Abstract: An airborne glide slope tracking system includes a radio altimeter for producing a signal indicative of the instantaneous altitude of the aircraft and a glide slope error indicator for producing a signal indicative of the angular deviation from glide slope. The two signals are multiplied to produce a glide slope error signal in feet and fed to a summing device directly and through a lead filter to generate a signal for a throttle servo to increase or decrease the thrust of the aircraft. A longitudinal accelerometer signal is then added to a signal indicative of a difference between an aircraft's reference angle of attack and actual angle of attack to produce a signal to a pitch command to provide an angle which will be sustained by the power of the aircraft.

Abstract: An airborne power control system for automatically controlling the power of an aircraft during landing is disclssed. The system includes a computer and a minimum airspeed program as a function of altitude. A radio altimeter or the like senses the instantaneous altitude of the aircraft while a pitot tube or the like measures indicated airspeed. A computer and program compare programmed airspeed with actual airspeed for a given altitude. And, a servomotor is provided for decreasing engine thrust where the actual airspeed exceeds the programmed airspeed at any given altitude. An inhibitor inhibits the decrease in engine thrust if the airspeed drops below the programmed airspeed.

Abstract: A robotic or remotely controlled flying platform (10) with reduced drag stabilizing control apparatus constructed having an air duct (12) with an air intake (14) on the top and an exhaust (16) at the bottom, containing supported therein a clockwise rotating fan (22) and a counter-clockwise rotating fan (24). Directly below the perimeter of the air duct exhaust are mounted a plurality of trough shaped air deflection assemblies (32) each including a rotatably adjustable half trough (44) for selectively scooping a portion of the drive air, and a stationary adjacent half trough (36) for receiving the scooped drive air and redirecting it outward and upward from the air duct.

Abstract: An airborne power control system for automatically controlling the power of an aircraft during landing is disclssed. The system includes a computer and a minimum airspeed program as a function of altitude. A radio altimeter or the like senses the instantaneous altitude of the aircraft while a pitot tube or the like measures indicated airspeed. A computer and program compare programmed airspeed with actual airspeed for a given altitude. And, a servomotor is provided for decreasing engine thrust where the actual airspeed exceeds the programmed airspeed at any given altitude. An inhibitor inhibits the decrease in engine thrust if the airspeed drops below the programmed airspeed.

Abstract: A method of automatically controlling an aircraft to avoid a vertical zone includes several steps. The aircraft first acquires limits of the zone to be avoided. The zone is modeled by a cylindrical volume which is limited by a horizontal contour with upper and lower altitudes of the zone. The cylindrical volume associated with a scheduled route of the aircraft is located and points of entry and exit in the cylindrical volume are determined. A new flight altitude is calculated in order to avoid the zone. A point of change of altitude is calculated to reach an avoidance altitude. The new flight altitude is updated and the point of change of altitude is input into an automatic pilot.

Abstract: A self adaptive limiter for use in aircraft control systems during approach and landing is disclosed. Estimated flight path angle is continuously computed during the glidepath tracking phase until a predetermined lock altitude above ground when a nominal flight path angle is latched. Nominal vertical speed is continuously computed below the lock altitude as a function of the latched nominal flight path angle and ground speed of the aircraft. A vertical speed limit function is generated as a function of the nominal vertical speed and radio altitude. During approach and landing a pitch limit is computed from the vertical speed limit, vertical speed, ground speed, and pitch. When a pitch command to the autopilot exceeds the pitch limit(i.e. commands excessive pitch down attitude), it is limited to the pitch limit thus preventing the aircraft from descending below a safe altitude.

Abstract: An underspeed protection system for an aircraft under autopilot control selects a target speed based upon the greater of a minimum maneuver speed and a stick shaker speed. The system then compares a monitored speed to the target speed to produce an error signal. The system also monitors vertical speed to determine if the aircraft begins to descend. If the error signal due to the underspeed condition causes the aircraft to descend, the system provides a hold zero vertical speed signal in place of the error signal such that the aircraft seeks to maintain its altitude. The hold zero vertical speed signal overrides the underspeed error signal such that the aircraft does not pitch forward to seek an increased speed.

Abstract: The disclosure is an aircraft collision-avoidance device used notably during landing approach. The device includes at least means of fixing the position of the aircraft, a database, first means of comparison, means of extraction of the point of touch-down on the runway, means of calculating the real glide slope vector of the aircraft, means of extraction of the theoretical glide slope vector of the aircraft, and second means of comparison used to compare the real and theoretical slope vectors. It is applicable in particular to airliners during final approach to landing.

