United States Patent (19)
`Gold et al.
`
`54 INERTIAL VELOCITY COMMAND SYSTEM
`
`(75) Inventors: Phillip J. Gold, Shelton; Donald L.
`Fogler, Jr., Milford, both of Conn.;
`James B. Dryfoos, Wallingford, Pa.
`
`73 Assignee: United Technologies Corporation,
`Hartford, Conn.
`(21) Appl. No.: 253,477
`(22) Filed:
`Jun. 3, 1994
`(51) Int. C.
`
`B64C 11/34; G05D 1/08;
`G05D 1/10
`52 U.S. Cl. ...................................... 244/76 R; 24.4/17.13;
`244/178; 24.4/181; 244/195; 364/434; 364/453
`58) Field of Search ............................... 244/76 R, 17.13,
`244/175, 178, 181, 182, 191, 195, 228;
`364/443, 453, 432, 434
`
`(56)
`
`References Cited
`U.S. PATENT DOCUMENTS
`3,995,144 11/1976 Johnson et al. .................... 235/615E
`4,032,759 6/1977 Danik .............
`... 235/50.25
`4,070,674
`1/1978 Buell et al. ............................. 364/453
`4,106,094 8/1978 Land ..............
`... 364/453
`4,173,784 11/1979 Heath et al. ...
`... 364/453
`4,232,313 ll/1980 Fleishman ............................... 364/453
`4,645,141
`2/1987 McElreath ............................ 244/1713
`
`
`
`US005553812A
`11) Patent Number:
`(45) Date of Patent:
`
`5,553,812
`Sep. 10, 1996
`
`4,924,400 5/1990 Post et al. ............................... 364/433
`5,001,646 3/1991 Caldwell et al. ....................... 364/434
`5,008,825 4/1991 Nadkarni et al. ....................... 364,434
`5,195,700 3/1993 Fogler, Jr. et al. ...
`... 244/17.2
`5,238,203 8/1993 Skonieczny et al. ................... 364/434
`Primary Examiner-Andres Kashnikow
`AG, Examiner-Virna Lissi Mojica
`Attorney, Agent, or Firm-Michael Grillo
`57)
`ABSTRACT
`A velocity command system is provided with a velocity
`stabilization mode wherein aircraft flight path referenced
`velocities are determined with respect to an inertial frame of
`reference, the flight path referenced velocities are held
`constant during pilot commanded yaw maneuvers so that the
`aircraft maintains a fixed inertial referenced flight path
`regardless of the pointing direction of the aircraft. Velocity
`control with respect to an inertial frame of reference is
`accomplished by controlling the aircraft flight path based on
`aircraft body referenced commanded lateral and longitudinal
`acceleration and based on aircraft body referenced lateral
`and longitudinal centrifugal acceleration. Operation in the
`velocity stabilization mode is provided in response to the
`manual activation of the velocity stabilization mode by the
`pilot, provided that the aircraft is already operating in the
`ground speed mode and the aircraft is not in a coordinated
`turn.
`
`12 Claims, 5 Drawing Sheets
`
`- 22
`
`AIRCRAFT POSITION/
`ATITUDE SENSORS
`
`DJI-1005
`IPR2023-01104
`
`

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`U.S. Patent
`U.S. Patent
`
`Sep. 10, 1996
`Sep. 10, 1996
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`Sheet 1 of 5
`Sheet 1 of 5
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`5,553,812
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`Sheet 2 of 5
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`5,553,812
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`U.S. Patent
`U.S. Patent
`fig. 3
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`Sep. 10, 1996
`Sep. 10, 1996
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`fig. 3
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`U.S. Patent
`U.S. Patent
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`Sep. 10, 1996
`Sep. 10, 1996
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`Sheet 3 of 5
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`U.S. Patent
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`Sep. 10, 1996
`Sep. 10, 1996
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`

`

`1.
`INERTIAL VELOCITY COMMAND SYSTEM
`
`5,553,812
`
`The Government has rights in this invention pursuant to
`Contract No. DAAJ09-91-C-AO04 awarded by the Depart
`ment of the Army.
