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WO1991015827A1 - Systeme de guidage de tangage - Google Patents

Systeme de guidage de tangage Download PDF

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Publication number
WO1991015827A1
WO1991015827A1 PCT/US1991/002067 US9102067W WO9115827A1 WO 1991015827 A1 WO1991015827 A1 WO 1991015827A1 US 9102067 W US9102067 W US 9102067W WO 9115827 A1 WO9115827 A1 WO 9115827A1
Authority
WO
WIPO (PCT)
Prior art keywords
signal
pitch
pitch angle
aircraft
providing
Prior art date
Application number
PCT/US1991/002067
Other languages
English (en)
Inventor
J. Howard Glover
Original Assignee
Sundstrand Data Control, Inc.
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Priority claimed from US07/505,214 external-priority patent/US5136518A/en
Application filed by Sundstrand Data Control, Inc. filed Critical Sundstrand Data Control, Inc.
Publication of WO1991015827A1 publication Critical patent/WO1991015827A1/fr

Links

Classifications

    • G—PHYSICS
    • G05—CONTROLLING; REGULATING
    • G05D—SYSTEMS FOR CONTROLLING OR REGULATING NON-ELECTRIC VARIABLES
    • G05D1/00—Control of position, course, altitude or attitude of land, water, air or space vehicles, e.g. using automatic pilots
    • G05D1/04—Control of altitude or depth
    • G05D1/06—Rate of change of altitude or depth
    • G05D1/0607—Rate of change of altitude or depth specially adapted for aircraft
    • G05D1/0615—Rate of change of altitude or depth specially adapted for aircraft to counteract a perturbation, e.g. gust of wind
    • G05D1/0623—Rate of change of altitude or depth specially adapted for aircraft to counteract a perturbation, e.g. gust of wind by acting on the pitch
    • G—PHYSICS
    • G05—CONTROLLING; REGULATING
    • G05D—SYSTEMS FOR CONTROLLING OR REGULATING NON-ELECTRIC VARIABLES
    • G05D1/00—Control of position, course, altitude or attitude of land, water, air or space vehicles, e.g. using automatic pilots
    • G05D1/04—Control of altitude or depth
    • G05D1/06—Rate of change of altitude or depth
    • G05D1/0607—Rate of change of altitude or depth specially adapted for aircraft
    • G05D1/0653—Rate of change of altitude or depth specially adapted for aircraft during a phase of take-off or landing
    • G05D1/0661—Rate of change of altitude or depth specially adapted for aircraft during a phase of take-off or landing specially adapted for take-off

