Abstract: An apparatus and method for inserting a waypoint into a preexisting flight plan which includes selecting a waypoint on a graphical display of a portion of the flight plan and automatically generating a proposed changed flight plan based upon inserting the waypoint into the nearest leg of the flight plan.

Abstract: A method and computer program product are provided for controlling the actuators of an aerodynamic vehicle to affect a desired change in the time rate of change of the system state vector. The method initially determines the differences between anticipated changes in the states of the aerodynamic vehicle based upon the current condition of each actuator, and desired state changes. The differences between the anticipated and desired state changes may be weighted based upon a predetermined criteria, such as the importance of the respective states and/or the weight to be attributed to outliers. The differences between the anticipated and desired state changes are converted to the corresponding rates of change of the actuators. These changes in the actuators may be limited to within predefined bounds. Control signals are then issued to the actuators to affect the desired change in the time rate of change of the system state vector.

Abstract: Process and device for controlling at least one aerodynamic elevator surface of an aircraft during a takeoff. The device (1) comprises a control member (3), calculation means (5) and motor means (7) for controlling the aerodynamic elevator surface (2), as a function of an attitude angle control law received from the calculation means (5). If the duration T between the moment at which the pilot actuates the control member (3) so as to control the aerodynamic elevator surface (2) in such a way as to increase the pitch attitude angle of the aircraft and the moment at which the aircraft should reach the minimum takeoff speed is less than or equal to a nominal duration, the calculation means (5) determine and transmit a minimum control law. Otherwise, they determine and transmit a modified control law which is such that at a duration T after the actuation of the control member (3), the pitch attitude angle is substantially equal to a nominal value.

Abstract: An electrical fly-by-wire system for operating an aircraft rudder includes a low-pass filter, arranged between a rudder bar and an actuator of a rudder. The low-pass filter receives a control command from the rudder bar corresponding to the degree of travel the rudder bar has experienced from a neutral position. Based on the amplitude of the control command, the filter generates an operating command for the actuator. Additionally, the filter operates such that the higher the fraction of the rudder bar's travel away from the neutral position, with respect to its maximum value of travel, the higher the filter's time constant is set.

Abstract: Process and device for controlling at least one aerodynamic elevator surface of an aircraft during a takeoff.

Abstract: A navigation and guidance computer system using navigation sensors to determine aircraft position, and automatically control the aircraft to fly from an initial position to intercept and follow a crew-selected rendezvous orbit in accordance with Air Traffic Control and military specified rules. It creates maneuvers that result in smooth, efficient motion of the aircraft from an initial position to the execution of a rendezvous orbit. It controls the aircraft by applying commands to the aircraft's autopilot. The system creates the smooth and coordinated maneuvers by considering and anticipating all constraints on the aircraft flight path by creating a single maneuver that is a re-entry maneuver, an over fly maneuver, and a turning maneuver all as a single turn that is contained entirely within the allowable airspace. All maneuvers are created and executed recognizing the turning capability of the aircraft considering allowable bank angles, current wind conditions and true airspeed.

Abstract: A flight plan is modeled by a trajectory with a set of segments having finite lateral and altitude extents between nominal way points. A terrain database stores hierarchical patches with maximum and minimum altitudes and pointers or altitudes for subpatches. Linear-programming inequality constraints match segments with patches. A search finds any terrain locations that impinge upon the trajectory tube, indicating an error condition for the plan. Moving hazards are modeled with segmented trajectories. Conflict with a moving hazard occurs when moving bubbles within both of the trajectories overlap in both space and time.

Abstract: An aircraft display and control system generally includes a processor, a cursor control and selection device, an aeronautical information database, a geographic database, and a plurality of display devices. Users, such as an aircraft pilot and copilot, can perform flight plan entry and modification by manipulating graphical information on the display devices using cursor control. In one embodiment, the present invention allows multiple members of an aircraft crew to share control of common flight information display areas, aids the crew's situational awareness by providing software-implemented dynamic symbology and highlighting to indicate cursor location, current panel of entry, and current focus for keyboard and cursor events.