`
`TECHNICAL FIELD
`The present invention relates to aircraft velocity com
`mand systems, and more particularly to a velocity command
`system having an inertial reference which allows a pilot to
`yaw the aircraft during operation in a velocity stabilization
`mode without changing the flight path of the aircraft with
`respect to the inertial reference.
`
`10
`
`BACKGROUND OF THE INVENTION
`In automatic flight control systems, it is well known to
`provide a velocity stabilization mode (velocity hold mode)
`wherein the automatic flight control system maintains the
`aircraft's body referenced velocities constant during opera
`tion in the velocity hold mode. This is particularly useful in
`a variety of aircraft operating conditions to reduce the pilot
`workload, maximize fuel efficiency, maintain aircraft posi
`tion with respect to a mission flight plan, and for a variety
`of other purposes.
`During operation in a velocity hold mode, there are a
`number of situations wherein a pilot may desire to yaw or
`turn the aircraft while maintaining the velocity hold mode.
`For example, the pilot may wish to yaw the aircraft to
`perform a sensor sweep, or for weapons firing. However,
`during operation in an aircraft body referenced velocity hold
`mode, the aircraft flight path will change in response to a
`pilot yaw input. This is illustrated in FIG. 1 wherein an
`aircraft 10 maintains a constant body referenced velocity
`both before 11 and after 11a a yaw maneuver, and therefore
`assumes a new flight path after the maneuver. During certain
`aircraft operations, and particularly during nap of the earth
`flight operations wherein the aircraft maintains an altitude
`close to the ground, a change in aircraft flight path in
`response to a yaw input may be undesirable. This is true
`when there are obstacles, e.g., buildings, trees, mountains
`etc., on either side of the flight path which the aircraft may
`encounter on the change in flight path.
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`DISCLOSURE OF INVENTION
`Objects of the invention include provision of an improved
`aircraft velocity command system having an inertial frame
`of reference.
`Another object of the present invention is to provide an
`aircraft velocity command system which allows the pilot,
`during operation in a velocity stabilization mode, to yaw the
`aircraft from side to side without changing the inertial flight
`path of the aircraft.
`55
`According to the present invention, a velocity command
`system is provided with a velocity stabilization mode
`wherein aircraft flight path referenced velocities are deter
`mined with respect to an inertial frame of reference, the
`flight path referenced velocities are held constant during
`pilot commanded yaw maneuvers so that the aircraft main
`tains a fixed inertial referenced flight path regardless of the
`pointing direction of the aircraft.
`In further accord with the present invention, velocity
`control with respect to an inertial frame of reference is
`accomplished by controlling the aircraft flight path based on
`aircraft body referenced commanded lateral and longitudinal
`
`60
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`2
`acceleration and based on aircraft body referenced lateral
`and longitudinal centrifugal acceleration.
`In still further accord with the present invention, opera
`tion in the velocity stabilization mode is provided in
`response to the manual activation of the velocity stabiliza
`tion mode by the pilot, provided that the aircraft is already
`operating in the ground speed mode and the aircraft is not in
`a coordinated turn.
`The inertial referenced velocity command system of the
`present invention allows a pilot to yaw the aircraft nose from
`side to side during operation in the velocity stabilization
`mode without changing the flight path of the aircraft with
`respect to an inertial frame of reference. Therefore, during
`restricted flight path operations, such as during nap of the
`earth flight operations, a pilot can turn the aircraft to perform
`a sensor Sweep or to aim weapons without changing the
`aircraft inertial flight path. Conventional velocity hold sys
`tems maintain aircraft body referenced velocity and do not
`maintain the aircraft's inertial flight path when yaw inputs
`are applied. It would require a significant amount of pilot
`workload in order to manually perform an inertial reference
`velocity hold task.
`The foregoing and other objects, features and advantages
`of the present invention will become more apparent in light
`of the following detailed description of exemplary embodi
`ments thereof as illustrated in the accompanying drawings.
`
`BRIEF DESCRIPTION OF THE DRAWINGS
`FIG. 1 is a diagram showing a change in vehicle flight
`path in response to a pilot yaw command during operation
`in a velocity hold mode of a vehicle having a body refer
`enced velocity hold control system;
`FIG. 2 is a diagram showing the response of a vehicle to
`a pilot yaw command during operation in a velocity stabi
`lization mode, the aircraft having an inertial referenced
`velocity command system of the present invent)on;
`FIG. 3 is a schematic block diagram of the velocity
`command system of the present invention in relation to an
`aircraft flight control system;
`FIG. 4 is a simplified schematic block diagram of the
`flight control system and velocity command system of FIG.
