Showing papers on "Helicopter rotor published in 1975"

Computation of helicopter rotor wake geometry and its influence on rotor harmonic airloads

Abstract: Thesis. 1975. Ph.D.--Massachusetts Institute of Technology. Dept. of Aeronautics and Astronautics.

Helicopter power plant control

Abstract: A control system for a helicopter power plant with three power units driving the lifting rotor system. Each power unit comprises a gas-coupled gas turbine engine, a hydromechanical fuel control receiving electrical inputs to set a gas generator governor, an engine electronic control, and engine accessories. Power plant operation is normally controlled by the pilot through condition levers and switches on a control quadrant, through a collective pitch control for the rotor blades, and through a master beeper switch which is operable to trim the setting of a rotor speed governor and the gas generator governors. A condition lever transmits a speed command signal through the electronic control of each power unit to its gas generator governor. A collective pitch signal from the rotor control system is another factor in setting the gas generator governor of each unit. The power plant includes a power management control which equalizes the power outputs of the engines and includes the rotor isochronous governor. The power management control receives inputs of rotor system speed command and actual rotor speed, and of the torques of the individual power plants and transmits a governor trim signal to the several electronic controls. All signals are transmitted electrically between the rotor system, control panel, power management control, and engine electronic controls. The signal representing condition lever power setting decreases with increasing power demand. The collective pitch signal decreases with increasing pitch. A failure to transmit either signal thus represents a high power request -- a fail-safe feature. A starting system for each engine includes logic circuits to test the operation of the electrical signal transmission system and closure of a fuel shutoff valve as a prerequisite to starting of the engine.

Isolated Airfoil-Tip Vortex Interaction Noise

Abstract: An experimental investigation was conducted in the UARL Acoustic Research Tunnel to define the noise characteristics associated with the interaction of a stationary tip vortex and a downstream stationary airfoil. This model test geometry simulated, in its simplest form, the tip vortexblade interaction which occurs on single rotor helicopters during hover. For moderate to high lift test conditions, the vortex-airfoil interaction was found to cause local blade stall with an attendant increase in the blade far-field noise. These results indicated that this interaction may be an important source of helicopter broadband noise during hover. Cross-correlation measurements conducted amongst surface-mounted and far-field microphones demonstrated that the operative noise mechanism was "trailing edge noise" arising from the interaction of stall generated eddies with the airfoil trailing edge. This mechanism would be expected to be responsible for increased noise at stall conditions in other, nonrotary wing, applications.

Wind tunnel investigation of helicopter-rotor wake effects on three helicopter fuselage models

Abstract: The effects of rotor wake on helicopter fuselage aerodynamic characteristics were investigated in the Langley V/STOL tunnel. Force, moment, and pressure data were obtained on three fuselage models at various combinations of windspeed, sideslip angle, and pitch angle. The data show that the influence of rotor wake on the helicopter fuselage yawing moment imposes a significant additional thrust requirement on the tail rotor of a single-rotor helicopter at high sideslip angles.

Aircraft with retractable rotor wing

Abstract: An aircraft having a rotor wing which in one position is locked against theirframe for forward flight, and in another position is extended clear of the airframe. In the extended position the wing is rotated and the tip portions are controllable in the manner of a helicopter rotor. A single power source provides propulsive and wing rotation power. In the preferred form the power source is a turbojet engine and the wing is rotated by a tip jet powered drive beam, separated from the wing to avoid aerodynamic interference. For forward flight the drive beam is enclosed in the airframe as the wing is retracted to fixed position.

Helicopter yaw and propulsion mechanism

Abstract: A mechanism is provided for controlling the directional heading of a helicopter, particularly during hover and autorotation. In addition the mechanism provides additional thrust which enables the helicopter to increase its maximum velocity. The mechanism includes three nozzles with valves positioned therein, located in the rearward portion of the helicopter fuselage. Two of the nozzles are located on opposite sides of the aircraft fuselage to provide thrust for use in controlling the directional heading of the aircraft. This thrust is also used to counteract the torque generated by the driving of the lifting rotor of the helicopter. The third nozzle is positioned so that the thrust generated by it will increase the forward velocity of the helicopter. A propeller or fan located in the body of the aircraft forces air from intakes through the fuselage to the nozzles. The position of the valves is determined by suitable control linkage.

