected lead-lag damper so as to accommodate a large degree of blade pitch change motion and part misalignment due to other blade motions, which bearing is maintained within a minimum space envelope. The compound bearing includes a slip-type journal bearing enveloping the blade end-shaft and a spherical, laminated elastomeric bearing enveloping the journal bearing and connected to the rotor connected part, such as the lead-lag damper.

Articulated helicopter rotor

Abstract: One or more helicopter blades are supported from a helicopter rotor in articulated fashion by a spherical elastomeric bearing which permits blade motion about the blade lead-lag axis and which includes a piston-cylinder type lead-lag damper. Lead and lag stops positioned selected distances from the blade shaft so that during rotor start-up operation, the uncentered blade will increase in lead angle to eventually be supported by the lag stop and the bottomed-out damper piston and so that, during rotor braking operation, the uncentered blade will increase in lead angle and eventually be supported by the lead stop and the bottomed-out lead-lag damper. This construction does not include a centering bearing.

Elastomeric helicopter rotor head with dynamic and static blade coning and droop stops

Abstract: An elastomeric helicopter rotor having a blade mounted for universal motion about the intersection of the blade pitch change, flapping and lead-lag axes and including blade coning and droop stops, including a ring member rotatably mounted about the pitch change axis on the blade and coning and droop stops constituting segments of circular members whose centers lie on the blade lead-lag axes and with the ring member and coning stop and droop stop members presenting mating surfaces to one another so that as the blade moves in lead-lag motion while the blade ring member is in line or surface mating contact with either the coning stop surface or the static or dynamic droop stop surfaces, a relative rotation will be established therebetween for full support of the blade throughout the lead-lag motion, without affecting blade pitch angle or flapping angle, without preventing independent blade pitch change, and without scuffing of parts.

System for continuously trimming helicopter rotor blades during flight

Abstract: A vibration sensor is mounted on the air frame to provide a continuous output proportional to the unbalance of the rotor blades with respect to one another. This output is fed to a trim tab on the particular blade causing the unbalance to alter that blade''s lift slightly until the balance is reduced to an acceptable level.

Investigation of Airfoil Dynamic Stall and Its Influence on Helicopter Control Loads

Abstract: : Measurements were made of the unsteady normal force and pitching moment on an NACA 0012 airfoil model oscillated both sinusoidally and nonsinusoidally over a range of incidence angles, including a substantial penetration into stall. The sinusoidal normal force and pitching moment data were reduced and tabulated as functions of the angle of attack, the angular velocity parameter, and the angular acceleration parameter. This generalized form of the data was used to reconstruct the measured sinusoidal aerodynamic response of the model airfoil with excellent results. Additional correlations were made using nonsinusoidal pitch schedules which included periodic ramp changes in angle of attack and a flexured angular blade response to a one-per- rev sinusoidal incidence angle change typical of that for a helicopter blade. The agreement between predicted and measured normal force and moment loops was very good for the ramp motion.

A method for analyzing dynamic stall of helicopter rotor blades

Abstract: A model for each of the basic flow elements involved in the unsteady stall of a two-dimensional airfoil in incompressible flow is presented. The interaction of these elements is analyzed using a digital computer. Computations of the loading during transient and sinusoidal pitching motions are in good qualitative agreement with measured loads. The method was used to confirm that large torsional response of helicopter blades detected in flight tests can be attributed to dynamic stall.

Articulated helicopter rotor utilizing plural elastomeric bearings for articulated support of the blade from the rotor hub

Abstract: An articulated helicopter rotor in which the blade is supported for articulated motion from the rotor by elastomeric bearings and including an elastomeric shear bearing adapted to take rotor in-plane and rotor out-of-plane shear loads.

Helicopter rotor blade

Abstract: A helicopter rotor blade having a simple and economical construction which includes internal counterweight means to shift the center of gravity to a forward location A skin is wrapped around a former which has opposed top and bottom surfaces The former cradles the counterweight member at the leading edge of the blade The specific gravity of the counterweight member is greater than that of the former, and the trailing edges of the skin are joined together to close out the blade The skin may be attached to the former by conventional shanked fasteners, or by resilient bonding means The counterweight member may be pinned to the former to hold the blade together in the unlikely event of its fatigue failure

Investigation of the Vortex Noise Produced by a Helicopter Rotor

Abstract: : Karman-street-type vortex shedding from a lifting surface was analyzed as a source of noise from a helicopter rotor in hover and forward flight. Experimental pressure-time histories were analyzed, and high resolution spectra were developed over a frequency range of 0 to 5000 Hz using a 0.7-Hz filter. The results of the investigation indicated that 'vortex noise' is the major source of acoustic radiation from a helicopter rotor in hover or low-speed flight and that it is concentrated in the frequency range of 200 to 500 Hz.

Experimental Investigation of Effects of Blade Section Camber and Planform Taper on Rotor Hover Performance

Abstract: : An experimental and analytical investigation was conducted to determine the effects of blade section camber and blade planform taper on helicopter rotor hover performance and to assess the accuracy of several theoretical methods in predicting such performance. The tests were conducted using small scale model rotors (nominally 4 feet in diameter) operating in and out of ground effect at full-scale tip speeds. The NACA 23112 airfoil, which is a section specifically designed to produce very low pitching moments such as required for helicopter blade applications, was selected for the cambered blade. A nominal taper ratio of 2:1 was selected for the tapered blade tests.

Main rotor free wake geometry effects on blade air loads and response for helicopters in steady maneuvers. Volume 1: Theoretical formulation and analysis of results

Abstract: A mathematical model and computer program were implemented to study the main rotor free wake geometry effects on helicopter rotor blade air loads and response in steady maneuvers. The theoretical formulation and analysis of results are presented.

The pressure-jet helicopter propulsion system

Abstract: The helicopter rotor probably represents one of the most difficult of all power-transmission problems in the fact that high power is required at relatively low rotational speeds. High power at low speed defines high torque, and high torque in a mechanical transmission requires large, heavy gears. In addition, since every action has an equal and opposite reaction, the shaft torques must be reacted, and this requires a tail rotor system for the shaft helicopter. A jet-propulsion system offers certain distinct advantages in comparison with a shaft-drive system for powering a helicopter rotor. It delivers a driving torque to the rotor in the form of a tangential force at the tip of the rotor rather than by twisting a shaft at the centre of the rotor.

Helicopter Rotor Wake Geometry and Airloads and Helicopter Rotor Wakes

Abstract: : Computer programs have been developed for the calculation of helicopter rotor tip vortex geometry in hover and forward flight and for the calculation of helicopter rotor harmonic airloads in forward flight. Calculated forward flight tip vortex geometries compare well in general with experimental smoke studies although there are differences in detail. The hovering tip vortex geometry agrees qualitatively with experiment but does not move downward fast enough. Airloads were computed using both the classical rigid wake assumption and the distorted tip vortex geometry obtained from the computer program. When compared with experimental airloads measurements the rigid wake airloads give better results than the distorted wake airloads.