Abstract: The helicopter rotor blade connection has a support beam which connects the lift-generating wing section of the rotor blade with the rotor hub, and transmits centrifugal and lateral forces as well as bending and torsional moments. The connection is torsionally elastic and, at least in a main plane, bending elastic. A damping device for damping bending vibrations of the support beam is constructed alternately of elastomer layers and intermediate layers which are stiff in the direction of their extension. According to the invention, the damping device consists of a plurality of separate individual dampers which are arranged with their elastomer and intermediate layers in the direction of the main bending plane of the support beam.

Abstract: The present invention relates to a method and a device for reducing the vibration generated on the structure (14) of a helicopter (2) by the aerodynamic flow through the main rotor (3).According to the invention, said device (1) includes:at least one sensor (16, 17, 18) measuring the vibration generated on said structure (14); anda calculating means (19) determining, from said measurements, an alternating variation in pitch of the blades (6) of the tail rotor (5) of the helicopter (2), capable of generating an alternating force (T1) opposing said vibration, and transmitting the pitch variation thus determined to a blade (6) pitch control system (7).

Abstract: A control system for an aircraft that controls the set speed of the propeller governor and the degree of opening of the throttle valve using a single control lever. The speed setting mechanism of the propeller governor is directly connected to the power lever by a push and pull cable, and the set speed of the propeller governor changes in proportion to the stroke of the power lever. Therefore, the propeller speed changes in proportion to the stroke of the power lever. On the other hand, the throttle valve is connected to the power lever by a nonlinear cam and a push and pull cable. Therefore, the degree of opening of the throttle valve does not change in proportion to the stroke of the power lever (i.e., the change in the degree of opening of the throttle valve has a nonlinear characteristic), and the engine output power changes in the nonlinear manner with respect to the change in the propeller speed.

Abstract: A control system has an HHC actuator having a basal portion and a movable portion. A load acting on the control system is detected by a semiconductor gauge. The load detected by the semiconductor gauge is compared with a threshold value in a computer. If the detected load is larger than the threshold load, the movable portion is fixed to the basal portion by a fixing mechanism. By detecting the sign of stall of blades, the HHC actuator is fixed in a neutral position, allowing safe flight. A fixing switch is provided for allowing the fixing mechanism to operate regardless of the result of the comparison with the computer.

Abstract: A device for stabilizing yaw motion is provided for use on a helicopter having a main rotor rotatable about a main rotor axis, a power source, and a tail rotor boom with a longitudinal axis. The device includes a tail rotor and a gyroscopic mechanism. The tail rotor is mounted to the tail boom of a helicopter and is rotated about a tail rotor axis by the power source to generate a thrust force transverse to the tail boom and offset from the main rotor axis. The gyroscopic mechanism automatically varies the thrust force generated by the tail rotor to oppose yaw motions. The gyroscopic mechanism includes first and second variable pitch gyro arms extending along first and second gyro arm pitch axes, respectively, a support mechanism, and a pitch adjustor linkage.

Abstract: A wine turbine with blades having a high lift-to-drag ratio is provided. A hydraulic or other control system is used to control cone angle, such as in response to flap moment. Pitch of blades is controlled, such as in relation to the blade flap angle. Teeter motion is controlled preferably as a function of teeter angular velocity.

Abstract: A pitch link having an adjustable length is constructed as a diamond-shaped frame by coupling and fixing four flexible beams same in length and rigidity each at end portions thereof. Spherical bearings are provided on opposite ends with the longer diagonal line as an axis of the pitch link. A turnbuckle shaft is screwed into vertex portions on the shorter diagonal line, which is turned so as to adjust a length of the shorter diagonal line, thereby adjusting a length of the pitch link. The turnbuckle shaft is rotated by a detachable motor. Vibrations of a body and the rotational position of a rotor are detected by a vibrosensor and a phase detector respectively while an aircraft equipped with the pitch link is flying, an amount for adjusting the pitch link is calculated by a vibration analyzer from the electrical signals, and the motor is driven through a FM data transmitter/receiver to adjust the pitch link.

Abstract: A wind turbine has a rotor with at least one blade which has an aileron which is adjusted by an actuator. A hinge has two portions, one for mounting a stationary hinge arm to the blade, the other for coupling to the aileron actuator. Several types of hinges can be used, along with different actuators. The aileron is designed so that it has a constant chord with a number of identical sub-assemblies. The leading edge of the aileron has at least one curved portion so that the aileron does not vent over a certain range of angles, but vents if the position is outside the range. A cyclic actuator can be mounted to the aileron to adjust the position periodically. Generally, the aileron will be adjusted over a range related to the rotational position of the blade. A method for operating the cyclic assembly is also described.