`3;
`FIG. 5 is a more detailed schematic block diagram of the
`velocity command system of FIGS. 3 and 4; and
`FIG. 6 is a schematic block diagram of an alternative
`embodiment of the velocity command system of FIG. 5.
`
`BEST MODE FOR CARRYING OUT THE
`INVENTION
`The inertial referenced velocity command system of the
`present invention is particularly well suited for allowing a
`pilot to yaw an aircraft from side to side while maintaining
`a specific or selected course, without changing the aircraft
`flight path with respect to an inertial frame of reference.
`Referring to FIG. 2, an aircraft 10 is shown operating in a
`velocity stabilization mode having a fixed velocity along a
`selected flight path as indicated by velocity vector 12. In
`response to a yaw command by a pilot during operation in
`a velocity stabilization mode, the aircraft 10a yaws while
`maintaining a constant flight path reference velocity 12a
`with respect to an inertial frame of reference.
`The velocity command system will be described with
`respect to a model following flight control system for a
`rotary winged aircraft; however, it will be understood by
`
`

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`those skilled in the art that the velocity stabilization control
`of the present invention is applicable to conventional control
`systems, and to both fixed and rotary winged aircraft.
`Referring to FIG. 3, a flight control system 21 includes a
`primary flight control system (PFCS) 22, an automatic flight
`control system (AFCS) 24 and a velocity command system
`(VCS) 25. The PFCS receives displacement command out
`put signals from a displacement collective stick 26 on lines
`27. The AFCS also receives collective stick discrete output
`signals on the lines 27. The PFCS and AFCS each receive
`the force output command signals of a four-axis sidearm
`controller 29 on lines 30 and sensed parameter signals from
`sensors 31 on lines 32. The pilot command signals on lines
`27 and 30 and the sensed parameter signals on lines 32 are
`shown consolidated within trunklines 33 and 34 in the PFCS
`and AFCS, respectively.
`The PFCS and AFCS each contain control modules for
`controlling the yaw, pitch, roll and lift axes of the aircraft.
`These modules are shown by blocks 35–38 for the PFCS and
`blocks 39-42 for the AFCS. The PFCS modules provide
`rotor command signals, and the AFCS modules provide
`conditioning and/or trimming of the PFCS rotor command
`signals. The PFCS and AFCS modules are interconnected
`through bus 43.
`The PFCS and AFCS use a model following algorithm in
`each control axis to provide rotor command signals on
`output lines 44 to a rotor mixing function 45 which com
`mands displacement of mechanical servos 46 and linkages
`47 to control the tip path plane of a main rotor 50. Addi
`tionally, the rotor mixing function 45 controls tail rotor
`servos 48 which control the thrust of a tail rotor 51 through
`linkages 49. The sensed parameter signals from sensors 31,
`on lines 32, provide the PFCS and AFCS with the aircraft's
`angular rate and attitude response to the main rotor and tail
`rotor command signals. Additionally, the sensors provide
`information related to velocity, altitude, acceleration, etc.,
`which information may or may not be used by the flight
`control system.
`The VCS25 receives the force output command signals of
`the sidearm controller 29 on lines 30 and sensed parameter
`signals from sensors 31 on lines 32. The pilot command
`signals on lines 30 and the sensed parameter signals on lines
`32 are shown consolidated within trunklines 59 in the VCS.
`The VCS contains control channel modules for providing
`pitch axis and roll axis velocity error signals to the AFCS.
`These modules are shown by blocks 60 and 61, respectively.
`The VCS modules are interconnected to the PFCS and AFCS
`modules through the bus 43. As described in greater detail
`hereinafter, during operation in the velocity stabilization
`mode, the VCS velocity error signals are fed back through
`PFCS proportional and integral paths to drive the error
`signals to Zero.
