The Boeing Company

Kioumars Najmabadi

Edward E Coleman

Monte R Evans

Brian Kenyon Rupnik

Robert J Bleeg

Chuong B Tran

Sherwin Chunshek Li

Brian T Whitehead

Richard S Breuhaus

John Koon-hung Ho

David W Grubb

Dorr Marshall Anderson

Timothy A Nelson

Tze L Siu

Paul C Strefling

Tristan C Flanzer

Michael D Bills

Khaled Abdel-Motagaly

Alexander C Ho

Paul Jay Cline

Nikos D Mills

Dianne Lynn McMullin

Heidi M Haugeberg

Uy-Loi Ly

Michael S McCuen

William F Shivitz

Rafael E Caicedo

Christopher D Dallara

Leo L Engstrom, Jr

Hisako Yamashiro

Thomas C Ramsey, Jr

Vincent M Walton

Steven Cook Stone

Christopher J Borland

David P Marquis

Kevin H Milligan

William F Carver

Robert James McIntosh

Bruce Mays Fritchman

Richard D Johnson

John Koon-Hung Ho

Thomas C Ramsey

Paul J Cline

Robert J Mcintosh

John K Ho

Dianne L Mcmullin

Systems, methods, and apparatus for sensor fault detection and identification using residual failure pattern recognition are disclosed. In one or more embodiments, a method for sensor fault detection and identification for a vehicle comprises sensing, with sensors located on the vehicle, data. The method further comprises performing majority voting on the data for each of the types of data to generate a single voted value for each of the types of data. Also, the method comprises generating, for each of the types of data, estimated values by using some of the voted values. In addition, the method comprises generating residuals by comparing the estimated values to the voted values. Further, the method comprises analyzing a pattern of the residuals to determine which of the types of the data is erroneous to detect and identify a fault experienced by at least one of the sensors on the vehicle.

Sensor fault detection and identification using residual failure pattern recognition

An example method of limiting an aircraft to a pitch axis envelope includes determining aircraft state limits associated with multiple pitch axis variables of an aircraft, determining predicted aircraft states, comparing the predicted aircraft states to the aircraft state limits to produce aircraft state errors, translating the aircraft state errors into a set of positive and negative limit elevator commands, selecting a highest priority positive limit elevator command, selecting a highest priority negative limit elevator command, limiting a primary pitch axis control law elevator command of the aircraft to a value that is less than or equal to the highest priority positive limit elevator command and greater than or equal to the highest priority negative limit elevator command, and controlling the aircraft according to the primary pitch axis control law elevator command limited to the value.

Systems and methods for pitch axis envelope limiting of an aircraft

A flight control system for an aircraft is disclosed. The flight control system includes one or more processors and a memory coupled to the processors. The memory stores data comprising a database and program code that, when executed by the one or more processors, causes the flight control system to receive as input a plurality of first operating parameters that each represent an operating condition of the aircraft. The flight control system is further caused to determine a drag coefficient and a lift coefficient based on the plurality of first operating parameters. The flight control system is also caused to determine an estimated dynamic pressure based on both the drag coefficient and the lift coefficient.

Flight control system for determining estimated dynamic pressure based on lift and drag coefficients

A flight control system for an aircraft is disclosed and includes one or more processors and a memory coupled to the processors. The memory stores data comprising a database and program code that, when executed by the one or more processors, causes the flight control system to receive as input a measured angle of attack that is based on a raw angle of attack and an estimated angle of attack based on a total pressure. The flight control system is further caused to compare the measured angle of attack with the estimated angle of attack to determine an error. In response to determining that the error between the measured angle of attack and the estimated angle of attack exceeds a threshold value, the flight control system determines the presence of a fault with an angle of attack value.

Flight control system for determining a fault based on error between a measured and an estimated angle of attack

A flight control system for an aircraft is disclosed, where the flight control system detects a first common mode pneumatic event. The flight control system includes one or more processors and a memory coupled to the processors. The memory stores data comprising a database and program code that, when executed by the one or more processors, causes the flight control system to receive as input a measured dynamic pressure and an estimated angle of attack. The flight control system is further caused to determine a rate of change of the measured dynamic pressure and compare the rate of change of the measured dynamic pressure with a dynamic pressure threshold value. The flight control system is further caused to determine a rate of change of the estimated angle of attack and compare the rate of change of the estimated angle of attack with a threshold angle of attack.

