Abstract: The output of a bridge measurement circuit is sampled at a time when there is no desired, or expected, bridge output signal, and a balance voltage applied to the bridge measurement circuit is automatically adjusted such that the sample bridge output signal is equivalent to a reference output signal. After the bridge is balanced, the balance voltage is automatically modified by a calibration voltage calculated to provide a known change in the sample bridge output signal, and the sample bridge output signal is compared to a reference output to automatically check the bridge measurement circuit gain. The balance voltage may be automatically modified by multiple calibration voltages, and the sample bridge output signal provides an indication of the linearity of the amplifier gain. The sample bridge output signal is monitored after a change in the balance voltage to automatically provide an indication of the bridge measurement circuit filter cutoff frequency.

Abstract: A helicopter external cargo suspension system (12) has load sensors (19,20) at each suspension point for sensing the magnitude of a load (32) at each suspension point. A delay period (FIG. 6) is determined based on the magnitude of the load, and a load is released if one of the following failure modes occurs for a period which exceed the delay period: one of the suspension points is unloaded, the magnitude of the load exceeds a maximum threshold magnitude, or if the distribution of the load, as indicated by the difference in magnitude at each suspension point, exceeds a difference threshold magnitude. If the magnitude of the load is below a minimum threshold magnitude, the load will be carried by one of the suspension points in response to a failure of the other suspension points.

Abstract: An integrated fire and flight control (IFFC) system determines a ballistic firing solution based on the position of targets relative to a helicopter and also based on the type of weapons to be fired. An elevation command is determined based on the required change in helicopter attitude to achieve the ballistic firing solution that, combined with the estimated time required to perform the aim and release of weapons, provides an estimate of deceleration and velocity loss that will occur. A forward acceleration and velocity profile is determined based on the desire to make a symmetrical maneuver sequence involving a nose down acceleration to achieve the acceleration and velocity profile that will be canceled by the subsequent deceleration and velocity loss during the pitch up maneuver to the ballistic firing solution.

Abstract: A mission computer (10) stores a nominal route plan which includes information indicative of the precise arrival time at specific navigation positions, e.g., waypoints, and the planned fuel consumption during travel to the specified waypoints. The mission computer (10) provides a time and fuel display system (15) with the route plan information. The time and fuel display system (15) generates a display (200) which indicates the total distance to be traveled during the mission (202), a distance traveled indicator (212) to reflect the progress of the aircraft along the mission, and a time bar (205) and a fuel bar (207) the height of which with respect to the indicator (212) are indicative of the percentage that the current time and fuel values are over or under the planned values for the current waypoint. The distance traveled indicator (212) is labeled with a waypoint designation (215) which is indicative of the designation of the waypoint towards which the aircraft is currently flying.

Abstract: A solid state power controller is automatically reset if an overload trip occurs during a start-up delay period. If an overload trip occurs during load starting, the SSPC remains tripped for a reset delay period, and thereafter the SSPC is automatically reset, thereby allowing the starting process to continue by heating a lamp filament, accelerating a motor, etc., without indicating a system fault to an operator. At the expiration of the start-up delay period, the SSPC no longer automatically resets in response to an overload condition.

Abstract: RF energy received by an antenna (10) of a radio direction finding system is evaluated for determining the presence of a RF signal transmission source at bearing values at which the received RF energy is statistically strong (33) and which have a relatively stable signal strength (35). The scanning pattern (sequence) ( 12, 15) of the antennae is altered to increase the dwell time of the antennae at bearings having a statistically strong and relatively stable signal strength combination, and decrease the percentage of time spent scanning the remainder of the 360 degree search spectrum. External sensors (30) are used to sense changes in antennae orientation with respect to a fixed reference to thereby reduce bearing errors (27) during changes in the attitude and heading of the platform carrying the radio direction finding system.

Abstract: A fire control system (55) azimuth command and elevation command provides a flight control system attitude reference in response to operation of the flight control system in a coupled aiming mode. The coupled aiming mode is engaged in response to the continuous operation of a pilot switch (920) the azimuth command and elevation command being below respective threshold magnitudes (940, 941), and the operation of the fire control system (55). During operation in the coupled aiming mode, the azimuth command and elevation command replace the yaw attitude feedback error signal and pitch attitude error signal, respectively, as the aircraft attitude reference.

Abstract: An automatic engine speed hold control system (430) allows an engine (420) to quickly achieve a commanded engine speed and thereafter maintain a constant commanded engine speed regardless of changes in engine loading. An engine speed error signal (444) is computed as the difference between an operator commanded engine speed (440) and actual engine speed (434,436). The engine speed error signal (444) is scaled based on an engine speed error gain (448), and the scaled engine speed error (452) is then provided via an integral path (456) and a proportional path (454) to a summing junction (478) where it is summed with load anticipation trim signals (474,467) to provide a final trim signal (485).

Abstract: In response to the loss of a generator in an electrical power system, all operating electrical loads are automatically switched to a second generator, and the electric power system continues to provide power to all of the loads. The total power required to operate all operating loads is permitted to exceed the nominal rated capacity of the operating generator, and the winding temperature or other critical hotspot temperature of the operating generator is monitored to determine whether thermal stress limits will be exceeded. If the critical hotspot temperature approaches the thermal stress limit, certain electrical loads are disconnected via a pre-programmed load shed priority schedule, and thereafter loads are re-connected as other loads are de-energized or if the critical hotspot temperature decreases.