Definitions

  • This invention relates generally to aircraft pitch guidance systems, and more particularly to wind shear recovery systems with stall protection.
  • Aircraft pitch guidance systems for guiding a pilot during a wind shear condition are known.
  • One such system is disclosed in United States Patent No. 4,347,572 to Berwick, Jr., et al.
  • the Berwick, Jr., et al. system utilizes various inputs including air mass derived signals such as the angle of attack signal derived from an angle of attack vane to provide the pilot of an aircraft with a guidance signal defining the maximum climb rate obtainable without stalling the aircraft.
  • air mass derived signals such as the angle of attack signal derived from an angle of attack vane
  • the pitch attitude of the aircraft is compared with a pitch reference signal to provide a pitch guidance command signal to the pilot that defines the optimum pitch angle required for maximum climb during a wind shear condition.
  • the pitch reference value is modulated by a stall warning discrete upon the occurrence of a stall warning to reduce the pitch guidance command signal during a stall warning to prevent the pilot from attaining a pitch attitude capable of causing a stall.
  • the pitch guidance command signal is again increased toward the pitch reference value to permit the pilot to increase the pitch atti ⁇ tude to the reference value.
  • the pitch reference value is also modulated as a function of altitude and descent rate to increase the pitch reference at low altitudes and high descent rates.
  • FIG. 1 is a functional block diagram of the pitch guidance system as deployed in an aircraft
  • FIG. 2 is a functional block diagram of the pitch guidance system according to the invention
  • FIG. 3 is a logical flow diagram illustrating the operation of the pitch angle reference modulator utilized in the pitch guidance system according to the invention.
  • FIG. 4 is a functional block diagram showing 0 additional inputs to the pitch guidance system of
  • FIG. 2 to optimize the pitch guidance signal for dif ⁇ ferent altitudes and descent rates.
  • the system according to the invention is 15 particularly suitable for providing guidance commands to the pilot of an aircraft to advise him of the op ⁇ timum pitch attitude required for recovery from a wind shear condition.
  • the system may be utilized in conjunction with a wind shear detector, for example,
  • pitch guidance control The primary requirement of pitch guidance control is that the system should assist the pilot in arresting the descent of the aircraft, while at the same time maintaining adequate stall margin.
  • the guidance system In ex ⁇ treme cases, where loss of aircraft energy due to wind shear is very large, the guidance system must adjust the pitch attitude such that the distribution of energy between potential energy (altitude) and kinetic energy (air speed or angle of attack) reflects the immediate requirements. For example, if the air ⁇ craft is very close to the terrain, then the guidance must emphasize potential energy (climb performance) , while if there is adequate terrain clearance, then the emphasis must be placed on maintaining kinetic energy (stall margin) .
  • the pitch guidance commands must not allow the aircraft to stall.
  • the nature of severe wind shear is such that it is very likely that severe turbulence will be encountered by the aircraft during the recovery maneuver.
  • One effect of this turbulence may be to cause the stall warning system to generate a stall warning even though the warning is caused by momentary increases in the measured angle of attack caused by the turbulence that can be tolerated. How ⁇ ever, a normal trained response of the pilot to a stall warning is to reduce the pitch attitude of the aircraft in order to reduce the -angle of attack below the stall warning threshold value.
  • a wind shear recovery guidance system 10 that provides a pitch guidance command signal to a pilot via an attitude director indicator 12 which may be an aural or visual indicator such as a gauge or a cathode ray tube.
  • the wind shear recovery guidance system 10 receives signals from a vertical gyro 14 that provides a signal repre- sentative of the pitch attitude of the aircraft, i.e., the angle of the longitudinal axis of the aircraft relative to the horizon.
  • wind shear recovery guidance system 10 Other inputs to the wind shear recovery guidance system 10 include a stall warning system 16, a flap position discrete 18 and other discretes 20 which may include signals represen ⁇ tative of the type of aircraft carrying the system and other aircraft parameters such as, for example, landing gear position.
  • the stall warning system 16 may be a conven- tional stall warning system, sometimes known as a
  • stick shaker that monitors various aircraft para ⁇ meters such as angle of attack and airspeed and pro ⁇ vides a warning to the pilot in the event of an impend ⁇ ing stall condition.
  • the stall warning may be communi- cated to the pilot in a variety of ways, one way being a physical vibration of the aircraft's control, hence the term "stick shaker”.
  • flap position and other aircraft parameters affect the optimum pitch angle required for maximum climb of the aircraft, and hence, signals representative of flap position and aircraft type discretes are provided to the wind shear recovery guidance system 10 from the flap position discrete 18 and the aircraft type discretes 20.
  • the aircraft type discretes define the type of aircraft that is carrying the system and the optimum pitch angles for maximum climb under various flight con- figurations.
  • the guidance system 10 includes a pitch reference value generator 21 that responds to the aircraft type discretes and 5 the flap position signal to provide a pitch reference value based on the type of aircraft and its flap posi ⁇ tion.
  • the pitch reference value generator may take various forms such as, for example, a look-up table that has stored therein pitch reference values for o various aircraft under various flap settings or a computation system that has stored therein a reference value for a particular aircraft and modifies this value as a function of flap setting.
  • the pitch guidance system 10 is illustrated 5 in greater detail in FIG. 2.
  • the pitch guidance system 10 provides a pitch guidance command signal to the attitude director indicator 12 based on the pitch attitude of the aircraft, the pitch reference value which, as stated above, is determined by various para- 0 meters including the performance characteristics of the aircraft as defined by the aircraft type discretes 20 and aircraft configuration parameters such as flap position as obtained from the flap position discrete 18.
  • the pitch reference value is operated on by a reference modulator 22 under the control of the stall warning discrete from the stall warning system 16.
  • the operation of the reference modulator 22 will be described in greater detail in conjunction with FIG. 3, but -briefly, the reference modulator 22 oper ⁇ ates on the pitch reference value to reduce the pitch reference value upon the occurrence of a stall warning.