Abstract: Apparatus for estimating a first height (1) of a vehicle (2) above a first reference surface (3), including a system (4) for determining position, velocity and attitude incorporating at least one sensing means (4b) operable to provide an output signal (5) indicative of a vertical specific force of the vehicle, error-estimating means (6) for receiving as input signals a horizontal reference velocity and position (4a) of the vehicle (2), and a radar altimeter measurement (7) of a second height (1b) of the vehicle (2) above a second reference surface (8) and for providing as an output signal (9) estimates of errors associated with the sensing means output signal (5), and integrating means (10) for receiving said sensing means and error-estimating means output signals (5, 9), and for subtracting the estimates of errors from the signal (5) indicative of vertical specific force while performing a double integration of the results of the subtraction, to provide an output indicative of the required estimated first he

Abstract: A system and method for intervention control of an aircraft in the event of pilot command error whether voluntary or involuntary. Impending detection of a chaotic condition associated with a maneuvering aircraft enable early prediction and control of the aircraft where solutions based upon performance prediction are available. A further feature of the present intervention control of the aircraft enables an equipment malfunction detection signal substitution of a satisfactory equipment signal.

Abstract: Electrical fly-by-wire system for operating an aircraft rudder.

Abstract: An apparatus, methods, and computer program products are provided for monitoring the attitude of an aircraft. The apparatus of the present invention includes a navigation system, such as a GPS, that provide values representing the velocity vector of the aircraft. Additionally, the apparatus includes two gyroscopes positioned with respect to the aircraft to sense the roll and pitch of the aircraft. A generator is connected to the navigation system and generates a calculated flight path and roll angle based on the velocity vector of the aircraft. Further, the apparatus of the present invention includes a combiner connected to both the generator and the gyroscopes. The combiner combines the calculated flight path angle to the sensed pitch angle and the calculated roll angle to the sensed roll angle and generates composite flight path and roll angles. The composite flight path and roll angles represent the attitude of the aircraft.

Abstract: Aircraft with electrical fly-by-wire controls, equipped with an automatic pilot.

Abstract: A method, apparatus and computer program product for assisting the pilot of an aircraft in making a go-around decision. Various aircraft parameters are monitored during the approach to land and a risk level assessed. When the risk level exceeds a specified threshold, an alert is provided.

Abstract: An improved aircraft flight management system (FMS) based on a layered subsystem architecture, residing on a computing platform and including an operator interface subsystem, a communications subsystem, a flight management subsystem, and a database management subsystem, wherein the architecture is predicated on the enforcement of subsystem-dependency rules wherein a given subsystem is allowed to depend only upon another subsystem in the same or lower hierarchical layer.

Abstract: An avionics system for retrofitting a GPS system to work with an existing autopilot by using signal switching and conditioning in an electronic flight display to accomplish the switching without the need for logic controlled relays in the GPS to autopilot interconnect wiring.

Abstract: A flight control system for an aircraft, particularly for a helicopter receives a number of items of information, especially the position of flight control members of said aircraft, and sends, depending on these items of information, control inputs to controls of the aircraft. The flight control system includes at least one sensor which measures values representative of the flight of the aircraft and computing means which generate, at least from these measured values and from the positions of flight control members, corrected control inputs which are sent to the controls and which enable control of the lateral speed of the aircraft with respect to the ground, without varying the course of the aircraft.

Abstract: A high integrity navigation system is created using a low integrity navigation computer monitored by a high integrity system. A navigation sensor (e.g. DGPS, GPS, ILS, MLS, and the like) is programmed with a desired trajectory. The navigation sensor calculates the vehicle position and generates a deviation signal indicative of the deviation of the vehicle from the desired trajectory. The navigation computer uses the deviation signal in the control law computations to generate steering commands for controlling the vehicle. The navigation sensor monitors the deviation of the vehicle from the desired trajectory. If the deviation exceeds a predetermined threshold, an integrity alarm signal is communicated to a display or other alarm to alert the operator.

Abstract: In a method of flight control in which a thrust command is computed based on the total aircraft energy error relative to flight path and speed control commands, and an elevator command is computed based on the energy distribution error relative to the same flight path and speed control commands, an improvement is provided including an elevator control command in response to a column control input by the pilot. In the short term, the computing establishes a change in flight path angle beyond the sustainable flight path angle at the trim speed for the prevailing thrust condition. In the long term, the computing establishes a change in speed relative to a set reference speed, the speed change being proportional to the column control input. In the long term, the computing establishes a flight path angle equal to a sustainable value for the prevailing thrust condition and the altered speed condition.