Flight control system for determining a common mode pneumatic fault

Aircraft and methods and systems for controlling performance of an aircraft. The methods and systems allow the aircraft to meet a performance requirement using a set of aircraft flight data and actuators connected to control surfaces. Performance data for primary and secondary actuators are obtained to select between a primary control law for controlling the primary control surface, a secondary control law for controlling the secondary control surface, and a blended control law that controls the primary and secondary control surfaces together. If the primary control surface cannot meet the aircraft performance requirement using the primary control law, the blended control law is implemented if the primary and secondary control surfaces can be used together to meet the performance requirement; otherwise the secondary control surface is used, using the secondary control law, to meet the aircraft performance requirement.

Methods and systems for controlling aircraft flight performance

A gust compensation system is configured to adaptively reduce gust loads exerted into an aircraft. The gust compensation system may include a first sensor proximate to a front of the aircraft. The first sensor is configured to output a first signal. A second sensor may be distally located from the front of the aircraft. The second sensor is configured to output a second signal. A gust signal sub-system is configured to receive the first and second signals and generate a gust signal based on analysis of the first and second signals. The gust signal sub-system outputs the gust signal and may modify a load parameter signal in response to the gust signal exceeding a load alleviation threshold.

Gust compensation system and method for aircraft

A method for vertical gust suppression due to turbulence for an aircraft having at least one of direct lift control surfaces or pitch control surfaces. The method includes sensing atmospheric turbulence, measuring the sensed atmospheric turbulence to generate turbulence data, generating a command based on the turbulence data, and applying the command to aircraft controls to actuate the direct lift control surfaces or the pitch control surfaces based on the turbulence data. Therefore, an aircraft response to the actuation of the direct lift control surfaces or the pitch control surfaces reduces a vertical acceleration, a pitch acceleration, a pitch rate, a pitch attitude or a structural load of the aircraft due to the turbulence. Thus, the method reduces the effects of vertical gusts of wind on the aircraft, improves the comfort level for aircraft passengers and crew, and reduces diversions the aircraft may take to avoid the turbulence.

Vertical gust suppression system for transport aircraft

A method for controlling control surfaces. A position limit is identified for movement of a control surface based on a load limit set for the control surface and a number of vehicle current operation parameters to form an identified position limit. Responsive to receiving a command to move the control surface on a vehicle to a new position, the control surface is commanded to move to a position within the identified position limit.

Horizontal tail load alleviation system

A computer implemented method, apparatus, and computer usable program product for symmetric and anti-symmetric control of aircraft flight control surfaces to reduce wing-body loads. Commands are sent to symmetrically deploy outboard control surfaces to shift wing air-loads inboard based on airplane state and speed brake deployment. Surface rate retraction on a wing with peak loads is limited to reduce maximum loads due to wheel checkback accompanied by utilization of opposite wing control surfaces to retain roll characteristics. Airloads are shifted inboard on a swept wing to move the center of pressure forward, thereby reducing the tail load required to perform a positive gravity maneuver. In a negative gravity maneuver, speed brakes are retracted, thereby reducing the positive tail load and reducing the aft body design loads. High gain feedback commands are filtered from wing structural modes above one hertz by a set of linear and non-linear filters.

Wing-body load alleviation for aircraft

Vibrations due to excitation of the natural modes of an aircraft's body are suppressed by an active multi-axis modal suppression system. Dedicated sensors are positioned in the aircraft at optimal locations for sensing modal induced vibrations. The sensor produced signals are processed through control logic which, in response thereto, and in response to aircraft inertial, velocity and altitude related signals produces output control signals. The control signals effect control surface deployment creating forces to suppress the natural mode induced vibrations on multiple geometric axis's. More particularly, a symmetric and anti-symmetric control surface deployments are used on one or more geometric axis'to damp lateral, longitudinal, vertical and most importantly torsional vibrational modes.

Aircraft multi-axis modal suppression system

Vibrations due to excitation of the natural modes of an aircraft's body are suppressed by an active multi-axis modal suppression system. Dedicated sensors are positioned in the aircraft at optimal locations for sensing modal induced vibrations. The sensor produced signals are processed through control logic which, in response thereto, and in response to aircraft inertial, velocity and altitude related signals produces output control signals. The control signals effect control surface deployment creating forces to suppress the natural mode induced vibrations on multiple geometric axis's. More particularly, symmetric and anti-symmetric control surface deployments are used on one or more geometric axis's to damp lateral, longitudinal, vertical and most importantly torsional vibrational modes.

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

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Bellevue, WA, United States of America

Kioumars Najmabadi