Abstract: A pitch control apparatus for a variable pitch propeller for drive connection to an aircraft engine, wherein an optimum rotation number of the engine is calculated in relation to a flight speed of the aircraft, an opening degree of the engine throttle and the density of atmospheric air in flight of the aircraft to maximize a thurst force acting on the aircraft, and wherein a pitch angle of the propeller blade is controlled in such a manner that the rotation number of the engine becomes substantially equal to the optimum rotation number.

Abstract: A variable-speed propeller in a ship drive has the hydraulic valves feeding the working chambers of the hydraulic drive for the variable-speed propeller formed as electromagnetic valves controlled by timing relays which can be switched by a rocker switch or push-button switches to adjust the propeller pitch by an angular increment controlled by the "on" time of the relays. The engine speed is detected and, when the engine speed exceeds the maximum speed of the engine, the relay associated with increased pitch is triggered to obtain an increment of engine-speed reduction, thereby preventing the system from inadvertently having the blade pitch reduced when the engine is at maximum speed.

Abstract: The existence of autorotation of a load (e.g., a helicopter rotor system) coupled to the free turbine of a gas turbine engine is determined and a signal indicative thereof enables an integrator which, upon autorotation, integrates rate-based recovery anticipation provided thereto, the integrator output providing a time-composite bias signal to the main control portion of the engine fuel control to increase the speed of the gas generator of the engine in anticipation of rotor reengagement with the free turbine. The amount of recovery error is determined from the rotor speed error and is multiplied with the integrator output, the product being fed, upon completion of autorotation recovery, to the integrator input whose output now provides a controlled removal of the bias signal from the main control portion so as to minimize transients induced therein.

Abstract: A multiple blade aircraft propeller hub has a bulkhead supporting concentric slip rings connected by slide blocks to the aircraft electrical system. Electrical conducting flexible leads connect the slip rings to electrical heating elements or de-icing boots carried by the propeller blades. Each of the leads includes an elongated flexible braided copper tube for conducting electricity and which loosely surrounds a flexible multiple strand stainless steel cable, and an extruded flexible vinyl insulating tube loosely surrounds the braided tube. The leads provide for minimizing stresses and metal fatigue in the electrical conducting braided tube with changes in blade pitch and thereby provide for a substantially extended service life.

Abstract: The speed 54,56 of the free turbine 40 of a helicopter engine 22 is compared 134,138 with the speed 142,140 of the helicopter rotor 10 to indicate 106,108 a specific magnitude of autorotation and the deceleration 150 of the rotor above either one of two threshold magnitudes 220,222 (dependent on the magnitude of autorotation) is utilized 81,68,69 to increase fuel flow 72 to the engine according to a specific schedule 160,162 determined by the type of autorotation, in anticipation of rotor speed droop which would otherwise occur during recovery from the autorotation maneuver.

Abstract: The invention concerns a control for an aircraft propeller. During operation in reverse pitch, as occurs upon thrust reversal during a landing maneuver, the speed of the propeller is controlled by adjusting pitch of the propeller, in order to prevent overspeeding. Further, during the transition period while the propeller is converting from forward pitch to reverse pitch, the propeller presents a light load to the engine. At this time, fuel flow to the engine is limited in order to restrict delivery of energy to the propeller, again, for the purpose of preventing overspeeding.

Abstract: A propeller pitch controlling arrangement having a fuel economizer feature for use on marine vessels, has a mode selector valve operable to a normal mode and an economizer mode. In the economizer mode, a first pilot signal is transmitted to a path selector valve to establish a first flow path to a propeller pitch servomechanism. A second pilot signal is transmitted from a relay valve when the marine vessel speed drops below a predetermined value. The second pilot signal moves the path selector valve to a second position establishing a second flow path to the propeller pitch servomechanism. A first selected fluid pressure from a regulating valve arrangement mechanically linked to the engine fuel rack, is directed to the propeller pitch servomechanism when the first flow path is established. This first selected fluid pressure is proportional to the movement to the engine fuel rack.

Abstract: A propulsion aggregate for an aircraft consisting of a multicylinder-injection internal combustion engine and of a propeller with adjustable blade pitch driven either directly by the engine or by way of a speed reduction gear. The internal combustion engine is equipped with a commercially available continuously operating fuel injection system which for safety reasons and for designing for optimum aircraft operation, is additionally provided with additional control installations. A throttle valve in the suction line of the internal combustion engine, the blade pitch of the propeller and, in case the internal combustion engine is charged by a turbocharger, also the charging pressure are adjustable by a common adjusting mechanism.