  • the pitch reference value as modified by the stall warning discrete is applied to a comparator 24 that compares the present pitch attitude of the aircraft as obtained from the vertical gyro 14 with the modu ⁇ lated pitch reference value to provide a signal repre ⁇ sentative of the deviation of the present pitch atti ⁇ tude of the aircraft from the modulated pitch reference value.
  • High frequency components of the pitch attitude signal from the vertical gyro 14 are passed through a high pass filter 26 (having a time constant of, for example, on the order of 2 seconds) and an amplifier 28.
  • the amplified high frequency signals from the amplifier 28 are subtracted from the output of the comparator 24 by a summing junction 30 to provide a damped pitch error signal to a command limiter 32.
  • the command limiter 32 limits the amplitude of the signal from the summing junction 30 to a value selected to avoid over controlling the aircraft, for example, to a range of zero to +20 degrees.
  • the logical operation of the reference modu ⁇ lator 22 is illustrated in FIG. 3.
  • the reference modulator monitors the pitch attitude signal from the vertical gyro 14 in conjunction with the pitch refer ⁇ ence value and the stall warning discrete to adjust the pitch guidance command signal as required relative to the pitch reference value and the stall warning discrete. As is illustrated in FIG. 3, the system first determines whether or not there is a stall warn ⁇ ing. If there is no stall warning, the system deter- mines whether the pitch reference signal is less than the reference value.
  • the output of the refer ⁇ ence modulator is simply limited to the pitch reference value. If the pitch reference signal is less than the reference value, the pitch command is increased by a predetermined number of degrees, for example, two degrees. After a delay of a predetermined number of seconds, for example, two seconds, the process is repeated, beginning with a determination of whether a stall warning is present.
  • the pitch guidance command signal is reduced by a predeter- 0 mined number of degrees, for example, two degrees, and after the predetermined time delay, the presence of a stall warning is monitored. If the stall warning is still present, the pitch guidance command signal is again reduced, but if not, the pitch guidance com- and signal is increased until it reaches the pitch reference value.
  • the reference modulator 22 modulates a total pitch refer ⁇ ence signal that is obtained from a combination of Q signals rather than a fixed pitch reference signal as a function of whether or not a stall warning is pre ⁇ sent.
  • the system 50 of FIG. 4 utilizes one of two fixed reference pitch signals depending on whether the aircraft is in a takeoff or an approach mode.
  • a pitch reference source 52 provides a signal representative of a 15° pitch reference
  • a pitch reference source 54 provides a 13° pitch reference.
  • the takeoff or approach reference source is selectable by a switch Q 55 that selects either the takeoff pitch reference 52 or the approach pitch reference 54 depending upon whether the aircraft is in a takeoff or approach mode.
  • the selection may be made manually, or preferably automatically by switching circuitry that monitors 5 various aircraft flight parameters such as flap posi ⁇ tion, landing gear position, engine power, airspeed and other parameters to automatically determine the ode of operation of the aircraft.
  • the reference signal from one of the sources 52 and 54 is applied to a summing junction 56 that sums it with various biasing signals that are determined by the altitude of the aircraft above ground and the vertical speed of the aircraft.
  • the vertical speed of the aircraft is ob ⁇ tained from a vertical speed source 58 that provides a signal representative of the vertical speed of the aircraft.
  • the vertical speed is obtained from a source that is not dependent on air data to avoid errors induced by turbulence that often accom ⁇ panies wind shear.
  • the vertical speed source 58 may include an inertial navigation system that provides an inertially derived Z-velocity signal. In a wind shear condition, if an aircraft has a negative (descending) vertical speed, it is more susceptible to the effects of downdrafts and negative shears than an aircraft having a positive vertical speed. Consequently, the vertical speed is monitored, and if the vertical speed is less than +8 feet per second, the pitch reference is increased.
  • the pitch reference is decreased. This is accomplished by a pair of func- tion generators 60 and 62.
  • the function generator 60 is operative for low positive vertical speeds and negative vertical speeds and serves to increase the pitch reference for vertical speeds below +8 feet per second. For example, between +8 feet per second and -32 feet per second during the takeoff mode, the pitch reference is increased by +8°. In the approach mode, the pitch reference is increased from 0° at a vertical speed of -4 feet per second to +8° at a vertical speed of -44 feet per second.
  • the function generator 60 provides a pitch bias signal to a multiplier 64 that is a function of the vertical speed signal from the vertical speed source 58.
  • the function generator 62 monitors the vertical speed from the vertical speed source 58 and decreases the pitch angle reference by up to -4° at vertical speeds between +24 and +44 feet per second.
  • the reduction in the pitch reference under high vertical speed conditions serves to reduce nuisance commands when the aircraft has sufficient vertical speed, or kinetic energy, to with ⁇ stand a wind shear without increasing the pitch angle.
  • the pitch bias signal from the function generator 62 is applied to a summing junction 66 and combined with the output of the function generator 60 that is applied to the summing junction 66 via a multiplier circuit 64.
  • the combined bias signals from the summing junc ⁇ tion 66 are applied to the summing junction 56 to provide the total reference signal to the reference modulator 22 which operates as it does in FIG. 2.
  • the function generators 60 and 62 bias the total pitch reference signal applied to the reference modulator 22 as a function of the vertical speed of the aircraft by increasing the pitch reference at negative vertical speeds and decreasing it at positive vertical speeds.
  • the output of the function generator 60 is modulated by the multiplier 64 as a function of the altitude of the aircraft above the ground.
  • the modu ⁇ lation is accomplished by a function generator 68 that monitors the output of a radio altimeter 70 and reduces the effect of the function generator 60 as a function of radio altitude. For example, below alti ⁇ tudes of 100 feet above ground, the function generator 68 applies a gain signal having a magnitude of 1 to -li ⁇ the multiplier 64.
  • the full pitch bias signal generated by the function generator 60 is applied to the summing • junction 66 (via the multiplier 64) so that it may fully modulate the total pitch reference signal applied to the reference modulator 22.
  • the pitch angle may be increased by up to 8° over the basic pitch command signal to thereby trade kinetic energy for altitude.
  • the achievement of the maximum climb rate is not as criti ⁇ cal as it is below 100 feet, and it is desirable to conserve kinetic energy to avoid a potential stall situation.
  • the function generator 68 gradually reduces the gain signal from a factor of 1 at 100 feet of radio altitude to 0 at a radio altitude of 500 feet.