Abstract: A fairing 122 is provided for covering a wire strike cutter 120. The fairing 122 comprises a microwave energy absorbing core 150 covered by dielectric skin 160,161,165,166. The fairing 122 reduces the radar signature of the wire cutter and fractures in response to a wire strike so as to permit the cutter to capture and cut a wire/cable. At least one projection 135 is formed on the fairing 122 for catching a slack wire/cable as it slides up the fairing, the projection holds the wire/cable until tension develops to fracture the fairing, thereby permitting the wire/cable to pass through to the wire cutter where it is subsequently cut. The fairing 122 minimizes drag and improves the aircraft's aerodynamic characteristics by directing the airflow around the wire cutter and equipment adjacent to the wire cutter.

Abstract: In a dual station sidearm control system, a pilot control station provides pilot control signals in response to pilot activation of a pilot sidearm controller (210) and a co-pilot station provides co-pilot control signals in response to co-pilot activation of a co-pilot sidearm controller (211). The pilot control signals and the co-pilot control signals are combined (220) to provide dual station control signals for controlling aircraft attitude in the yaw, pitch, roll and lift attitude axes. A priority detector function (222) is responsive to pilot control signals in one of the attitude axes for reducing the priority of co-pilot control signals with respect to pilot control signals in the one attitude axis as the magnitude of pilot control signals increases in the one attitude axis.

Abstract: A target tracking system (10) comprises sensors (12) which provide data corresponding to a region of interest, the data being time dependent and consisting of amplitudes, ranges and angles. A window (202) is placed around data of interest, the size of the window being determined based on target size, assumed speed and acceleration characteristic, and the window is thereafter broken down into a plurality of smaller windows (208), thereby forming a grid having a nodal point (210) at each corner of the smaller windows. The data within the window is stored in a matrix, and background noise is thereafter minimized by filtering the data past a threshold value (215). The filtered data is analyzed to determine its distance weighted contribution at each nodal point (219), and the weighted distances are summed for each nodal point resulting in a nodal point magnitude for each of the nodal points (220).

Abstract: An antenna test and measurement system (100) performs measurements of the transmitted or received signal strength of a vehicle mounted transmit or receive antenna (102) with respect to all aspects of the system with respect to the vehicle mounted antenna, and vehicle state parameters (131, 139), i.e., position and attitude, are simultaneously recorded (185, 186). The antenna measurements are correlated with the state parameter measurements such that variations in the measured signal strength caused by changes in vehicle state are apparent. A graphical display (189, 190) of antenna signal strength and vehicle state parameters are presented for the entire 360.degree. azimuth of an antenna test so that variations in the signal strength caused by changes in vehicle state parameters may be identified.

Abstract: A fiber optic cable harness break-out fitting (10) is provided with a pair of support legs (12, 14) for supporting contact with a fiber optic cable harness (42) having a plurality of fiber optic cables (45). A take-off tube (20) is positioned between the support legs (12, 14), and the fitting is positioned on the harness such that fiber optic cables (47) which break-out from the fiber optic cable harness are received through a throat (22) of the take-off tube (20). Support surfaces (37,39) are provided between each of the legs (12,14) and the take-off tube (20), the surfaces having a controlled radius to control the degree of bend in the break-out fibers (47). The legs (12,14) are mounted on the harness (42) using tie down material (53), and the end of the take-off tube is slotted (24) such that the break-out fibers exiting the tube are secured within the tube with tie down material (53), the slots providing the necessary flexibility to allow the take-off tube end to securely grip the break-out fibers.

Abstract: An engine failure monitor 50 for use with a multi-engine aircraft having at least two engines detects conditions indicative of a partial (700) or total (400) engine failure, including turbine shaft failures (600). In response to an engine failure, suitable inputs (134, 138, 712) are provided to an electronic engine control for operating the remaining engine. Additionally, indications indicative of the engine failure (132, 710) are provided to the cockpit.

Abstract: A control system for automatically loading and unloading a vehicle (10) from a cargo compartment comprises a plurality of distance measuring sensors (31,32,33,34) attached to specified locations on the exterior surfaces of the vehicle (54,56,57) and the cargo compartment (55). The sensors provide distance signals, indicative of the distance between each sensor and the nearest obstruction, to a microprocessor which compares each of the distance signals to reference signals indicative of the minimum allowable clearance between each specified location and an obstruction. During activation of an automatic loading/unloading switch (70), a winch (39) is activated to move the vehicle into or out of a cargo compartment if each of the distances is greater than the corresponding minimum allowable clearance.

Abstract: An aircraft landing locking pin (12) automatically repositions from an engage position, wherein a rotating portion (15) of the landing gear is prevented from rotating, to a released position, wherein the rotating portion if free to rotate, in response to excessive side loading on the landing gear. The locking pin comprises a spring loaded actuator (32,34,60) which urges a roller (19) of a lock pin assembly (20) in contact with a bushing (22). In response to a side loading on the landing gear in excess of a predetermined loading, the locking pin is automatically released wherein the roller is expelled from the bushing. The threshold loading is determined as a function of the spring force applied to the lock pin assembly via the actuator linkage. The locking pin automatically reengages when the lock pin assembly is aligned with the bushing.

Abstract: A fire control system (55) azimuth command and elevation command provides a flight control system attitude reference in response to operation of the flight control system in a coupled aiming mode. The coupled aiming mode is engaged in response to the continuous operation of a pilot switch (920) the azimuth command and elevation command being below respective threshold magnitudes (940, 941), and the operation of the fire control system (55). During operation in the coupled aiming mode, the azimuth command and elevation command replace the yaw attitude feedback error signal and pitch attitude error signal, respectively, as the aircraft attitude reference.

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.