Abstract: The speed (54, 56) of the free turbine (40) of a helicopter engine (20) is compared (103) with the speed (105, 106) of the helicopter rotor (10) to indicate (101, 102) autorotation, and the deceleration (108) of the rotor above a threshold magnitude (110) is utilized (81, 68, 69) to increase fuel flow (72) to the engine in anticipation of rotor speed droop which would otherwise occur during recovery from the autorotation maneuver.

Abstract: Damping of the primary bending mode of a tower (12) mounting a wind turbine having a control (36) for providing a pitch blade angle reference signal (40) to modulate the pitch of the turbine blades (1) through a pitch change mechanism (38) for constant power is provided by generating the pitch blade angle reference signal as the integral (104) of the summation (266) of a torque/power controlling blade pitch angle reference rate signal (98) with an estimated acceleration signal (255) generated by filtering (250, 252, 254) the blade pitch angle reference signal (40) with the following transfer function ##EQU1##

Abstract: An accelerometer (1) disposed on the support tower (12) of a wind turbine electric generating system in the vicinity of the rotor (10, 16) thereof provides a signal (1) indicative of acceleration of the tower in the direction of the rotor rotational axis. The signal (2) is passed through a band-pass filter (4) for summation (9) with a torque/power controlled (100) blade pitch angle reference rate signal (98), the integral (104) of which provides a blade pitch angle reference signal (40) to control the pitch angle of the rotor blades (10) through a pitch change mechanism (38), thereby to provide additional, positive aerodynamic damping to the tower while modulating blade angle for constant torque/power in response to wind turbulence.

Abstract: The invention relates to improved control systems for controllable pitch propellers of a twin screw vessel driven by a single prime mover operating at a predetermined rotational speed while the vessel is in the at sea mode, with provisions for maintaining the total load of said propellers upon the prime mover substantially constant under varying at sea conditions. A further element of the invention permits the transition of the control system to and from the maneuvering mode to the at sea mode under substantially smooth operating conditions of speed and power.

Abstract: A torque control system designed to economize on consumed fuel includes a source of motive power, a load and a transmission coupling the motive power to the load. A torque measuring device is attached to the output of the transmission and senses a torque applied to the load. The output of the torque sensor is compared with electrical analogs of optimized torque situations. The comparison results in the development of a control voltage which is applied to the source of motive power and the transmission to effect optimization of torque control.

Abstract: A control system for a helicopter power plant with three power units driving the lifting rotor system through a main gearbox. Each power unit comprises a gas turbine engine of the gas-coupled type, a hydromechanical fuel control receiving electrical inputs to set a gas generator governor in the fuel control, an engine electronic control, and various engine accessories.

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.

Abstract: A control system for a helicopter power plant with three power units driving the lifting rotor system through a main gearbox. Each power unit comprises a gas turbine engine of the gas-coupled type, a hydromechanical fuel control receiving electrical inputs to set a gas generator governor in the fuel control, an engine electronic control, and various engine accessories.The operation of the power plant is normally controlled by the pilot through condition levers and switches on a control quadrant, through a pitch control by which the collective pitch of the rotor blades is controlled, and through a master beeper switch which is operable to trim the setting of an isochronous governor for the rotor system and the gas generator governors. The condition lever of each engine transmits a speed command signal through the electronic control of each power unit to its gas generator governor.

Abstract: A coordinated fan pitch, fuel flow and fan exhaust nozzle area control for a gas turbine powered aircraft propulsor of the bypass duct, variable pitch and exhaust nozzle configuration serves to effectuate reverse pitch change through feather, by minimizing forward thrust excursion and shaft torque to obtain rapid reverse thrust response.

Abstract: A control for a variable pitch fan propulsor driven by a turbine type of power plant which fan is mounted in an engine bypass duct having a variable exit nozzle. The control serves to coordinate the control of fuel flow to the engine, the area of the fan exit nozzle, and the pitch of the fan blades by biasing the power lever position signal with Flight Mach No. An additional feature is the inclusion of fan surge control derived from signals of flight Mach No. and corrected free turbine speed.