`FIG. 4 illustrates the functional interconnection of the
`VCS 25 with the PFCS 22 and the AFCS 24. FIG. 4 Will be
`described with respect to the pitch axis modules 36, 40, 60,
`e.g., flight control pitch attitude reference and VCS pitch
`axis velocity error signals; however, it will be understood by
`those skilled in the art that the functional interconnection of
`FIG. 4 is equally applicable to flight control roll attitude
`reference and VCS roll axis velocity error signals.
`The PFCS receives a pitch axis command signal on line
`70, provided through trunk lines 33 and lines 30, from the
`sidearm controller 29 (FIG. 3). In the present embodiment,
`the sidearm controller is a four-axis force stick in which yaw
`axis command signals are generated by the pilot's lateral
`twisting (left or right) of the sidearm controller, pitch axis
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`4
`command signals are generated by the pilot's pushing and
`pulling (front or back) of the sidearm controller, and roll axis
`command signals are generated by the pilot applying a left
`or right force to the sidearm controller. The pitch command
`signal is presented to the input of a pitch axis command
`model 72. In the command model 72, pilot commands are
`shaped to yield a desired rate and/or attitude response. The
`desired rate response is provided on a line 75 in a feedfor
`ward path to an inverse model 76 of the vehicle dynamics.
`The inverse model 76 provides a control command signal on
`a line 90 which represents the approximate rotor command
`necessary to achieve the desired pitch axis rate of change of
`the aircraft for each pilot commanded maneuver, and pro
`vides the primary control input to the rotor mixing function
`45.
`The desired rate response on the line 75 is also provided
`to a summing junction 78 in the PFCS, and to a Body to
`Euler Transformation 79 in the AFCS. The summing junc
`tion 78 compares the desired rate response on line 75 with
`the aircraft's actual pitch rate, received (from sensors 31,
`through lines 32 and trunk 33) as a sensed pitch rate signal
`on line 84. The output of the summingjunction 78 is a pitch
`rate error signal on a line 85. The rate error signal is applied
`to a proportional gain function 87 (to reduce the error), the
`output of which is provided on a line 89 to a summing
`junction 88. The summing junction 88 also receives the
`control command signal on line 90 from the inverse model
`76, and a pitch command modifying signal on a line 92 from
`a summing junction 95. The output of the summing junction
`88 is provided on a line 100, and presented through the
`PFCS output trunk lines 44 to the mixing function 45.
`The magnitude and rate of change of the pitch command
`modifying signal from the AFCS is a function of the aircraft
`pitch error. The pitch command modifying signal is a
`calculated value provided by a model following algorithm
`within the AFCS, based on the actual aircraft response to the
`rotor command signal. The pitch command modifying signal
`modifies the rotor command signal to achieve the pilot
`commanded pitch attitude. A more detailed description of
`the model following algorithms of the AFCS, and AFCS
`architecture, is given in commonly owned U.S. Pat. No.
`5,238,203, entitled "High Speed Turn Coordination For A
`Rotary Wing Aircraft', the disclosure of which is incorpo
`rated herein by reference.
`A pitch attitude feedback error signal is used to provide
`the pitch command modifying signal. The attitude feedback
`error signal is developed in the AFCS. The desired rate
`response is provided on the line 75 to the Body to Euler
`Transformation 79. The Transformation 79 transforms the
`desired rate response, which is in terms of aircraft body axes,
`to an inertial axes reference. The output of the transforma
`tion 79 is provided on a line 101 to an AFCS attitude
`command model 103. The attitude command model 103 is
`an integral function which converts the desired rate response
`to a desired pitch attitude signal on a line 105. The desired
`pitch attitude signal is provided to a summing junction 108,
`the other input of which is a pitch attitude signal on a line
`107, provided from sensors 31, throughlines 32 and trunk 34
`(FIG. 3). The output of the summing junction 108 is a pitch
`attitude feedback error signal on a line 109 in terms of
`inertial axes, which is provided to a Euler to Body Trans
`form 110 which transforms the pitch attitude feedback error
`signal from an inertial axes reference back to an aircraft
`body axes reference on a line 111. The operation of both
`Transform functions 79, 110 are described in greater detail
`in the aforementioned commonly owned U.S. Pat. No.
`5,238,203, the disclosure of which is incorporated herein by
`reference.