Abstract: The piloting assistance method embodying the invention consists in displaying on a screen of the cockpit of the aerodyne, in a Cartesian reference space representing altitude/distance from ground along a horizontal plane, way points identified by a series of marks, plural way points, a reference vertical profile connecting the way points by connecting lines and altitude constraint symbols indicating a minimum altitude at which the aerodyne can pass, a maximum altitude at which the aerodyne can pass and an altitude at which the aerodyne must necessarily pass, a symbol indicating the position of the aerodyne. The invention enables the pilot to have a clear vision of the vertical situation of the aerodyne along its flight plan.

Abstract: In an aircraft, there is included a Flight Management System (FMS), Autopilot and Autothrottle for controlling the aircraft. The apparatus has a plurality of outputs for definition of the real-time targets, controlled by the Autopilot and Autothrottle, to guide the vertical position of the aircraft to a desired vertical position along the desired vertical flightplan (or profile) according to a set of operational procedures. The FMS includes an apparatus that comprises an element which provides information denoting actual vertical position of the aircraft, and an element which generates information specifying the desired vertical position of the aircraft along the predetermined desired flightplan.

Abstract: A path capture forcing function generator utilizes integrations from an initial altitude rate error signal and an initial altitude error signal to provide an altitude rate signal and an altitude signal which are compared to the altitude rate error and altitude error signals to provide path rate error and path error signals to the autopilot. In the process of providing the altitude rate and altitude error signals an altitude acceleration signal is established which is used as a feed forward predictor for expected normal acceleration of the aircraft.

Abstract: This invention provides a flare control modification for pilot-in-the-loop aircraft maneuver command electronic flight control systems. During flare pilot pitch control inputs are interpreted as incremental flight path angle commands above a reference flight path angle (nominally a -3.degree. glideslope). The flare control modification allows conventional piloting technique (i.e., pitch controller pull-and-hold) to be used during flare, while retaining the benefits of the maneuver command system masking effects of gusts, winds, windshear, and variations in airplane weight, balance, and aerodynamic configuration.

Abstract: An aircraft navigational system includes a digital computer, a first visual output for displaying to an aircrew a geometrical representation of an optimum descent guideslope relative to a symbol representing the location of the aircraft, and a second visual output for displaying optimum descent guideslope information, such as altitude and distance, alphanumerically. The visual display provides guidance information to the aircrew to control the descent of the aircraft along an optimum descent guideslope which is referenced to a selected end of descent waypoint such as a destination airport or to an instrument approach marker. The guideslope is calculated in accordance with predetermined airspeed and altitude requirements with define a descent profile which is fuel efficient and which complies with aviation regulations regarding maximum airspeed limitations.

Abstract: An improved apparatus for automatic deceleration of a helicopter to substantially zero air speed from a glide slope descent. Parameters available in a typical helicopter automatic flight control system, such as vertical speed, glide slope error, longitudinal acceleration, and pitch attitude are blended together in a manner as to approximate helicopter ground speed which is then used in a control law computation to decelerate the helicopter to a near hover speed as the vehicle approaches a preset altitude minimum.

Abstract: An aircraft navigational system for providing a geometic display to the pilot of the vertical position of the aircraft relative to a selected vertical flight path profile as the aircraft progresses between its departure and destination. The vertical profile, which has a climb portion, a cruise portion and a descent portion, is formed by a number of navigational points, each of which is defined by an altitude component and a geographical position component. The profile points which form the climb and descent profiles are calculated from controller assigned altitude and waypoint constraints as well as from airspeed and altitude operating limitations which define various selected aircraft performance climb and descent modes.

Abstract: The disclosed apparatus is useful in controlling an aircraft inertial flight path on the final approach to landing to coincide with a desired flight path represented by the intersection of the localizer and glide path electronic beam signals, and there being an inertially responsive unit on the aircraft.The apparatus includes:(a) Dual antenna systems carried on the aircraft spaced apart laterally to receive said beams, and(b) circuitry responsive to the beam signals received by the two antenna systems to coact with the inertially responsive unit to produce an output for controlling the inertial path of the aircraft to coincide with the desired path defined by said localizer and glide path signals.

Abstract: An improved inner loop control is disclosed for use in an aircraft pitch axis control system. Signals representative of vertical acceleration and pitch rate are complementary filtered and summed to produce an "inner loop" damping signal which is combined with "outer loop" command signals to produce an elevator control signal. The improved approach provides higher outer loop control law gains and better system stability, resulting in higher control bandwidth and tighter command tracking, especially in turbulence.

Abstract: A method for calibrating and correcting circuit errors induced in an aircraft ILS receiving system is disclosed herein. The spectral components of the input signal are processed through bandpass filters and then rectified and integrated in order to produce an output signal representative of the deviation from the desired centerline course. First and second test signals are generated corresponding to first and second known positions. These test signals are sequentially processed in the same manner as the operational input signals in order to develop first and second position errors which represent distortion and circuit induced errors within the processing system. A corrective function is then derived for a range of operational input signals by interpolating from the first and second position errors. All subsequently processed signals are adjusted by applying the corrective function.

