Abstract: A power transmission tool for the installation and removal of a spanner nut (15) on a threaded end of a drive shaft (27) comprises a torque reactor plate (32) with a central threaded aperture (35,37) for threaded engagement with a wrench shaft (39). The pitch of the wrench and reactor plate threads (37,39) is the same as the pitch of the drive shaft and spanner nut threads. The reactor plate (32) is fastened (60,63,65) to a drive flange (20), the flange being mounted on an end of the drive shaft and having splines to prevent relative movement therebetween. An aperture (73) is formed in the torque reactor plate for receiving a bar which is used to react the torque applied to the drive shaft/flange combination by the wrench via the nut, thereby preventing rotation of the flange and shaft combination when torque is applied to the wrench.

Abstract: A replaceable tip portion (20) for a helicopter rotor blade (14) is selectively swept, tapered and anhedral. The tip portion is formed of an upper and lower composite tip skin layer (35), each tip skin layer having bonded thereto a honeycomb core (38). The density of the core varies from a leading edge of the tip to a trailing edge of the tip from a high density to a lower density. A channel (recess) (40) is formed in the honeycomb for receiving a rotor spar tip end (30), and the tip portion is removably attached (60) to the rotor spar and the remainder of the rotor blade. The honeycomb core is attached to the upper and lower tip skin layers in halves, and is machined along the rotor blade chord plane (70) prior to assembly of the halves on the rotor spar, the machined core providing tight tolerance control in the construction of the blade tip portion.

Abstract: A helicopter engine speed reference (66) is increased (117,118) in response to heavy rotor loading (102) in excess of a threshold at the current descent rate (FIG. 4 ). The reference speed is faded up (117,118) at a rapid rate to 107% of rated speed (119). After a fixed time interval (129), reduce rotor loading (122), reduced pitch angle below a threshold magnitude (121) and reduced roll angle below a threshold magnitude (120) will cause the reference speed to be faded down slowly (146,147) to rated speed (148). Automatic increase in engine reference speed is overridden to 107% of rated speed (203) when a battle switch is activated (201) and weapons are ready (202). Increases in engine reference speed are prohibited (207) when the helicopter is operating in quiet mode (206) or the helicopter is resting on its wheels (208).

Abstract: A bleed air control system for an aircraft gas turbine engine operates in a fuel saving mode in response to operation of the aircraft in a steady-state cruise condition (215,350) within a threshold flight envelope (202,310) during steady-state engine operation above a threshold range. During operation in the fuel saving mode, a gas turbine engine compressor bleed valve (10) is closed. The rate of bleed valve closure is determined as a function of bleed valve position (12,20) and gas generator speed (26,30) at the instant the fuel savings mode is initiated.

Abstract: During operation of an integrated fire and flight control (IFFC) system in a coupled aiming mode, the occurrence of a weapon (172), weapon mount (145) or sensor (146) reaching a constraint limit (210) indices a fade-in of a constraint limiting axis command signal (140) for a respective attitude axis as the aircraft attitude reference, replacing the flight control system attitude feedback error signal (111) for the respective attitude axis. The constraint limiting axis command signals provide the aircraft attitude reference only while the pilot manually depresses and holds an enable switch (920). The pilot command stick input (90) remains the primary input to the IFFC system to thereby provide override capability for pilot command maneuvers.

Abstract: During operation of a flight control system in a coordinated area bombing mode, one pair in a group of pairs of fire control azimuth coordinate error signal and elevation coordinate error signals are faded-in as the aircraft yaw attitude reference and pitch attitude reference, respectively. Each pair is respectively indicative of the change in aircraft yaw attitude and pitch attitude for an aircraft reference axis to be aligned with an aiming line of sight. The aiming line of sight corresponds to a target location within a selected target area. A firing signal is provided in response to both the azimuth and elevation coordinate error signals being below respective threshold magnitudes, and the next pair in the group of pairs of azimuth and elevation coordinate error signals provide the aircraft attitude reference. A selected weapon is fired in response to the firing signal.

Abstract: A mode control (25) controls a helicopter flight control system (21) between a heading hold mode and a turn coordination mode according to switching patterns resembling the natural response of an expert pilot to flight conditions. Helicopter operating parameters (31,30) are detected (31,29) and each parameter's detected value is converted from a crisp value to a fuzzy input (100). A new mode fuzzy output is provided by applying a compositional rule of inference (105) across each fuzzy input and a composite mode selection rule base (110). The new mode fuzzy output is converted into a crisp value (112) which is used to determine whether the flight control system operates in the heading hold mode or the turn coordination mode.

Abstract: An inertial restraint mechanism restrains a rail mounted missile against forward movement along a missile launch rail in response to a deceleration force. The restraint mechanism comprises a mass the movement of which is opposed by a spring force. When the mass overcomes the spring force due to a deceleration force in excess of a threshold magnitude, a locking mechanism is activated to a missile restraining position, thereby restraining the missile against forward movement. The restraint mechanism is locked in the missile restraining position to prevent release of a missile in response to multiple deceleration forces experienced by the missile and launch rail. Additionally, a detent may be provided which prevents actuation of the restraint mechanism to the missile restraining position unless the missile is stored in a storage compartment.

Abstract: A rotor blade balance weight assembly comprises a weight retention housing (10) having a hollow core (22) and being closed on one end (12). The housing is disposed in an aperture which extends from the upper (52) to the lower (53) airfoil surfaces of the rotor blade (50). The closure end of the housing is contoured to match the airfoil upper surface contour. A stud (70) extends along the central axis of the housing, and washer-like balance weights (72) are received over the stud and secured by a locking nut (75). The open end of the housing is closed by a cover plate (25) that is contoured to match the airfoil lower surface contour, the cover plate being secured to the housing by a nut (78) in locking engagement with the stud. A pair of clips (37) are received within the housing, each clip having a tab surface (40) which extends through slots (42) in the housing, the clips being secured by the locking nut and stud.

