Abstract: A system for applying a predetermined proof load to a control cable and measuring the resultant cable length thereof includes a motor (32) which, under the control of a controller (14) drives an actuator (28) which, in turn, controllably displaces an end fixture (26). The exact distance moved by the end fixture (26) is monitored by an encoder (34) and a load cell (27) monitors actual loading on a cable under test. A home position sensor (30) provides input to the controller (14) to indicate whether or not the end fixture (26) is homed to its reference position. During cable testing, the fixed end of the cable is fixedly attached to a predetermined position in a channel (20). The free end of the cable is then attached to the end fixture (26). The controller (14) actuates the motor (32) to thereby provide a predetermined proof load profile to the cable under test. At the end of the proof loading test, the controller (14) determines, via the encoder (34), the distance L.sub.

Abstract: The present invention is directed to a non-normalized aircraft angle-of-attack indicating system. An angle-of-attack vane produces a signal .alpha.' indicative of raw aircraft angle-of-attack. A pitot-static system (68) produces output signals representative of aircraft static pressure and total pressure, P.sub.s and P.sub.t, respectively. The signals .alpha.', P.sub.s and P.sub.t are fed to an air data computer (66) which correspondingly produces an aircraft actual angle-of-attack value .alpha. and a calculated Mach number, M. Configuration sensors (80) produce output signals representative of gear position, g, flap position, f, and speed brake position, sb. The signals .alpha., M, g, f and sb are fed to a display computer (76). The display computer includes a submodule (90) which utilizes a lookup table to determine a value of angle-of-attack stick shaker, .alpha..sub.ss from the values of M, g, f and sb. A second submodule (92) utilizes a lookup table to produce a second angle-of-attack reference value .

Abstract: An improved aircraft flight management system includes a flight management computer (FMC), left and right control display units (CDU's) and a backup CDU. A triple redundant digital databus links the FMC and three CDU's. Upon detecting a failure in either the left or right CDU, the FMC utilizes reconfiguration rules stored in a look-up table to automatically cause the backup CDU to replace the failed CDU in operation, thereby relieving the flight crew of the burden of interfacing to the FMC through only one CDU. Upon detection of a failed databus, the system similarly utilizes structured data routing to reconfigure around the failed databuses. In addition, if any two CDU's have failed in a three CDU system, automatic data rerouting is implemented to the non-failed CDU to insure FMC to CDU communication.

Abstract: Signals from warning systems are passed to a hazard prioritization computer. The prioritization computer also receives inputs from the aircraft's air data and inertial reference system. The alert prioritization computer includes three functional modules: (1) hazard detection, identification and monitoring, (2) threat assessment and (3) display and alert prioritization logic. The hazard prioritization computer processes the warning system signals, along with stored data from a hazard database to compute a severity component of threat and a proximity component of threat. These two components are processed to produce an overall threat value for each hazard. This overall threat value is then processed to provide alert and display generation and prioritization for the flight crew and/or the aircraft's auto-flight system.

Abstract: Aircraft flight path changes commanded by a pilot via the wheel, column and pedal (12) are converted to position sensor signals (14) and passed as the input to primary flight computer (26). The primary flight computer (26) converts these pilot inputs and the inputs from autopilot (25) into desired surface actuator commands and then transmits them to actuator control electronics (18). The actuator control electronics (18) also receives a position feedback signal, representative of the position of the aircraft control surface (42). The actuator control electronics (18) produces a control signal which is fed to the input of an actuator (32) which includes a hydraulic system (34). The actuator responds to control input signals to drive linkage (40) which then positions the control surface (42).

Abstract: An aircraft automatic braking system processes the flight crew selected stopping position of the aircraft on the runway via a control display unit (30) and the aircraft's actual position, provided by a global positioning system (34), to generate a stop-to-position deceleration control signal in a provided control logic (36). If the flight crew selects the stop-to-position autobraking mode, the system determines whether or not a stop-to-position autobraking mode meets several predetermined criteria and, if the criteria are met, applies a control signal to the aircraft's braking system (62, 66) such that the aircraft is smoothly braked tending it to stop at the selected runway stopping position. The system eliminates the need for pilot lookup in a manual to determine a desired autobraking setting to choose based on altitude, temperature, approach speed and runway conditions and also operates to reduce pilot workload during limited visibility conditions.