Abstract: An individual guidance system for aircraft in an approach control area under automatic dependent surveillance which, by supplying the pilot with the required flight data automatically in units of micro air spaces, permits safe and accurate flight with little scope for human error, being an individual guidance system for aircraft in an approach control area under automatic dependent surveillance wherein the air-traffic control system divides the approach control area automatically into a group of micro air spaces, and establishes flight rules within the micro air spaces in order to guide aircraft by establishing no-fly air spaces.

Abstract: The airspeed limiting performance envelope of a helicopter is converted to a groundspeed envelope by factoring-in the wind speed and direction. A groundspeed command for an unmanned helicopter is provided as a function of the range to the helicopter's destination. The function may be calculated so as to provide a nominal groundspeed command equal to a nominal groundspeed command limit, or the function may be a fixed function to cause the aircraft to creep toward final approach. The cosine and sine of the desired relative flight direction of the aircraft, which is equal to the true bearing to the destination minus the true heading of the aircraft, are utilized to scale the groundspeed command into a longitudinal groundspeed command and a lateral groundspeed command. If one of the groundspeed commands is over a corresponding limit, the other groundspeed command is scaled back so that the vector addition of the two commands will cause flight in the desired direction.

Abstract: A system for protecting an aircraft in cruising flight against excessive load factors when a vertical gust of wind occurs is disclosed. The system has a pitch-attitude flight control system (21) which is under the control of a pilot of the aircraft, an automatic pilot (26), a switch (25) allowing the aerodynamic pitch-attitude control surfaces (22) of the aircraft to be controlled either by the flight control system (21) or by the automatic pilot (26), and a protection system (31) capable of acting on the switch (25) to disconnect the automatic pilot (26) and switch on the flight control system (21). The disconnection of the automatic pilot (26) and the switching-on of the flight control system (21) is delayed by a delay time which is at least approximately equal to a typical duration of a vertical gust of wind, and the switching-on of the flight control system (21) is not carried out unless the gust of wind persists when the delay time has elapsed.

Abstract: A stand alone terrain conflict detector of an aircraft includes a global positioning system (GPS) receiver, an inertial navigation system, navigational and topographical databases, a control panel, a central processing unit (CPU), which CPU generates position data, a current flight path vector and control signals, an obstacle detector which receives the position data and the current flight path vector and which generates a flight path signal, an alert signal identifying a terrain threat to the aircraft and a projected flight path vector, a video generator coupled to the obstacle detector and the CPU, and a display connected to the video generator. The display outputs one of a 2D image, a first 3D image and a second 3D image and the terrain threat generated by video generator.

Abstract: A system and method for determining the angular orientation of a body moving in object space. The invention is usable to detect and identify the angular orientation of the body relative to the theoretical plane of gravity and thereby relative to any other theoretical plane between zero and three-hundred sixty degrees therefrom. The invention is operable regardless of whether the body is rotating with an angular velocity, either natural and/or induced. The invention may employ one or more analog and/or digital sensors. The analog sensor may be an accelerometer. The system may be used to provide real-time orientation data to permit the body to be guided or directed towards a desired destination or simply to permit the orientation of the body to be known, controlled or varied as desired.

Abstract: A monitoring and control system for a dual-engine helicopter for OEI flight operations includes a parametric indicator operative in response to sensor signals for monitoring an engine gas generator speed parameter N1 during dual-engine and OEI flight operations. The indicator includes a rotatable needle that provides an analog indication of the current value of the N1 parameter. For dual-engine flight operations, the indicator includes a first DE indicia defining a normal operating range for the N1 parameter, a second DE indicia that defines a precautionary operating range for the N1 parameter, and a DE indicium that defines a take off power limit for the N1 parameter.

Abstract: A mobile craft is guided with respect to a trajectory. To guide the craft, the distance xtk is determined between the mobile craft and the point of orthogonal projection of the mobile craft on a segment of the trajectory. The angle tkae is determined between the direction of travel of the mobile craft and the direction at the point of orthogonal projection. A control value is computed for the lateral attitude from the distance and angle, two characteristic coefficients and a nominal control value. This control value is applied to the craft.