Abstract: Method and device for the reduction of the oscillations of a divergent nature induced in the fuselage (2) of a helicopter (1) by its rotary wings (3) when they are turning. According to the invention, the oscillations are detected (at 17, 18) and they are converted (at 19, 20) into electrical signals, which are used (at 13) in order to produce pitch variations of the blades (4) of the rotary wings (3), in such a way that a rotating torque vector opposing said oscillations results.

Abstract: The blade angle controlling pitch beam servo (26) of a helicopter tail rotor (22) is responsive to a main rotor torque signal (92) indicative of torque coupled to a helicopter main rotor (10) for providing torque compensation so that the helicopter airframe will not counter-rotate under the main rotor. The torque coupled to the main rotor is that amount of engine torque (76, 114) in excess of torque coupled to the tail rotor (81, 115) and helicopter auxiliaries (84, 116). The amount of torque compensation provided by the tail rotor is reduced by an amount indicative of aerodynamic forces (130) on the helicopter airframe.

Abstract: Fault-tolerant apparatus for controlling first and second input members of a mechanical actuator includes first and second electrical machines and at least two controllers. The first and second electrical machines, which apply braking forces to the first and second input members, respectively, are provided with stators having pluralities of isolated phase windings. Each controller is coupled to a phase winding of the first electrical machine, and also to a phase winding of the second electrical machine. Any one controller can excite the first electrical machine to apply the braking force to the first input member or the second electrical machine to apply the braking force to the second input member.

Abstract: A control system for providing individual blade control inputs to a four-bladed helicopter rotor. Limited authority series actuators located in a non-rotating portion of the helicopter are driven by a computer which receives inputs from individual blade sensors and aircraft stabilization sensors, summed with cockpit control inputs, and the resultant motions transmitted to the rotor blades through a conventional swashplate which drives four blades of the rotor and a translatable differential sleeve and summing linkage which drives only two blades. The differential sleeve is adapted for axially raising or lowering the swashplate. The combined motion of swashplate and differential sleeve is shown mathematically to result in individual blade control in pitch. Computer individual blade commands are converted from the rotating reference frame to cyclic, collective and differential blade pitch commands in the fixed reference frame through a coordinate transformation matrix.

Abstract: Aircraft speed and direction are determined by sampling the sinusoid-like pressure variations at the end of a rotor and performing a Fourier analysis on the pressure samples. Fourier quadrature component signals are used to determine aircraft direction and are also used to determine, along with a steady state component, the aircraft's total velocity. A discrete Fourier analysis may be performed on the pressure samples using a plurality of subrevolution pressure samples averaged over each subrevolution interval.

Abstract: Quad-redundant signal processors in a flight control computer control valves to modulate airflow in a circulation control rotor system. The valves are arranged in a series and each valve is responsive to two signal processors. Loss of a valve occurs when both signal processors for that valve fail. A first valve, and every fourth valve in the series thereafter is responsive to a first signal processor and a third signal processor. The next valve in the series, and every fourth valve in the series thereafter is responsive to a second signal processor and a fourth signal processor. The third valve in the series, and every fourth valve in the series thereafter is responsive to the first signal processor and the fourth signal processor. The next valve in the series, and every fourth valve in the series thereafter is responsive to the second signal processor and the third signal processor.

Abstract: The pitches of the several blades in a helicopter rotor assembly are controlled by individual blade control subsystems. Those blade control subsystems respond to the output signals from accelerometers mounted on the blades. The signals from the accelerometers are used to control the pitch of the blades on which the accelerometers are mounted as well as the pitch of other blades. By appropriate orientation of any accelerometer and filtering within a particular feedback subsystem, each subsystem is designed to correct particular modes of blade motion to the exclusion of other modes of motion by varying lift.

Abstract: A windmill employs a separate wind vane pivotably mounted on the chassis of the windmill. An internal sensor detects the relative angle between the wind vane and the windmill and activates the pitch-control mechanism of the windmill to feather the windmill blades whenever the angle exceeds a predetermined maximum.

Abstract: Disclosed is an overload protection apparatus for variable pitch propellers, particularly cycloidal propellers. The apparatus is of the kind having a hydraulic servomotor (2) with an adjustment rod (1) for actuating a propeller, and having mechanical restoration means comprising an auxilliary adjustment means for adjusting a restoring lever (5) via an intermediate lever (6). As disclosed, the restoring lever (5) is pivotally linked to a first arm of the intermediate lever (6) which in turn is pivotally mounted on said adjustment rod (1) at a fulcrum (9). The first and second arms of the intermediate lever extend from opposite sides of the fulcrum. The second arm of the intermediate lever preferably includes a slideway (15) which receives a slide-shoe (16) affixed to an auxilliary adjustment lever (7), which is preferably an angular lever on which the auxilliary adjustment means acts at the end (12).