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`The pitch attitude feedback error signal on the line 111 is
`applied to feedback shaping circuitry 112 and thereafter via
`a proportional path containing a gain function 125 to the
`summing junction 95. The pitch attitude feedback signal is
`also applied to the summing junction 95 via an integral path
`containing an integral function 130. The output of the
`summingjunction 95 is the pitch command modifying signal
`on the line 92.
`The command model 72 provides a desired attitude
`response signal on a line 140 to a velocity model 143 in the
`VCS. Additionally, the command model for the yaw axis
`provides a desired yaw axis rate response, i.e., desired
`heading rate, on a line 145 to the velocity model 143. In
`order to hold the flight path referenced velocities constant,
`centrifugal acceleration terms caused by the pilot yawing the
`vehicle are included in the velocity model. As described in
`greater detail hereinafter with respect to FIG. 5, the velocity
`model uses the heading rate signal on the line 145 and the
`desired attitude response for the pitch axis and the yaw axis
`to develop centrifugal acceleration terms. The acceleration
`terms are combined with derived pilot on-axis acceleration
`commands, and the result is integrated to yield a desired
`commanded velocity signal on a line 147. The commanded
`velocity signal on the line 147 is compared with ground
`reference velocity signals provided on a line 150 (from
`sensors 31 through lines 32 and trunk 33) in a summing
`junction 152. The output of the summing junction 152 is a
`velocity error signal which is fed back via the feedback
`shaping circuitry 122 through the proportional 125 and
`integral 130 paths to drive the velocity errors to zero.
`The feedback shaping circuitry 112 may be of any suitable
`type known in the art for providing selection and/or transi
`tion between attitude feedback signals and velocity error
`signals as applicable. For example, when the velocity sta
`bilization mode is not activated, the feedback shaping cir
`cuitry 112 provides the attitude feedback signals to the
`integral and proportional paths. However, upon activation of
`the velocity stabilization mode, the feedback shaping cir
`cuitry replaces the attitude feedback signals with the veloc
`ity error signals for the integral path and provides a com
`40
`bination of attitude feedback signals and velocity error
`signals to the proportional path. The above description of the
`feedback shaping circuitry 112 is provided for illustrative
`purposes only, and the circuitry will vary depending on the
`dynamics and characteristics of the plant being controlled.
`As will be apparent to those skilled in the art, transient free
`switching may be used to transition between attitude feed
`back signals and velocity error signals.
`Referring now to FIG. 5, the velocity command system 25
`for the pitch axis and the roll axis are shown in greater detail.
`Upon activation of the velocity stabilization mode, the
`command model 72 (FIG. 4) provides a desired rate
`response 75 and a desired attitude response 140 for the pitch
`and roll attitude axes so that the helicopter maintains an
`attitude necessary to maintain a constant or fixed desired
`velocity. The velocity command system 25 of the present
`invention is operative to maintain the desired velocity con
`stant with respect to an inertial frame of reference.
`In a pitch axis module 500, the desired pitch axis attitude
`response is provided on a line 140 to a gain function 501
`wherein the commanded attitude is multiplied by the gravi
`tational constant, g, i.e., 9.81 m/sec or 32.2 ft/sec. The
`output of the multiplication function 501 is an acceleration
`term provided on a line 503 to a summing junction 505. As
`will be understood by those skilled in the art, a pitch axis or
`roll axis command for a rotary wing aircraft is proportional
`to an acceleration command, i.e., by multiplying the axis
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`command by the gravitational constant. For a fixed wing
`aircraft (or other type of vehicle) an explicit acceleration
`command may be used to provide the pitch axis acceleration
`term on the line 503.
`The other input to the summing junction is a centrifugal
`acceleration term which will be described in greater detail
`hereinafter. The output of the summing junction 505 is an
`acceleration term which is provided on a line 510 to an
`integrator 520, the output of which is a desired pitch axis
`commanded velocity signal on the line 147. The desired
`commanded velocity signal is provided to the summing
`junction 152 the other input of which is the longitudinal
`ground speed signal on the line 150. The summing junction
`152 compares the commanded velocity with the actual
`velocity, and provides as its output a pitch axis velocity error
`signal on the line 155 which is provided via the PFCS
`integral path and proportional path to drive the velocity error
`tO Zer0.
