The Boeing Company

Sandel Avionics Inc.

Robert J Bleeg

Kioumars Najmabadi

Monte R Evans

Richard S Breuhaus

Edward E Coleman

Timothy A Nelson

Dorr Marshall Anderson

Neal Van Huynh

Gerald J Block

John A Standley

Delmar M Fadden

Richard W Taylor

Brian L Bocksch

Ralph Scott Pepper

James W Berwick, Jr

Jim Slattery

Vincent E McFaddin

John A Dley

Scott Pepper

Systems and methods are disclosed for a path guidance panel. The path guidance panel provides separate display screens for showing the lateral and vertical guidance information. Each screen shows the current state of the respective guidance system, and shows what, if anything, the respective guidance system will do next. This provides a convenient display for the pilot to see the current and future states of the guidance systems of the aircraft.

System and method for path guidance panel

A flight management system for controlling navigation of an aircraft includes a path guidance panel having at least one mode selector for allowing user selection of guidance mode in which the aircraft is to operate. The system also includes at least one display device having rendered thereon a navigational map. A computer module receives aircraft data from one or more aircraft components, at least one of the aircraft components being a guidance system. The computer module is configured to use information obtained from the guidance system to cause the display device to render on the navigational map a current flight plan path that is the aircraft is following.

System and method for aircraft operations including path guidance panel with conditional waypoints

Systems and methods for controlling aircraft flaps and spoilers are disclosed. Systems in accordance with some embodiments include a wing having a trailing edge, and a flap positioned at least partially aft of the wing trailing edge and being deployable relative to the wing between a first flap position and a second flap position by operation of a first actuator. A spoiler can be positioned at least proximate to the trailing edge and can be movable among at least three positions, including a first spoiler position in which the spoiler forms a generally continuous contour with an upper surface of the wing, a second spoiler position deflected downwardly from the first, and a third spoiler position deflected upwardly from the first. A second actuator can be operatively coupled to the spoiler to move the spoiler among the first, second and third spoiler positions in a manner that is mechanically independent of the motion of the flap.

Systems and methods for controlling aircraft flaps and spoilers

Systems and methods for controlling aircraft flaps and spoilers. Systems in accordance with some embodiments include a wing having a trailing edge, and a flap positioned proximate to the wing trailing edge and being deployable relative to the wing between a first flap position and a second flap position by operation of a first actuator. A spoiler can be positioned at least proximate to the trailing edge and can be movable among at least three positions, including a first spoiler position in which the spoiler forms a generally continuous contour with an upper surface of the wing, a second spoiler position deflected downwardly from the first, and a third spoiler position deflected upwardly from the first. A second actuator can be operatively coupled to the spoiler to move the spoiler among the first, second and third spoiler positions in a manner that is mechanically independent of the motion of the flap.

A bias correction system for use in a neutrally-stable aircraft having a control column and a pitch control system is provided. The position of the control column is generally represented by a control column position signal. The bias correction system is for removing control column bias when the control column is in a neutral position. The bias correction system includes a first combining unit for combining the control column position signal and a correction signal, and a switch. The switch includes activated and deactivated states. The switch is set to the deactivated state when the control column is physically displaced from its neutral position. The deactivated state allows the correction signal to remain at its last value, the activated state allows the correction signal to equal approximately the control column position signal. In one embodiment, the bias correction system includes a smoothing filter for receiving the control column position signal and outputting a smoothed correction signal when the switch is in the activated state. In another embodiment, the bias correction system includes a unit for limiting the correction signal to control column position signal values within a limited predefined range.

Aircraft pitch-axis stability and command augmentation system

A method of providing speed protection to a neutrally-stable aircraft including the calculation of at least one of a stall protection signal and an overspeed protection signal. The aircraft includes a pitch control system having a command pitch signal received from pilot input to a control column. The protection signal is provided to the pitch control system to limit the commanded pitch signal during flight. The calculation of a stall protection signal includes differencing the aircraft's current airspeed with a reference speed and multiplying the difference by a scaling factor that corresponds to the force required by the pilot to move the control column. The calculation of an overspeed protection signal includes calculating a first signal based on airspeed, calculating a second signal based on Mach, selecting the large of the two signal and passing it through a limiter. The limiter is a function of roll attitude and corresponds to the force required by the pilot to move the control column. Also provided is a method of calculating a flare compensation signal to emulate ground effects encountered by an aircraft while executing a landing. The flare compensation signal is determined using a table of gear heights versus flare compensation amounts. The flare compensation signal is limited to a value between zero and a positive upper limit.

