Bell Helicopter Textron Inc.

Textron Innovations Inc.

Textron Corporation

Kenneth E Builta

Kynn Jerald Schulte

James E Harris

Bryan P Honza

Carlos Alexander Fenny

Bryan Paul Honza

Jeffrey W Epp

Billy K Gore

Robert L Fortenbaugh

Paul Eugene Darden

Samuel W Ferguson, Iii

Kynn J Schulte

A method and apparatus provide for automatically controlling the flight of a tiltrotor aircraft while the aircraft is in flight that is at least partially rotor-borne. The method and apparatus provide for automatically tilting nacelles in response to a longitudinal-velocity control signal so as to produce a longitudinal thrust-vector component for controlling longitudinal velocity of the aircraft. Simultaneously, cyclic swashplate controls are automatically actuated so as to maintain the fuselage in a desired pitch attitude. The method and apparatus also provide for automatically actuating the cyclic swashplate controls for each rotor in response to a lateral-velocity control signal so as to produce a lateral thrust-vector component for controlling lateral velocity of the aircraft. Simultaneously, collective swashplate controls for each rotor are automatically actuated so as to maintain the fuselage in a desired roll attitude. The method and apparatus provide for yaw control through differential longitudinal thrust produced by tilting the nacelles.

Method and apparatus for flight control of tiltrotor aircraft

A method for operating an aircraft to prevent/recover from a stall condition includes the steps of detecting an actual vertical velocity of the aircraft, calculating vertical velocity error of the aircraft, the vertical velocity error being based upon a comparison between the actual vertical velocity of the aircraft and a commanded vertical velocity of the aircraft, and determining if the aircraft is in one of a near stalled condition and a stalled condition based upon at least the detected vertical velocity error and the polarity of the vertical velocity error. The method further includes the steps of taking control of the aircraft from an operator of the aircraft, reducing a bank angle of the aircraft, pitching the aircraft downward, and increasing the airspeed of the aircraft if the aircraft's airspeed is outside an airspeed window if the aircraft is in one of the near stalled condition and the stalled condition.

Stall prevention/recovery system and method

The difference between a first position of a first pylon of a tiltrotor aircraft and a second position of a second pylon of the aircraft is prevented from becoming too large. An actuator position error for the first pylon is calculated from a difference between the first position and a commanded first position of the first pylon. An actuator position error for the second pylon is calculated from a difference between the second position and a commanded second position of the second pylon. An absolute value of the actuator position error for the first pylon is compared to the preset limit. If the absolute value of the actuator position error for the first pylon is greater than or equal to a preset limit, the actuator position error for the second pylon is calculated from the difference between the first position and the second position.

Conversion system fault management system for tiltrotor aircraft

A system for controlling flight of an aircraft has sensors, a receiver, and a digital control system, all of which are carried aboard the aircraft. The sensors determine the position of the aircraft relative to the earth and the inertial movement of the aircraft. The receiver receives transmitted data communicating the position and movement of a reference vehicle relative to the earth. The control system calculates the position and velocity of the aircraft relative to the reference vehicle using the data from the sensors and the receiver and then commands flight control devices on the aircraft for maneuvering the aircraft in a manner that maintains a selected position and/or velocity relative to the reference vehicle. The system allows use of a graphical or tactile user interfaces.

Control system for vehicles

A flight control system for an aircraft is configured for receiving command signals representing commanded values of a location of a geospatial point and a radius about the geospatial point for defining a circular groundtrack. A sensor determines a geospatial location of the aircraft and provides a location signal representing the location of the aircraft. A controller for commanding flight control devices on the aircraft controls the flight of the aircraft and is configured to receive the command signals and the location signal. The controller uses the command signals and location signal to operate the flight control devices to control the flight of the aircraft for directing the aircraft generally toward a tangent point of the circular groundtrack and then maintaining a flight path along the circular groundtrack.

Control system for automatic circle flight

A system for controlling flight of an aircraft has sensors (), a receiver (), and a digital control system (), all of which are carried aboard the aircraft. The sensors () determine the position of the aircraft relative to the earth and the inertial movement of the aircraft. The receiver () receives transmitted data () communicating the position and movement of a reference vehicle relative to the earth. The control system () calculates the position and velocity of the aircraft relative to the reference vehicle using the data from the sensors () and the receiver () and then commands flight control devices () on the aircraft for maneuvering the aircraft in a manner that maintains a selected position and/or velocity relative to the reference vehicle. The system allows use of a graphical or tactile user interfaces.