`The roll axis velocity error is developed in a roll axis
`module 502 in the same way that the pitch axis velocity error
`is developed. The desired roll axis attitude response is
`provided on a line 140a to a gain function 521 wherein the
`attitude command is multiplied by the gravitational constant,
`g. The output of the gain function 521 is provided on a line
`523 to a summing junction 525. The other input to the
`summing junction 525 is a centrifugal acceleration term
`which will be described in greater detail hereinafter. The
`output of the summingjunction 525 is provided on a line 530
`to an integrator 540. The output of the integrator 540 is a
`desired roll axis commanded velocity signal on a line 147a
`which is provided to a summing junction 152a where it is
`compared with a lateral ground speed signal on a line 150a.
`The output of the Summing junction 152a is a roll axis
`velocity error signal on a line 155a.
`In order to hold the flight path (inertial) referenced
`velocities constant, centrifugal acceleration terms which are
`caused by the pilot yawing the vehicle must be included in
`the pitch axis and roll axis velocity models. A longitudinal
`centrifugal acceleration term is provided on a line 550 to the
`summing junction 505. The longitudinal centrifugal accel
`eration term is determined by a multiplication function 555
`as the product of the commanded heading rate on line 145
`and the desired roll axis commanded velocity signal on the
`line 147a. The output of the multiplication function 555 is
`provided via normally closed switch 557 to the summing
`junction 505 via the line 550. The normally closed switch
`557 opens in response to the aircraft operating in the
`automatic turn coordination mode. The lateral centrifugal
`acceleration term is provided as the output of multiplication
`function 565 via normally closed switch 567 and line 560 to
`the summing junction 525. The lateral centrifugal accelera
`tion term is determined by the multiplication function 565 as
`the product of the commanded heading rate on the line 140
`and the desired pitch axis velocity command signal on the
`line 147. If the aircraft is in a coordinated turn, switches 557
`and 657 are opened, and the centrifugal acceleration terms
`are set equal to zero.
`The velocity stabilization described thus far is applicable
`to an aircraft flight control system implemented using model
`following algorithms. However, the inertial reference veloc
`ity command system of the present invention may be imple
`mented in a conventional flight control system using the
`simplified coordinate system transformations illustrated in
`FIG. 6. Referring to FIG. 6, aircraft body reference veloci
`ties (velocity vectors Vx and Vy) are converted to inertial
`reference velocities (velocity vectors Vn and Ve) using a
`body to inertial transformation function 600. The body to
`
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`7
`inertial transformation 600 uses equations 1 and 2 below to
`transform body referenced velocities to inertial referenced
`velocities:
`
`5,553,812
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`V=V, cos ()-V, sin ()
`
`(eq. 1)
`
`(eq. 2)
`V=V. sin (Y)-V, cos ('')
`wherein Vn is aircraft inertial referenced longitudinal veloc
`ity, e.g., in the North/South direction; Ve is aircraft inertial
`referenced lateral velocity, e.g., in the East/West direction;
`Vx is aircraft body referenced longitudinal velocity; Vy is
`aircraft body referenced lateral velocity; and I is vehicle
`heading.
`Upon activation of the inertial reference velocity synchro
`nization mode, synchronization functions 603 and 604 store
`the instantaneous values of the inertial referenced velocity
`vectors. During operation in the inertial referenced velocity
`synchronization mode, the synchronization functions 603
`and 604 compare the aircraft's actual inertial referenced
`velocity to the stored inertial referenced velocity values to
`develop inertial referenced velocity error signals, dVn and
`dVe. Thereafter, the inertial referenced velocity error signals
`are provided to an inertial to body transformation function
`610 which converts the error signals from an inertial refer
`ence back to a body reference. The inertial to body trans
`formation function 610 uses equations 3 and 4 below;
`
`dV=dV, cos ()+dV. sin (F)
`
`(eq. 3)
`
`(eq. 4)
`dV=-dV, sin (Y)+dV. cos (Y)
`wherein dVn is aircraft inertial referenced longitudinal
`velocity error; dVe is aircraft inertial referenced lateral
`velocity error; dVx is aircraft body referenced longitudinal
`velocity error, and dVy is aircraft body referenced lateral
`velocity error.