Abstract: A system and method for conducting one engine inoperative (OEI) flight procedures training in a dual-engine helicopter includes a multi-function OEI training switch that is operative to initiate OEI flight procedures training by selecting one of said engines as the single operative engine for OEI flight procedures training, and a training function module that is: (1) operative to establish suppressed 30-second, 2-minute, and maximum continuous OEI operating limits for selected engine operating parameters to limit the actual power provided by said powerplant system during OEI flight procedures training; and (2) operative to generate biasing factors to control the operation of the parametric indicators for the selected engine operating parameters during OEI flight procedures training.

Abstract: An aircraft overspeed protection system produces proportional and integral commands for input to an autopilot controlling the aircraft. The proportional and integral commands are produced by comparing actual monitored speed of the aircraft with a target speed from a target speed selector. The target speed selector selects as the target speed a trigger speed above a nominal maximum operating airspeed of the aircraft until the trigger speed is reached by the aircraft. When the trigger speed is reached by the aircraft, the overspeed control goes into Overspeed Protect Command active mode and the target speed selector selects a new target speed below the nominal maximum operating speed of the aircraft. The overspeed protection system remains in Overspeed Protect Command active mode until the pilot takes a positive action by selecting a new autopilot mode or by disengaging the autopilot.

Abstract: In order to pilot an aerodyne between two geographical positions associated with a transition route constraint, the method according to the invention comprises the application to each of these positions of a first transformation transforming the rhumb lines into straight lines, the construction of a path joining these positions while complying with the route constraints associated with these points, by means of arcs of circles and segments of straight lines tangential to the arcs of circles, the computation of the respective positions of the intermediate points spaced not far apart on said path, the application to said intermediate points of a transformation that is the reverse of said first transformation, so as to determine the geographical position of these intermediate points, and the automatic control of the piloting of the aerodyne so that its path flies over each of said intermediate points.

Abstract: The invention relates to novel basic architecture for helicopter autopilots, transforming the conventional objective of maintaining fuselage trim to an objective of maintaining airspeed, and without making use of coupler techniques. On each of the pitch and roll control systems, a known fraction of the output signal (24S) from the flight control is injected by being summed with the standard fuselage trim input (1S). The signal (4S) formed in this way possesses the novel static and dynamic properties required for the control loop to make the helicopter statically stable in speed throughout its flight envelope. The novel architecture makes it possible to replace the conventional fuselage trim signal (1S) with an accelerometer signal. The signal (4S) is displayed in particular on a standard "artificial horizon" type instrument (33), thereby making it possible simultaneously to monitor rapid variations of fuselage trim and to control airspeeds.

Abstract: This device relates to assistance in the piloting of an aircraft or helicopter in conditions of poor visibility, at the landing stage, during the final approach before touchdown. In case of poor visibility, the landing operations are facilitated by ILS radioelectrical installations that enable an automatic approach to the runway threshold. All that remains to be done is to perform a flare-out for touchdown which has to be done in a certain zone at the start of the runway. It happens, in the event of poor visibility, that this touchdown zone is difficult to assess. This causes risks of overshooting. The disclosed device is aimed at putting out an alarm in the event of overshooting of the nominal touchdown zone.

Abstract: Aircraft input signals of selected altitude (h.sub.s), selected vertical speed (h.sub.s), current altitude (h), current vertical speed (h), current pitch attitude (.theta.), current normal acceleration (N.sub.z) and current pitch rate (Q) are processed to produce an elevator command (.delta..sub.e) that will result in a smooth vertical transition of the aircraft for various autopilot vertical maneuvers, including altitude capture and hold, glideslope capture and hold, and flare control.

Abstract: Apparatus and method for controlling an unmanned generally aerodynamically symmetric aircraft includes detecting the presence of misalignment between the horizontal component of the center of gravity vector and the horizontal component of the direction of flight vector and further includes rotating the aircraft to align the horizontal components of the vectors.

Abstract: A method and apparatus are provided for addressing the effect of centripetal acceleration upon estimates of cross-track velocity, for determination of east gyro bias error, generated with a taxiing aircraft. After initial estimates of crab angle, ratio of crab angle to centripetal acceleration and lever arm are provided, velocity, heading angle and heading angle rate are observed as the aircraft taxis. An estimated value of centripetal acceleration is taken as the product of heading angle rate and heading velocity. Cross-track velocity is computed from cross-heading velocity and this is integrated to generate cross-track position. A Kalman filter generates various gains, including one associated with the ratio of crab angle to centripetal acceleration, for error allocation.