Tow-position helicopter main rotor

Abstract: A helicopter main rotor which can be supported in two positions of elevation above the fuselage, and apparatus and method for moving the helicopter rotor with blades attached from a first elevated position above the fuselage wherein the rotor is supported from an extension shaft projecting from a first shaft to a second elevated position above the fuselage wherein the rotor is supported from the first shaft.

Automatic one-per-rev control system

Abstract: Vibration in the helicopter fuselage produced through the helicopter rotor at a frequency of one for each rotation of the rotor (commonly identified in the art as a one-per-rev vibration) is controlled by applying a correction signal to the primary cyclic control system of the rotor assembly. To generate the correction signal, a transducer responds to the one-per-rev vibrations to generate a signal coupled to circuitry for generating a vibration related signal dependent upon the amplitude and phase of the one-per-rev vibrations. The vibration related signal is applied to a sample and store network wherein it is demodulated by a reference signal related to the rotational speed of the helicopter to establish a control voltage which varies at a rate in dependence upon variation in the maximum amplitude of the one-per-rev vibrations. This control voltage is integrated and modulated by the reference signal thereby resulting in the correction signal for application to the primary cyclic control system. The reference signal for both the demodulation and modulation steps is generated by combining a reference pulse having a phase corresponding to a desired cyclic pitch generated for each revolution of the helicopter with a pulse train synchronized with the reference pulse at a repetition rate related to rotation of the helicopter rotor. In one embodiment of the invention a manually adjustable ride control voltage is generated that has a magnitude related to a desired vibration correction. This ride control voltage is switched alternately with the control voltage to be modulated by the reference signal.

Helicopter rotor blade

Abstract: A helicopter main rotor has blades with an inboard airfoil having a high lift-drag ratio for efficient hovering and a "shock-free" outboard airfoil for high speed cruise. Each rotor blade has a structural section composed of metal spars and metal skins forming torque boxes accompanied by a nonstructural fairing and after body. Secured within the torque boxes and the after body are honeycomb cellular structures of a metallic and nonmetallic material, depending on the area wherein the cellular structure is secured. At a point along the blade radius, between 60% and 75%, the upper metal skin of the aft torque box changes from an outer contour surface of the blade to a flat surface parallel to and above the cord line. A fairing of fiberglass skin is secured over a nonmetallic cellular structure over the flat surface of the upper metal skin to complete the upper outer airfoil contour outboard of the 75% radius dimension. The lower outer airfoil contour is the metal skin of the aft torque box followed by a fiberglass skin joined at a trailing edge with the fiberglass upper surface.

Helicopter power transmission systems

Abstract: This invention relates to power transmission systems for helicopters having two or more engines, the transmission system comprising for each engine, drive shafting and gearing including a final drive gear for transmitting torque from each engine to a common output gear for connection to the rotor system, and an idler gear interconnecting the respective gearings to distribute available torque amongst the full complement of final drive gears. A particular embodiment is shown and described for a twin-engined helicopter in which an additional final drive gear is meshed with the output gear and driven from the idler gear so that, even in the event of an engine failure the available torque from the remaining engine is transmitted through all three final drive gears to the rotor system. In the particular embodiment the drive to the additional final drive gear is also arranged for connection to a tail rotor.

Droop stops for helicopter rotor having elastomeric bearings

Abstract: A fully articulated helicopter rotor having an elastomeric type main bearing which is reactive of all blade motions about the intersecting blade pitch change, lead-lag, and coning axes, including blade droop supporting and limiting members for static and dynamic operating conditions wherein coupling influences between the separate blade motions are precluded. In reacting these blade weight loadings and dynamic thrust forces, the droop members specifically provide full area contact bearing surfaces to distribute the large forces from the blades into the rotor hub proper while avoiding scuffing or rolling contact between the parts, and the high bearing stresses which would result from either point or line contact only.

Helicopter rotor balancing apparatus and method

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 and an out-of-track condition of the rotor blades, the method employing two vibration pickups, the method including: A. operating the pickups to produce vibratory signals corresponding to up and down vibratory motion of the structure at port and starboard sides of said axis, B. combining said signals to produce a resultant oscillatory output signal characteristic of rotor vibratory motion due substantially to only one of said conditions, C. using said resultant output signal to alleviate said one condition.