`Thereafter, the conventional control system provides the
`body referenced error signals via feedback paths, which may
`or may not contain proportional and integral paths, to drive
`the velocity errors to zero.
`The invention is described herein as being responsive to
`command signals provided by a four-axis side arm controller
`used with a model following control system. However, the
`invention will work equally as well with a flight control
`system wherein control surface commands are provided by
`conventional control inputs such as a mechanical displace
`ment control stick and pilot operated pedals. Alternatively,
`other combinations of mechanical, electro-mechanical and
`electronic control may be used. It is also anticipated that the
`invention would be equally applicable to remotely piloted
`vehicles.
`Although the centrifugal acceleration term (on lines 550
`and 560 of FIG. 5) are shown and described as being
`developed in response to commanded heading rate, it is
`expected that actual heading rate may also be used to
`develop the centrifugal acceleration terms. However, it may
`be desirable when using actual heading rate to provide
`filtering to reduce the effects of wind gusts and external
`forces on measured heading rate.
`Although the invention has been described and illustrated
`with respect to exemplary embodiments thereof, it should be
`understood by those skilled in the art that the foregoing and
`various other changes, omissions and additions may be
`made therein and thereto without departing from the spirit
`and scope of the present invention.
`We claim:
`1. An aircraft flight control system for providing control
`surface command signals to the aircraft control surfaces
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`thereby controlling the yaw, pitch, roll and lift attitude axes
`of the aircraft in flight, comprising:
`velocity stabilization means for providing a velocity sta
`bilization signal in response to the activation of a
`velocity stabilization mode; and
`velocity command means responsive to the presence of
`said velocity stabilization signal for maintaining a
`constant aircraft velocity with respect to an inertial
`reference.
`2. An aircraft flight control system according to claim 1
`wherein said velocity command means further comprises:
`means for providing reference signals corresponding to a
`desired constant aircraft velocity with respect to an
`inertial reference;
`means for providing inertial velocity signals indicative of
`aircraft velocity with respect to said inertial reference;
`means responsive to said reference signals and said iner
`tial velocity signals for providing inertial velocity error
`signals indicative of the difference there between; and
`said flight control system providing control surface com
`mand signals to drive the magnitude of said inertial
`velocity error signals to zero when said velocity stabi
`lization signal is present.
`3. An aircraft flight control system according to claim 1
`wherein said velocity command means further comprises:
`means for providing reference signals corresponding to a
`desired constant aircraft velocity with respect to an
`inertial reference;
`means for providing aircraft velocity signals indicative of
`aircraft velocity with respect to an aircraft frame of
`reference;
`means for converting said aircraft velocity signals to
`inertial velocity signals indicative of aircraft velocity
`with respect to said inertial reference;
`means responsive to said reference signals and said iner
`tial velocity signals for providing inertial velocity error
`signals indicative of the difference there between;
`means for converting said inertial velocity error signals to
`aircraft velocity error signals indicative of the required
`change in aircraft velocity for the aircraft to operate at
`said constant velocity with respect to an inertial refer
`ence; and
`said flight control system providing control surface com
`mand signals to drive the magnitude of said aircraft
`velocity error signals to zero when said velocity stabi
`lization signal is present.
`4. An aircraft flight control system according to claim 1
`wherein said velocity command means further comprises:
`velocity error means for providing velocity error signals
`for the pitch axis and the roll axis indicative of the
`aircraft attitude rate of change for the pitch axis and the
`roll axis, respectively, required to maintain said con
`stant aircraft velocity with respect to an inertial refer
`ence; and
`said flight control system being responsive to the presence
`of said velocity stabilization signal for providing con
`trol surface command signals to drive the magnitude of
`said velocity error signals to zero.
`5. An aircraft flight control system according to claim 4
`wherein said means for providing velocity error signals
`further comprises:
`control means operable by a pilot to provide axis com
`mand signals for controlling the yaw, pitch, roll and lift
`attitude axes of the aircraft;
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`means responsive to said axis command signals for pro
`viding corresponding desired attitude signals indicative
`of a desired aircraft attitude in the yaw, pitch