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  • Engineering & Computer Science (AREA)
  • Aviation & Aerospace Engineering (AREA)
  • Radar, Positioning & Navigation (AREA)
  • Remote Sensing (AREA)
  • Physics & Mathematics (AREA)
  • General Physics & Mathematics (AREA)
  • Automation & Control Theory (AREA)
  • Traffic Control Systems (AREA)

Abstract

Un système de guidage du tangage d'un avion utilise les informations de tangage dérivées de l'inertie, pour donner au pilote les informations définissant l'angle de tangage optimal pour une ascension maximum dans des conditions de cisaillement du vent. Le système emploie un modulateur de référence du tangage (22) qui reçoit une valeur discrète d'avertissement de décrochage provenant d'un dispositif avertisseur de décrochage, pour réduire l'angle de tangage indiqué lors de l'apparition d'un signal avertisseur de décrochage afin de réduire la possibilité de décrochage de l'avion dans des conditions de fonctionnement détériorées telles que des conditions de vent arrière ou de panne de moteur. Le système utilise les informations de tangage dérivées de l'inertie plutôt que les informations relatives à l'angle d'attaque et dérivées de la masse d'air, pour éviter les phénomènes transitoires dans le signal du détecteur d'angle d'attaque, et l'angle de tangage indiqué est modulé en fonction de l'altitude (70) et de la vitesse verticale (58) pour optimiser l'angle de tangage à des altitudes et des vitesses de descente différentes.
PCT/US1991/002067 1990-04-04 1991-03-26 Systeme de guidage de tangage WO1991015827A1 (fr)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
US505,214 1990-04-04
US07/505,214 US5136518A (en) 1988-04-11 1990-04-04 Pitch guidance system