Abstract: An improvement to an aircraft flight management system automatically allows changes entered into one control display unit (CDU) to be transmitted to the remaining CDU's upon failure of the onboard flight management computers. Each CDU includes memory allowing it to store the initial flight plan as being implemented by the flight management computer. Upon failure of all onboard flight management computers, detection of the failures by the CDU's initiates an alternate navigation mode. In this mode, entries made by a flight crew in one CDU are automatically transmitted to the remaining CDU's and, correspondingly, to the navigation displays, to thereby avoid redundant data entries in the remaining CDU's of a new, modified flight plan. In addition, all navigation displays are automatically caused to display the new, desired flight plan.

Abstract: A method for controlling the level of detail displayed in a computer generated screen display of a complex structure in which the structure is modeled as a root volume bounding all of the parts of the structure, with the individual parts comprising sub-objects bounded in sub-volumes. The method begins by selecting the root volume (202) and then enters the step of determining whether or not all sub objects within the selected volume are turned off (204). If they are, the entire display is culled. If they are not, the system then computes the screen coordinates of the selected volume (208). The method proceeds to the step of determining whether or not the selected volume is on the screen or off. If it is off the screen, the entire volume is culled. If it is on screen, however, a determination is made as to whether or not all sub-objects are turned on (214). If all sub-objects are turned on, the number of screen pixels for the selected volume is computed (216).

Abstract: An aircraft avionics system utilizes ARINC 429 compliant equipment, connected to one or more ARINC 429 data buses, and ARINC 629 equipment, coupled to one or more ARINC 629 data buses. Communication between the ARINC 429 and ARINC 629 equipment is facilitated by a data conversion gateway function (DCGF) system. The DCGF transforms ARINC 429 signals on one ARINC 429 bus for transmission over a second ARINC 429 bus, or over one or more ARINC 629 buses. The DCGF also transforms ARINC 629 signals on one ARINC 629 bus for transmission over a second ARINC 629 bus, or over one or more ARINC 429 buses.

Abstract: An improvement to an aircraft flight management system allows the flight crew to enter a predetermined position, which may be either the aircraft's current position or a defined waypoint, and a course line, or vector from the predetermined position. The system responds by displaying to the flight crew the predetermined position and course line, and then proceeds to link the course line to the existing flight plan. The flight crew may then either execute the modified flight plan or erase the modification and enter an alternate plan. The procedure of entering the predetermined position and course line, and linking to the existing flight plan may all be done with the autopilot activated, whereby appropriate lateral navigational signals are generated and supplied to the autopilot flight director to direct the aircraft along the newly specificed route. In addition, accurate performance data, such as distance to go, estimated time of arrival, and fuel remaining at each waypoint are displayed to the flight crew.

Abstract: An aircraft navigation control system includes a flight management system control display unit (12), a navigation display unit (40), logic circuitry (60) and a cursor control input device (70). The flight management system control display unit (12) includes a keyboard (30) to allow desired route waypoint modification via keyboard entries. These entries appear on a display (20) which provides a text listing of the selected waypoints. The navigation display (40) produces a graphical depiction of the selected waypoints as provided on the flight management system control display unit (12). In addition, the user may modify or create a new desired route by movement of a cursor (4) on the navigation display (40) to thereby highlight and select various waypoints as depicted on the display (40). By use of this graphical interface via a cursor control device (70) and the cursor (74) on the navigation display (40), a user may create, by a connect the dot approach, a new desired navigation route.

Abstract: A DC-to-DC switching power supply (12) receives a DC biased power signal at a first input (14), a control signal at its control input (16) and produces an output DC voltage at its output terminal (20). A power supply controller (30) compares the produced output signal V.sub.out with an input reference signal V.sub.ref. The reference signal is an analog signal which may be produced via digital-to-analog converter (34) which converts a digital input command into a corresponding analog signal V.sub.ref. The resulting error signal is then passed through a proportional gain block 36 and an integral gain block (38). A signal proportional to the differential of the output voltage is passed through a differential gain block (44). The outputs from the proportional gain block (36) integrator gain block (38) differential gain block (44) are summed in the second summer (40) and applied as the input to a delta-sigma converter (42).

Abstract: Disclosed is a method for processing an arbitrary collection of objects, forming a complex structure, into a hierarchy of bounding volumes, from a root volume bounding all objects, to sub-volumes bounding individual objects or assemblies thereof, for use as successive approximations to said objects in a computer generated display. The method includes the first step of creating a bounding volume for each of the objects. Selected bounding volumes are then processed through a combining algorithm determining whether or not, based upon a geometric relationship between the bounding volumes and the higher level, root volume, the selected bounding volumes can be combined. If it is determined that the bounding volumes can be combined, a new bounding volume is created with the combined volumes comprising sub-volumes thereof. This process systematically repeats and attempts to combine all sub-volumes.