Abstract: A method for revising the lateral flight plan of an aircraft uses a designator and validator device connected to the flight management system of the aircraft to obtain on the ND screen of this system, in addition to a geographical representation of the flight plan selected by the pilot, a touch-sensitive area associated with each point of the ND screen and dynamically assigned function areas for constructing menus. A cursor is moved on the ND screen by action of the pilot on the designator device. The cursor can be moved to a function area or a touch-sensitive area. activating the function represented by the function area or selecting the touch-sensitive area being achieved by action of the pilot on the validator device. The method simplifies the task of the pilot who has only one display screen to monitor.

Abstract: An improved fly by wire or fly by light system which provides for a smooth manual override of the autopilot function by allowing the pilot to modify the autopilot command to the control surfaces by physically manipulating the yoke to a position different than that commanded by the autopilot.

Abstract: An aircraft rudder command system for allowing a pilot to directly input a sideslip command for yaw-axis control through use of the rudder pedals is disclosed. The aircraft rudder command system includes a pedal input device for receiving a pedal input signal indicative of pilot rudder pedal input, a signal-receiving device for receiving feedback signals indicative of the current state of aircraft operation, a command generator system responsive to the pedal input signal and at least one of the feedback signals for generating a sideslip angle command, command limiting means for generating a limited sideslip angle command, and a feedback control system for transmitting a sideslip control rudder command to a rudder actuation system. The rudder actuation system causes the rudder to move in such a manner so as to produce an actual aircraft sideslip angle which follows the limited sideslip angle command. The aircraft rudder command system may also include a sideslip estimator.

Abstract: A helicopter flight control system includes a model following control system architecture having provisions to compensate for Euler singularities which occur when the pitch attitude of the helicopter starts to approach ninety degrees. The control system processes information from a variety of helicopter sensors in order to provide the control commands to the helicopter main and tail rotors. The present invention synchronizes a sensed attitude signal, and a desired attitude signal as the pitch attitude of the helicopter approaches ninety degrees to compensate for the Euler singularities.

Abstract: An aircraft flight control system including model following control laws includes improved logic and algorithms to limit the error between a desired parameter value from the output of a model and an actual parameter value. Such logic is operable to sense the amount of said parameter error and to limit the amount of the error if it exceeds a predetermined value. The difference between the predetermined limit and the actual error is fed back to the model such that the output of the model is adjusted so that the error between the desired and actual parameter values does not exceed the predetermined value.

Abstract: A digital power controller for an aircraft having a computer controlled propulsion system is disclosed. A movable handle is in electronic communication with a computer of the propulsion system. The handle is positionable is a neutral location for maintaining the aircraft at a substantially constant velocity and for resetting power command functions. The handle is also positionable in locations forward the neutral location for commanding increasing acceleration rates of the aircraft. The handle is positionable in locations aft the neutral location for commanding increasing deceleration rates of the aircraft. A switch associated with the movable handle is engagable at the forward end aft positions for commanding velocity hold and for permitting reposition of the handle to the neutral location for resetting power command functions.

Abstract: A control system for an aircraft or other man-machine system wherein the usual visual feedback system is characterized and is optimally supplemented by a secondary feel oriented feedback arrangement in which input signals are derived from either of two supplementary feedback signal sources and the resulting algorithms characterized mathematically. The disclosure includes several exemplary arrangements of the feedback systems in which some of the input parameters are of selected value. Mathematical characterization of the feedback paths is used.

Abstract: A throttle controller includes a manual mode where engine output is a function of throttle lever angle, and an alternatively selectable Speed Hold/Thrust Hold mode which is initiated when the throttle is placed in a center Hold position. Throttle operation during the Speed Hold mode is governed by (i) a selected speed entered at a mode control panel, or (ii) a speed existing when the lever was placed in the Hold detent, given that no speed is entered and selected at that mode control panel or (iii) a selected speed/thrust designated in a flight plan entered into the aircraft's flight management computer. During the Thrust Hold mode, the throttle controller maintains the level of thrust which existed when the throttle lever was placed in the Hold detent. When out of the Hold detent, the controller operates to control output thrust as a function of Throttle Lever Angle (TLA).