Philosophy, design, and evaluation of soft-mounted engine rotor systems

Abstract: Gas turbine engine rotors are conventionally supported by bearings mounted on relatively stiff supports. The resulting vibratory loads and deflections can be reduced significantly by judiciously soft-mounting the bearings through squirrel cages and/or squeeze films. In addition to minimizing loads and stresses in an engine, it is important that clearances during conditions of maneuvers, thermal bow, and rotor whirl due to unbalance (even under extraordinary conditions such as loss of blades) be controlled. For high-speed rotors, it becomes necessary to support the rotors on resiliently mounted bearings to achieve vibration-free, long-life, close-clearance engines. In this paper, the design philosophy, criteria, and methods of evaluation for soft-mounted turbine engine rotor systems used in General Electric aircraft engine design are described. A major constituent of this method is a computer program for system vibration and static analysis [VAST]. This program is capable of finding natural frequencies, normalized modes, and responses due to any distribution of exciting forces considering gyroscopic and shear-deflection effects. Aircraft mounting and excitations from the helicopter rotor are also included in the computer analysis. General Electric's T700 turboshaft engine, under development for the U.S. Army, serves to illustrate the squeeze film, softmounting concept of design. Results from tests of the T700 engine, Advanced Technology Axial Centrifugal Compressor (ATACC), T64 turboshaft, TF34 turbofan, and other engines are summarized verifying the advantages of soft-mounted rotor systems.

Aeroelastic stability of trimmed helicopter blades in forward flight

Abstract: Equations for moderately large amplitude coupled flap-lag motion of a torsionally rigid hingeless elastic helicopter blade in forward flight are derived. Quasi-steady aerodynamic loads are considered and the effects of reversed flow are included. By using Galerkin's method the spatial dependence of the problem is eliminated and the equations are linearized about a time dependent equilibrium position determined from the trimmed equilibrium position of the rotor in forward flight. In the first trim procedure the rotor is maintained at a fixed value of thrust coefficient with forward flight and horizontal and vertical force equilibrium is satisfied in addition to maintaining zero pitch and roll moments. The second trim procedure maintains only zero pitch and roll moment simulating conditions under which a rotor would be tested in the wind tunnel.

The development of experimental techniques for the study of helicopter rotor noise

Abstract: The features of existing wind tunnels involved in noise studies are discussed. The acoustic characteristics of the MIT low noise open jet wind tunnel are obtained by employing calibration techniques: one technique is to measure the decay of sound pressure with distance in the far field; the other technique is to utilize a speaker, which was calibrated, as a sound source. The sound pressure level versus frequency was obtained in the wind tunnel chamber and compared with the corresponding calibrated values. Fiberglas board-block units were installed on the chamber interior. The free field was increased significantly after this treatment and the chamber cut-off frequency was reduced to 160 Hz from the original designed 250 Hz. The flow field characteristics of the rotor-tunnel configuration were studied by using flow visualization techniques. The influence of open-jet shear layer on the sound transmission was studied by using an Aeolian tone as the sound source. A dynamometer system was designed to measure the steady and low harmonics of the rotor thrust. A theoretical Mach number scaling formula was developed to scale the rotational noise and blade slap noise data of model rotors to full scale helicopter rotors.

A theoretical study of the application of jet flap circulation control for reduction of rotor vibratory forces, addendum

Abstract: The theoretical potential of a jet flap control system for reducing the vertical and horizontal non-cancelling helicopter rotor blade root shears was investigated. It was determined that the dominant contributor to the rotor power requirements is the requirement to maintain moment trim as well as force trim. It was also found that the requirement to maintain moment trim does not entail a power penalty.

Helicopters safety drive

Abstract: A helicopter is provided with an emergency drive system adapted to rotate the helicopter rotor in the event of motor failure, the drive system being powered by compressed gas stored in the helicopter. The system may comprise nozzles aimed at the plane of rotation of the rotor, or a turbine coupled to the rotor drive shaft.