Publications (1)

Publication Number Publication Date
WO1991015827A1 true WO1991015827A1 (fr) 1991-10-17

Family

ID=24009463

Family Applications (1)

Application Number Title Priority Date Filing Date
PCT/US1991/002067 WO1991015827A1 (fr) 1990-04-04 1991-03-26 Systeme de guidage de tangage

Country Status (3)

Country Link
EP (1) EP0480001A4 (fr)
CA (1) CA2063823A1 (fr)
WO (1) WO1991015827A1 (fr)

Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20220258878A1 (en) * 2021-02-18 2022-08-18 Airbus (Sas) Method and device for assisting an aircraft crew in recovering the aircraft from a stall situation

Citations (9)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3744309A (en) * 1970-07-20 1973-07-10 United Control Corp Pitch signal calculator for aircraft
US3967799A (en) * 1972-11-17 1976-07-06 Sundstrand Data Control, Inc. Head up display and pitch generator
US4485446A (en) * 1981-09-08 1984-11-27 The Boeing Company Aircraft lift control system with acceleration and attitude limiting
US4590475A (en) * 1983-10-24 1986-05-20 The United States Of America As Represented By The Secretary Of The Army Stall avoidance system for aircraft
US4763266A (en) * 1982-12-27 1988-08-09 The Boeing Company Aircraft flight command and display system
US4786905A (en) * 1986-03-26 1988-11-22 Sundstrand Data Control, Inc. Wind shear detector head-up display system
US4841448A (en) * 1987-09-01 1989-06-20 Flight Dynamics, Inc. Windshear flight recovery command system
US4891642A (en) * 1988-01-11 1990-01-02 Sundstrand Data Control, Inc. Wind shear detection system
US4947164A (en) * 1988-01-21 1990-08-07 Sundstrand Data Control, Inc. Flight path responsive aircraft wind shear alerting and warning system

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US924059A (en) * 1909-01-02 1909-06-08 Charles W Grobe Dust-pan attachment.
GB2134866B (en) * 1980-11-28 1985-06-19 Sundstrand Data Control Angle of attack based pitch generator and head up display
DE3578659D1 (de) * 1985-12-18 1990-08-16 Boeing Co Flugregel- und anzeigesystem fuer windscherung.
US4725811A (en) * 1986-02-13 1988-02-16 Sundstrand Data Control, Inc. Wind shear detection and alerting system
WO1989009955A1 (fr) * 1988-04-11 1989-10-19 Sundstrand Data Control, Inc. Systeme de guidage a compensation de cisaillement du vent a protection contre le decrochage

Patent Citations (9)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3744309A (en) * 1970-07-20 1973-07-10 United Control Corp Pitch signal calculator for aircraft
US3967799A (en) * 1972-11-17 1976-07-06 Sundstrand Data Control, Inc. Head up display and pitch generator
US4485446A (en) * 1981-09-08 1984-11-27 The Boeing Company Aircraft lift control system with acceleration and attitude limiting
US4763266A (en) * 1982-12-27 1988-08-09 The Boeing Company Aircraft flight command and display system
US4590475A (en) * 1983-10-24 1986-05-20 The United States Of America As Represented By The Secretary Of The Army Stall avoidance system for aircraft
US4786905A (en) * 1986-03-26 1988-11-22 Sundstrand Data Control, Inc. Wind shear detector head-up display system
US4841448A (en) * 1987-09-01 1989-06-20 Flight Dynamics, Inc. Windshear flight recovery command system
US4891642A (en) * 1988-01-11 1990-01-02 Sundstrand Data Control, Inc. Wind shear detection system
US4947164A (en) * 1988-01-21 1990-08-07 Sundstrand Data Control, Inc. Flight path responsive aircraft wind shear alerting and warning system

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Title
See also references of EP0480001A4 *

Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20220258878A1 (en) * 2021-02-18 2022-08-18 Airbus (Sas) Method and device for assisting an aircraft crew in recovering the aircraft from a stall situation

Also Published As

Publication number Publication date
EP0480001A4 (en) 1993-09-08
CA2063823A1 (fr) 1991-10-05
EP0480001A1 (fr) 1992-04-15

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