Abstract: A helicopter engine fuel control anticipates changes in main rotor torque in response to lateral cyclic pitch commands, to thereby minimize engine and main rotor speed droop and overspeed during left and right roll maneuvers. A fuel compensation signal (100,101) is summed with a helicopter fuel control (52) fuel command signal (67) in response both to a lateral cyclic pitch command signal (LCP) (107) from a pilot operated cyclic pitch control exceeding a left or right threshold magnitude (201,210) and a total lateral cyclic pitch command signal (TCP) (108) from a lateral cyclic pitch control system exceeding a left or right threshold magnitude (202,207,215,220). The magnitude of the fuel compensation signal is dependant upon the direction of TCP and LCP, e.g., left or right, and helicopter roll acceleration (115). Alternatively, the magnitude and duration of the fuel compensation signal is dependant upon the rate of change in commanded lateral cyclic pitch (107,400,407).

Abstract: A helicopter engine fuel control anticipates sudden changes in engine power demand during yaw inputs to thereby minimize engine and main rotor speed droop and overspeed during yaw maneuvers. The rate (121,123) of yaw control (107) position change generates (110) a yaw component (104) of a helicopter fuel control (52) fuel command signal (70). The magnitude of the yaw component is also dependant upon the rate of yaw control position change (703). The fuel command signal yaw component (104) is overridden (113,115) when rotor decay rate (209,217) has been arrested during a sharp left hover turn (216); when the yaw component is removing fuel (239) during rotor droop (238); and when the yaw component is adding fuel (228) during rotor overspeed (227).

Abstract: During operation of a flight control system in a coupled aiming mode, wherein a fire control system (55) azimuth command and elevation command provide an aircraft attitude reference, a bank angle calculation function (1077) provides a bank angle signal to place the aircraft in a roll angle which results in a substantially coordinated turn. The bank angle signal is determined primarily as a function of an aiming line of sight heading rate for small azimuth commands, and is determined primarily as a function of aircraft heading rate for large azimuth commands. Additionally, the bank angle initially comprises a component as a function of aircraft lateral acceleration for driving aircraft lateral acceleration to zero, and after the aircraft assumes a roll attitude for turn coordination, the bank angle comprises a component as a function of aircraft side slip for driving aircraft side slip to zero.

Abstract: An automated test system for testing a filter pin connector (320) uses an insulation resistance tester (102) for testing the filter pin connector insulation resistance, a capacitance meter (104) for measuring the capacitance of each filter pin connector pin contact pair (10), a signal generator (100), and a radio frequency (RF) microvoltmeter for measuring the pin contact pair attenuation of test signals provided by the signal generator. A test fixture includes a pair of RF multiplexers (112, 113) and a high voltage multiplexer (115) for providing a signal path between each pin contact pair and the test equipment. A microprocessor (120) controls the test equipment and multiplexers for measuring the insulation resistance, capacitance, and signal attenuation of each pin contact pair.

Abstract: A memory seat control system provides repositioning of one memory seat to a recall position corresponding to another memory seat when the absolute location of both seats along corresponding travel axes is known. Additionally, the control system allows a recall position for one seat, e.g., a passenger's seat, to be entered from another seat, e.g., a driver's seat, when the absolute location of both seats along corresponding travel axes is known.

Abstract: A fail-operational control system utilizes low current switches for controlling vehicle high current loads. In a normal mode of operation, the control system utilizes a CPU to control the loads in response to input signals provided to the CPU by the low current switches. In a fail-operational mode of operation, during failure or removal of the CPU, certain high current loads are switched directly by corresponding low current switches to provide continued operation of the certain high current loads.

Abstract: A remote operating system for remote-controlled operation of a device provides secure communication of encoded messages between a transmitter (100) and receiver (150) of the system, and provides automatic re-synchronization of the transmitter and receiver without revealing a loss of synchronization to the operator. A pseudo random binary number (PRBN) generator (105) in the transmitter (100) produces a sequence of identification numbers. Each time the transmitter (100) is activated, the identification number contained in the transmitted encoded message is selected as the next number in the sequence of identification numbers. A PRBN generator (170) in the receiver (150) produces a sequence of reference numbers that is identical to the sequence of identification numbers. The receiver responds to a command code portion of the transmitted encoded message for operation the device when there is identity between the reference number and the identification number.

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: A vehicle mounted radar tracking system monitors vehicle heading rate, vehicle heading acceleration, vehicle roll rate and vehicle pitch rate, and provides a turn detect signal in response to heading rate, heading acceleration, roll rate or pitch rate being in excess of corresponding enable threshold magnitudes. In response to the turn detect signal, the tracking system dead reckons the positions of targets being tacked, and tracking system parameters are varied to make it more responsive to target motion. The turn detect signal is removed in response to heading rate, heading acceleration, roll rate and pitch rate being less than corresponding disable threshold magnitudes, and the tracking system is returned to normal operation.

Abstract: A vehicle mounted radar system has an antenna (12) which rotates about a reference axis (19). During angular movement of the antenna with respect to the reference axis, a radar indicated target azimuth (.phi.') is modified to provide a corrected target azimuth (.phi.) which is substantially indicative of the actual or true target azimuth. A tracking system (25) uses a predicted radar indicated target position (P.sub.n+1 (x,y)) for target tracking and correlation. The tracking system uses a smoothed target position (STP(x,y)), indicative of actual target position, for providing a visual display (39) of target track.