A flight control system for an aircraft receives a selected value of a first parameter, which is either the airspeed or inertial velocity of the aircraft. A primary feedback loop generates a primary error signal that is proportional to the difference between the selected value and a measured value of the first parameter. A secondary feedback loop generates a secondary error signal that is proportional to the difference between the selected value of the first parameter and a measured value of a second flight parameter, which is the other of the airspeed and inertial velocity. The primary and secondary error signals are summed to produce a velocity error signal, and the velocity error signal and an integrated value of the primary error signal are summed to produce an actuator command signal. The actuator command signal is then used for operating aircraft devices to control the first parameter to minimize the primary error signal.

Automatic velocity control system for aircraft

One embodiment of the present invention is a method for automatically reducing the effect of a component of an external force that is laterally incident on a rotorcraft. A signal of the rotorcraft indicative of and proportional to the component is monitored. An absolute value of the signal and a preset high limit are compared. If the absolute value is greater than the preset high limit, manual heading control of the rotorcraft is disabled and the heading of the rotorcraft is adjusted with respect to the external force so as to decrease the lateral component of the external force experienced by the rotorcraft.

Automatic heading control system for tiltrotor aircraft and helicopters

One embodiment of the present invention is a method for automatically controlling the conversion of a tiltrotor aircraft. An airspeed command for the tiltrotor aircraft is received. The airspeed command is converted to a pylon position. A difference between the airspeed command and a measured airspeed is calculated. The difference between the airspeed command and a measured airspeed is converted to a dynamic pylon position. A total pylon position is calculated from the pylon position and the dynamic pylon position. A pylon of the tiltrotor aircraft is moved to the total pylon position. Another embodiment of the present invention is a system for calculating a position of a pylon of a tiltrotor aircraft based on an airspeed command. The system includes an airspeed command module, a pylon trim position module, a dynamic pylon position module, and a pylon position module.

Automatic conversion system for tiltrotor aircraft

An improved tilting rotor aircraft, a control system for an elongated shaft, and a method of controlling an elongated shaft are provided. Sensors are utilized to detect an amount of twist on a flexible and elongated shaft. A controller receives the signal and generates a command signal. A plurality of actuators receive the command signals and compensate for the twist in the flexible and elongated shaft.

Method and apparatus for controlling force fights in a rotating shaft

A improved tilt rotor aircraft comprising an airframe, at least one wing member coupled to the airframe, and at least one tilt rotor assembly carried by the at least one wing member. The tilt rotor assembly is actuated by a conversion actuator. The improved tilt rotor aircraft includes a low-height tunable downstop assembly for absorbing oscillatory vibration loads, such as pitch loads and yaw loads. The low-height tunable downstop assembly comprises a pivotable striker assembly having an adjustable stiffness, and a cradle assembly adapted to releasably receive the striker assembly. The striker assembly is attached to the tilt rotor assembly and the cradle assembly is attached to the wing member. A plurality of sensor modules are bonded to the cradle assembly. The sensor modules include a plurality of strain gages arranged to sense shear strain in the cradle assembly. The sensor modules are electrically coupled to sensor circuits and flight control computers. As the conversion actuator pivots the tilt rotor assembly into the airplane mode, the conversion actuator forces the striker assembly into contact with the cradle assembly at a selected preload. The sensor modules directly sense preload induced strain in the cradle assembly and send corresponding signals to the flight control computers via the sensor circuits. If the signals from the sensor modules indicate that the preload is insufficient to maintain stability in the aircraft mode, then a warning signal is sent to the pilot and a control signal is sent to the conversion actuator to increase the preload. Likewise, if the signals from the sensor modules indicate that the preload is too high, a warning signal is sent to the pilot and a control signal is sent to the conversion actuator to decrease the preload.

Method and apparatus for sensing preload in a tilt rotor downstop

A helicopter elevator (12, 13) is driven by an electrical actuator (20) to position the elevator (12, 13) to an optimum angle of incidence for given flight conditions. The actuator (20) is included within a servo loop which responds to a command signal derived from an electronics module (69). Transducers which measure longitudinal cyclic position (40), collective pitch position (41) and air speed (42) generate signals which are summed in a summation unit (43). The output of the summation unit (43) is passed through a signal multiplier (45) which has the gain thereof set inversely proportional to air speed. A trigger unit (56) activates the signal multiplier (45) at an air speed above a preset threshold. An integrator (54) integrates the sum of the outputs of the longitudinal cyclic position transducer (40) and air speed transducer (42) to compensate for different center of gravity aircraft loadings. Duplicate electronic systems are provided to drive a pair of servo motors (36a, 36b) with brakes provided to lock the actuator (20) when an electronic failure is detected.

Electrically controlled elevator

A fail safe aircraft control system wherein duplicate control loops are employed to produce identical control signals for the system actuator. Output signals from the two loops are compared to detect when the outputs differ by a preset limit as an indication of a component failure. When such a failure is detected the system actuator is deactivated thus preventing the failed augmentation system from affecting control of the aircraft.

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Euless, TX, United States of America

Kenneth E Builta