Abstract: A throttle controller includes a manual mode where engine output is a function of throttle lever angle, and an alternatively selectable Speed Hold/Thrust Hold mode which is initiated when the throttle is placed in a center Hold position. Throttle operation during the Speed Hold mode is governed by (i) a selected speed entered at a mode control panel, or (ii) a speed existing when the lever was placed in the Hold detent, given that no speed is entered and selected at that mode control panel or (iii) a selected speed/thrust designated in a flight plan entered into the aircraft's flight management computer. During the Thrust Hold mode, the throttle controller maintains the level of thrust which existed when the throttle lever was placed in the Hold detent. When out of the Hold detent, the controller operates to control output thrust as a function of Throttle Lever Angle (TLA).

Abstract: A servo control apparatus comprising: a unit for generating a required position value; an actuator adapted to drive a load and having its maximum displacement in one direction and its minimum displacement in the other direction; a drive unit for driving the actuator; a position detecting unit for detecting a displacement of the actuator; and a control unit for correcting the required position value on basis of the maximum and minimum displacements of the actuator and for computing a deviation between an output of the position detecting unit and the corrected required position value; the load being controlled by the deviation.

Abstract: An automatic control system for an aircraft has a first controller connected to the pilot's operating controls. The first controller manipulates the operating controls so that goals, expressed in terms of selected aircraft parameters, are achieved. A second controller supplies a series of goals to the first controller so that the aircraft will perform desired maneuvers. A third controller acts as a mission planner, and supplies desired maneuvers to the second controller in accordance with overall mission plans.

Abstract: An aircraft guidance system for optimizing the flight path of an aircraft in the presence of a windshear maximizes the time the aircraft remains in the air and the distance traveled regardless of the magnitude of the windshear, in the presence of horizontal or vertical windshear components, while effectively minimizing excitation of the aircraft's phugoid mode. A flight path angle is commanded sufficient to clear any obstacle that may be found in the airport vicinity. For longitudinal or horizontal shears, a slightly positive constant flight path angle which is a function of the magnitude of the vertical wind is added to the slightly positive flight path angle command to produce a modified command that compensates for the decrease in flight path angle relative to the ground caused by the vertical wind. The system inhibits exceeding stick shaker angle of attack by reducing the command signal until the actual angle of attack is equal to or less than the stick shaker angle of attack.

Abstract: A biomechanical feedback arrangement wherein a varying force tending to improve the neuromotor tracking response of a human subject, particularly in the presence of lateral or front-back G force fields, is added to the test subject input member of a feedback control system. Use of the biomechanical feedback in a high-performance aircraft and in a ground-based simulator apparatus is also disclosed, along with comparison results from simulator testing of human subjects.

Abstract: Apparatus and method of anti-armor missile trajectory shaping for optimum rhead penetration of armor by a guided missile. The guidance system utilizes a terminal homing guidance unit in conjunction with a programmed control signal through the missile's autopilot to cause the missile to cruise at low altitudes and then dive onto the armor target. The terminal dive angle can be selected dependent upon the target's armor characteristics.

Abstract: An aircraft spoiler control system wherein each of a plurality of spoilers is driven by an electrically responsive actuator. Sense signals representative of the aircraft's control wheel rotation, speedbrake lever deflection, flap position and air/ground status are processed by logic which produces a corresponding control signal for each actuator. Fault detection circuitry switches in a back-up actuator control signal if a fault occurs in the active control circuit.

Abstract: An aircraft heading reference system including a rate gyro for sensing the yaw rate .psi. of the aircraft and producing electrical signals representative thereof; a rate gyro for sensing the pitch rate .theta. of the aircraft and producing electrical signals representative thereof; an air speed sensor for sensing air speed .nu. of the aircraft and producing electrical signals representative thereof; a roll angle computer circuit responsive to the electrical signals representative of the yaw rate, pitch rate, and air speed for determining an electrical signal representative of the roll angle .phi. according to the expression:.psi.+.theta. tan .phi.=(tan .phi./cos .phi..multidot.(g/.nu.)and a yaw and pitch computer circuit, responsive to the pitch rate .theta., yaw rate .psi., and roll angle .phi. electrical signals for resolving the yaw and pitch rates about the roll angle .phi. to provide an electrical signal representative of the heading rate of the aircraft.