Method of speed protection and flare compensation for use with aircraft

Disclosed is a pitch-axis stability and command augmentation system (19) in which a pilot column input (.delta..sub.C) is provided to a pitch command processor (26), the output of which is a C*U feedforward command (C*U.sub.FFC) that is supplied as an additive input to a combining unit (20). The combining unit (20) receives a second additive input of an augmented feedback command signal (AFB.sub.COM). The resulting output is filtered and generates an elevator command signal (.delta..sub.e,FILT). The command processor (26) additionally supplies a corrected column position signal (.delta..sub.C,COR) to a pitch command C*U processor (26) that converts the corrected column position signal (.delta..sub.C,COR) into a C*U pitch command (C*U.sub.PilotCmd), representative of the pilot's requested elevator pitch which is generated by movement of the control column. A computed C*U processor (32) forms a computed C*U signal (C*U.sub.Computed) that is based on the current state of the aircraft. The C*U processor (32) includes a speed stability processor (150) that computes the difference between a reference speed (V.sub.REF) and calibrated airspeed (V.sub.CAS). The speed stability processor (150) includes phugoid damping (400) in generating a speed error signal (E.sub.z), and providing a means by which the pilot may set a reference airspeed (V.sub.REF) by using a trim device. An underspeed protection signal (C*U.sub.US) is supplied by a stall protection processor (416). A flare compensation signal (C*U.sub.FR) is supplied by a flare compensation processor (414). An overspeed protection processor (412) supplies an overspeed protection signal (C*U.sub.OS). A combining unit (424) is provided to combine the stall protection signal (C*U.sub.US), flare compensation signal (C*U.sub.FR), and overspeed protection signal (C*U.sub.OS) to form a compensation and protection signal (C*U.sub.C&P). A combining unit (22) combines the pilot-requested C*U command (C*U.sub.PilotCmd), the computed C*U signal (C*U.sub.Computed), and the compensation and protection signal (C*U.sub.C&P). The resulting signal is an error signal (E.sub.C*U) which is supplied to an integrator (28) that outputs a C*U integral command signal (IC*U.sub.COM). Another combining unit (23) receives the C*U integral command signal (IC*U.sub.COM) and combines it with a pitch rate damping command signal (Q.sub.COM). The output of the combining unit (23) is the augmented feedback command signal (AFB.sub.COM).

Disclosed is a pitch-axis stability and command augmentation system (19) in which a pilot column input (.delta..sub.C) is provided to a pitch command processor (26), the output of which is a C*U feedforward command (C*U.sub.FFC) that is supplied as an additive input to a combining unit (20). The combining unit (20) receives a second additive input of an augmented feedback command signal (AFB.sub.COM). The resulting output is filtered and generates an elevator command signal (.delta..sub.e,FILT). The command processor (26) additionally supplies a corrected column position signal (.delta..sub.C,COR) to a pitch command C*U processor (26) that converts the corrected column position signal (.delta..sub.C,COR) into a C*U pitch command (C*U.sub.PilotCmd), representative of the pilot's requested elevator pitch which is generated by movement of the control cola. A computed C*U processor (32) forms a computed C*U signal (C*U.sub.Computed) that is based on the current state of the aircraft. The C*U processor (32) includes a speed stability processor (150) that computes the difference between a reference speed (V.sub.REF) and calibrated airspeed (V.sub.CAS). The speed stability processor (150) includes phugoid damping (400) in generating a speed error signal (E.sub.z), and providing a means by which the pilot may set a reference airspeed (V.sub.REF) by using a trim device. An underspeed protection signal (C*U.sub.Us) is supplied by a stall protection processor (416). A flare compensation signal (C*U.sub.FR) is supplied by a flare compensation processor (414). An overspeed protection processor (412) supplies an overspeed protection signal (C*U.sub.OS). A combining unit (424) is provided to combine the stall protection signal (C*U.sub.Us), flare compensation signal (C*U.sub.FR), and overspeed protection signal (C*U.sub.OS) to form a compensation and protection signal (C*U.sub.C&P). A combining unit (22) combines the pilot-requested C*U command (C*U.sub.PilotCmd), the computed C*U signal (C*U.sub.Computed), and the compensation and protection signal (C*U.sub.C&P). The resulting signal is an error signal (E.sub.C*U) which is supplied to an integrator (28) that outputs a C*U integral command signal (IC*U.sub.COM). Another combining unit (23) receives the C*U integral command signal (IC*U.sub.COM) and combines it with a pitch rate damping command signal (Q.sub.COM). The output of the combining unit (23) is the augmented feedback command signal (AFB.sub.COM).

Aircraft pitch-axis stability and command augmentation

A controller for reducing unwanted sideways motion of an aircraft by reducing lateral side loads, resulting from air mass, turbulence and gusts. The controller functions in a manner that in the presence of higher frequency side loads, the rudder is caused to move in a relieving direction so that the net force across the vertical stabilizer is reduced. A pressure differential across opposite sides of the vertical stabilizer is measured and used to generate a first rudder deflection signal. To maintain stability of the aircraft, a beta-dot signal from the yaw damper module is gain adjusted and filtered to generate a second rudder deflection signal which is added to the first rudder deflection signal. The resulting combined signal reduces lateral side loads at higher frequencies without comprising aircraft directional stability.

Apparatus and method for reducing aircraft loads resulting from

An aircraft guidance system operable during both takeoff and go-around maneuvers allows safe operation even under adverse wind shear conditions. Initially the system controls vertical speed to achieve an acceptable, minimum climb rate regardless of airspeed loss. Should excess thrust exist such that the acceptable minimum climb rate is exceeded, increased vertical speed and airspeed are commanded. When the climb rate significantly exceeds the acceptable minimum, airspeed is controlled to a specified value. Should the aircraft ever approach the 'stick shaker' condition, an override command limits aircraft angle of attack.

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Mercer Island, WA, United States of America

Robert J Bleeg