Laser Velocimeter Measurements of the Helicopter Rotor‐Induced Flow Field

Abstract: The use of a two-color laser velocimeter to measure the flow velocities in the wake of a helicopter rotor is discussed, including methods for obtaining two components of both instantaneous and time-averaged velocities. Results are presented from an experiment using a 2.13 m diameter model helicopter rotor operating at a tip speed ratio of 0.18 in a wind tunnel. The location of the tip vortex from the preceding blade was determined on the advancing side, and the diameter of the vortex core was found to be 15 percent of the blade chord (1.5 percent of the radius). The effects of the airfoil's bound vorticity were observed in the velocity distributions vry near the blade. These effects suggest that the laser velocimeter may be used to determine the aerodynamic loading (circulation) at a spanwise station on the blade. Also, the structure and boundary of the time-averaged wake were investigated.

Analog mixer to vary helicopter rotor phase angle in flight

Abstract: An analog mixer for a helicopter rotor control system mounted for tilting motion about a variable tilt axis and connected to the rotor control swashplate to cause synchronous tilting and including means to impart tilt generating control inputs to the analog mixer, and means to vary the position of the analog mixer tilt axis as a function of helicopter speed to thereby vary the phase angle of the helicopter rotor without imparting control inputs to the rotor control swashplate.

Combined helicopter flight controller

Abstract: A helicopter rotor control mechanism is provided in which cyclic and collive rotor control functions are performed by a single control wheel having three degrees of freedom and operable by either or both hands of the pilot as desired. The control wheel is rotatable and translatable on one axis and tiltable about another to perform lateral, collective and pitch controls, respectively, through non-interacting linkages driving conventional collective control, lateral control and pitch control push rods connected to a conventional rotor control head of the swash plate type.

Spherical bifilar tuning pin bushing

Abstract: A multibladed helicopter rotor carries at least three bifilar absorbers which are tuned to eliminate in-plane vibrations of the rotor which otherwise would be transmitted to the helicopter fuselage. Improved tuning pin assemblies and cooperating tracking inserts for the absorber reduce the undesirable frictional damping heretofore encountered in bifilar absorbers by centering the pendulous elements of the absorbers relative to their support members and reacting forces normal to centrifugal force by permitting flapping movement of the bifilar mass in the vertical plane.

100-kW hingeless metal wind turbine blade design, analysis and fabrication

Abstract: The design, fabrication and analysis of aluminum wind turbine rotor blades is discussed The blades are designed to meet criteria established for a 100-kilowatt wind turbine generator operating between 8 and 60-mile-per-hour speeds at 40 revolutions per minute The design wind speed is 18 miles per hour Two rotor blades are used on a new facility which includes a hingeless hub and its shaft, gearbox, generator and tower Experience shows that, for stopped rotors, safe wind speeds are strongly dependent on blade torsional and bending rigidities which the basic D spar structural blade design provides The 025-inch-thick nose skin is brake/bump formed to provide the basic 'D' spar structure for the tapered, twisted blades Adequate margins for flutter and divergence are predicted from the use of existing, correlated stopped rotor and helicopter rotor analysis programs

Laser velocimeter measurements of rotor blade loads and tip vortex rollup

Abstract: A method for obtaining and analyzing the instantaneous velocities of helicopter rotor flow fields through use of a laser velocimeter capable of simultaneously sensing two components of velocity is described. Rotor blade aerodynamic loads may be computed from the velocity distributions near the blades. The experiment was conducted with a 2.13 m (7 ft) diameter model helicopter rotor operating in a wind tunnel. Velocity distributions are presented which document the flow field near the advancing blade. Circulation is calculated from the velocity measurements, and the radial distribution of circulation is discussed. The influence of the tip vortex from the preceding blade is apparent in this distribution. Tip vortex rollup on the advancing blade was documented by making a series of measurements at various distances behind the blade. Effects of blade drag are evident in the velocities behind the blade trailing edge.

Digital computer program DF1758 fully coupled natural frequencies and mode shapes of a helicopter rotor blade

Abstract: The analytical techniques and computer program developed in the fully-coupled rotor vibration study are described. The rotor blade natural frequency and mode shape analysis was implemented in a digital computer program designated DF1758. The program computes collective, cyclic, and scissor modes for a single blade within a specified range of frequency for specified values of rotor RPM and collective angle. The analysis includes effects of blade twist, cg offset from reference axis, and shear center offset from reference axis. Coupled inplane, out-of-plane, and torsional vibrations are considered. Normalized displacements, shear forces and moments may be printed out and Calcomp plots of natural frequencies as a function of rotor RPM may be produced.