Abstract: A helicopter rotor having blade trailing edge tabs which are actuated in response to control rod loading to reduce or cancel the blade pitching moments imposed upon the blade and hence the control rods by blade aerodynamic and inertial loads, or other loads, encountered during rotor operation.

Abstract: Individual blade control of a helicopter rotor is accomplished by rotation of a mast (10) about a standpipe (12) where an individual blade (16) is coupled to the mast through a hub (18). Pilot input signals are transmitted by fly-by-optic transmission lines (14) through a fiber optic slip ring (40,42) to signal generators (134, 136, 138) for each of three sections of triplex actuators (76-79) coupled to individual rotor blades. Hydraulic fluid for positioning each of the actuators (76-79) is provided by electrohydraulic power packages (88, 90 and 92) where each package includes a pump and alternator. As each blade rotates about the mast (76) the pitch angle changes by individualizing the output of the signal generators (134-138) for each actuator.

Abstract: A helicopter mast is driven by a transmission with a swashplate individually linked to pitch horns on each rotor blade. At least four actuators are coupled to the swashplate, three of which normally control the attitude and position of the swashplate. A separate power unit is provided for each actuator, each power unit being independently driven from the transmission. Means are then provided to shift control of the swashplate from one of the three actuators to a fourth actuator upon any one becoming disabled.

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.

Abstract: A stabilizer bar has an aerodynamically shaped member, such as a winglet, on opposite ends thereof, and the bar is connected to rotate on its longitudinal axis and is mounted to tilt about a pivot axis at right angles thereto connecting it to a rotor shaft that also carries a pair of rotor blades the pitch of which can be changed about their longitudinal axis, that is parallel with the stabilizer bar pivot axis. A pair of oppositely directed mixing levers on opposite sides of the rotor shaft and connected parallel with the stabilizer bar for tilting movement therewith are connected by first control linkages to diametrically opposite sides of a wobble plate assembly, and by second control linkages to vary the pitch of the pair of rotor blades. An arm attached to the stabilizer bar in the plane of the aerodynamically shaped members has an end disposed in alignment with the pivot axis of the stabilizer bar that is connected by a control rod to the wobble plate assembly 90.degree.

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.

Abstract: A rotorcraft is subject to a rocking motion when on the ground, which is called ground resonance, and in the air which gives the fuselage a waddle type of motion about its center of gravity. both are a form of aeromechanical instability. This fuselage motion is coupled with and accompanied by bending oscillations of the blades with respect to the hub of the rotor of the rotorcraft. An oscillation damper is described which damps the oscillations of the aircraft and thereby eliminates or substantially reduces the rocking and waddle motion of the fuselage. The damper includes a seismic device having a mass-spring combination tuned to the frequency of the rocking motion which will typically be one half or approximately one half of the rotor revolutions per minute at cruising speed and is positioned laterally of and spaced from the center of gravity of the fuselage.

Abstract: In a helicopter, a pilot-actuated lever controls, through linkage mechanisms, a servo valve to drive a hydraulic piston; the piston moves a swash plate which in turn controls movement of rotor blade pitch positioning mechanisms against the force of blade loading, caused by aerodynamic forces. Blade loading, heretofore monitored visually by the pilot on a cruise guide indicator instrument, is used herein to control a secondary input to the servo valve, thereby to alter the position linkage mechanism which causes the force of a spring attached thereto to impose a force on the collective pitch control for blade loadings in excess of one-third of maximum allowable blade loads, in a direction tending to drive the control to lower collective pitch (lower blade loading). This provides "feel" to the pilot in proportion to blade loading when at critical magnitudes.

Abstract: In eliminating one-per-rev vibration in a helicopter due to rotor unbalance, a one-per-rev vibration signal manifestation of rotor unbalance is generated and sampled when the rotor blades are successively in alignment with a reference direction during each cycle of rotation of said rotor. A difference signal is generated from the samples to represent by magnitude and phase the character of the unbalance. The rotor is modified in magnitude proportional to the magnitude of the signal and in direction dependent upon the span and chord components of the signal. In one aspect the force track error signal is applied to the input of a cyclic pitch control actuator system dynamically to automatically correct span unbalance forces.