Abstract: In an aircraft autopilot there is provided a first circuit selectively actuatable to connect the glide slope adapter to electrical power during the existence of a glide deflection signal, the glide slope adapter being disconnected from electric power when either (a) the altitude hold is inoperative, or (b) the altitude hold is operative and at least one of the following exists: the radio switch LOC NORM is off, or glide slope coupling has been achieved; a second circuit for maintaining the autopilot under control of a preset heading on the pictorial navigation indicator when the radio coupler is preset for radio navigation control including a third circuit for placing the automatic pilot under control of the radio coupler and interrupting heading control by the pictorial navigation indicator upon actual intercept of the omni occurring; and a selectively actuatable fourth circuit for disconnecting the console pitch signal from the amplifier and interconnecting a pitch signal of predetermined fixed magnitude to

Abstract: A method for controlling the decelerated approach of an aerodyne on radioignment generally referred to as "glide," in particular for the purpose of reducing its fuel consumption and reducing the noise, mainly in the final stage of the approach, by approaching the glide at high speed, initiating the descent according to the glide angle, also at high speed then beginning a decelerated phase in order to reach a predetermined speed V sel+k, k being generally equal to 10 knots, on reaching a predetermined altitude, h.sub.F, and then passing progessively from the speed V sel+k to the speed V sel. From the time when the aerodyne reaches the axis of the glide path, a course is determined in two sections, as regards the variables of speed and altitude, for reaching the point situated at the altitude h.sub.

Abstract: The invention relates to a system utilizing an exponential control law for glide slope capture. The capture maneuver from above or below the beam, is a function of decreasing glide slope beam error. The present autopilot approach coupler is an altitude rate command system which provides switchless signal processing during glide slope capture, and tracking.

Abstract: A landing approach system for aircraft utilizing ILS (instrument landing system) beam guidance information which converts angular glide slope information into linear altitude deviation relative to the glide slope beam center, using computed range information.

Abstract: In a helicopter having an automatic pilot system for controlling the main rotor collective pitch, main rotor longitudinal and lateral cyclic pitch, and tail rotor collective pitch, in response to inputs provided from normal sensors, navigational instruments and controls, (which include among others as is known, doppler radar, a pilot panel, an air speed indicator, a radar altimeter, a barometric altimeter and a longitudinal accelerometer), commands to control automatic approach to a desired hover position are provided to the autopilot system in response to signals normally supplied by a TACNAV unit as well as the sensors, navigational units and controls; controls are provided for altitude and rate of descent, so as to terminate in hover at a desired altitude, as well as for turns, courses, and speed so as to approach the desired point of hover upwind and to decelerate in a controlled fashion to hover.

Abstract: In the vertical path control of an area navigation system when the aircraft ascends or descends to a waypoint with an "at-or-above" or "at-or-below" altitude requirement, vertical steering is effected at a constant airspeed with regard to a reference airspeed. An alert device is included to provide a warning to the pilot when the aircraft flight path angle is less than the straight line flight path angle to the waypoint for "at-or-above" waypoints or when the flight path angle of the aircraft is greater than the straight line flight path angle to the waypoint for waypoints with an "at-or-below" altitude requirement. Altitude error is displayed on a vertical deviation indicator of an aircraft flight instrument in accordance with the difference between the actual aircraft altitude and the altitude of the next waypoint having a firm altitude requirement.

Abstract: An aircraft flight control system utilizing inertial sensor input signals for providing glide slope and flare control signals in the aircraft. The autopilot approach and landing coupler system is a vertical velocity command control system utilizing signals representative of integrated vertical acceleration for providing feedback signals representative of vertical velocity and utilizing signals representative of ground velocity to provide the switchless glide slope and flare vertical velocity command signals.

Abstract: The invention relates to a system utilizing an exponential control law for glide slope capture. The capture maneuver from above or below the beam, is a function of decreasing glide slope beam error. The present autopilot approach coupler is an altitude rate command system which provides switchless signal processing during glide slope capture, and tracking.

Abstract: A vertical navigation steering control for aircraft for providing vertical steering of an aircraft along flight paths defined by waypoints which recognizes the problems introduced by lateral steering and accounts for shortening of horizontal distance between waypoints due to lateral path changes. Further features include means for capturing a NAV path, following a path between waypoints, transitioning from one path to another at a waypoint, and means for protection against excessive climb rates.

Abstract: A vertical glide slope control system including a complementary filter. Complementary aircraft generated and processed signals are added to a glide slope error signal which may include distortions. The results of the summation are applied to a first order lag filter having a low breakpoint (i.e., a low-pass band). The low-pass band eliminates high frequency distortions contained in the glide slope error signal without loss of stability. Aircraft stability is not lost due to the inclusion of the complementary aircraft signals added prior to the passage of the signals through the first order lag filter.