Abstract: A method and apparatus for implementing a databus voter to select the command signals from one of several redundant asynchronous digital processing units finds particular application in the flight control art. Asynchronous, triple redundant primary flight computers (PFC's) (12, 14, 16) produce output command signals carried over a data buses (18) to control redundant actuators (32, 34, 36) in an aircraft flight control system. To prevent the transmission of a fault and a force fight among the redundant actuators, which drive a common aircraft control surface (50), a databus voter (52, 54, 56) associated with each PFC receives all PFC commands and, in accordance with a selection algorithm, outputs a voted one of the PFC command signals. A databus voter monitor, (144) associated with each databus voter, deactivates its associated databus voter if it detects a voted command signal which is inconsistent with the selection algorithm.

Abstract: A fail-operational/fail-operational fault tolerant clock includes a voting core comprised of triple redundant clock modules (30, 40, 50) and a floating hot spare module (60). Each module includes a voter (84) and fault detection, identification and reconfiguration circuitry (82) which operates to substitute the floating hot spare module produced clock signal for a failed voting core module signal without the introduction of transients or an asynchronous voted output. The modules (30, 40, 50, 60) are all preferably formed on a single semiconductor chip which includes isolation guard rings (32, 42, 52, 62) and independent power leads (34, 44, 54, 64) in addition to isolation buffering and point-to-point wiring to enhance fault tolerance.

Abstract: An improvement to an autopilot flight director system allows the flight crew to make a late runway change without disengaging the autopilot. The improved system allows a transition to a new runway assignment if radio altitude is greater than 1500 feet. Control logic determines whether or not localizer and/or glideslope have been captured and enters predetermined modes as a function thereof to effect a transition to the new approach while enabling instrument landing system retuning while maintaining autopilot engagement. Flight crew workload during approach is significantly reduced as a result of implementation of the improved system.

Abstract: Asynchronous, triple redundant primary flight computers (PFC's) (12, 14, 16) produce output command signals carried over a data bus (18) to control redundant actuators (32, 34, 36) in an aircraft flight control system. To prevent the transmission of a fault and a force fight among the redundant actuators, which drive a common aircraft control surface (50), a databus voter (52, 54, 56) associated with each PFC receives all PFC commands and, in accordance with a selection algorithm, outputs a voted one of the PFC command signals. A databus voter monitor, (144) associated with each databus voter, deactivates its associated databus voter if it detects a voted command signal which is inconsistent with the selection algorithm.

Abstract: An aircraft automatic throttle system is responsive to a throttle command signal to drive the aircraft throttle through a clutch and detent mechanism which allows for pilot override. Full throttle is achieved only if the throttle is driven against the stop position at the engine with a controlled, limited force. The instant invention is an improved servoamplifier system and method including a position sensor for sensing the position of the throttle, a summing circuit for summing the throttle command signal with the position sensor signal to produce an error signal and an amplifier for amplifying the error signal and producing an output signal to drive the servo driver. Unique signal processing is provided which monitors the amplifier output signal and detects a predetermined condition thereof representative of the throttle being driven against the forward stop at the engine.

Abstract: An onboard flight management system produces cost-effective trajectory target signals in response to an input desired time of arrival. Aircraft monitoring systems produce signals corresponding to real-time aircraft parameters including fuel flow rate, airspeed, windspeed, thrust, drag and mass. Stored in memory are aircraft characteristics including time cost, fuel cost, and acceleration due to gravity. A flight computer responds to an input desired time-of-arrival and the real-time parameters and stored characteristics signals and produces cost-effective trajectory target signals. The computer employs multiply regressed drag and fuel flow models which avoid trajectory discontinuities and the trajectory targets are produced using a Fibonacci search technique, which is computationally very efficient.

Abstract: An integral turbo jet/ram jet propulsion system has co-axial forward and rear portions and includes a converging supersonic compression chamber and an inlet in the forward portion for directing air into the supersonic compression chamber. A diverging subsonic compression chamber is disposed rearwardly of the supersonic compression chamber. An inlet sonic throat connects the supersonic compression chamber and the subsonic compression chamber and directs a flow of air from the supersonic compression chamber through the throat and into the subsonic compression chamber. A high by-pass turbo jet engine is disposed adjacent to the subsonic compression chamber for heating and expanding the air which passes into the engine. A subsonic expansion exhaust chamber and supersonic expansion exhaust chamber are disposed rearwardly of the engine with an exhaust sonic throat between the two.