e
`
`United States Patent (19)
`Caldwell et al.
`
`5,001,646
`Patent Number:
`(11
`45 Date of Patent: Mar. 19, 1991
`
`(56)
`
`4,563,743 1/1986 Murphy et al. ..................... 318/584
`54 AUTOMATED HELICOPTER FLIGHT
`4,564,908 1/1986 Clelford et al...
`. 364/433
`CONTROL SYSTEM
`4,617,662 10/1986 Auer, Jr. et al. ..................... 37/25
`4,648,345 3/1987 Whan et al. ........................ 114/338
`75 Inventors: Donald G. Caldwell, Mesa; Stephen S.
`4,679, 51 7/1987 Morris et al. ....................... 364/510
`Osder, Scottsdale, both of Ariz.
`4,697,768 10/1987 Klein ................................... 364/434
`4,713,757 12/1987 Davidson et al. ................... 364/200
`(73) Assignee: McDonnell Douglas Corporation,
`4,801,110 l/1989 Skutecki........................... 244/7.13
`Long Beach, Calif.
`Primary Examiner-Parshotam S. Lall
`21 Appl. No.: 286,235
`Assistant Examiner-Michael Zanelli
`(22
`Filed:
`Dec. 19, 1988
`Attorney, Agent, or Firm-Curt L. Harrington; George
`W. Finch; John P. Scholl
`51 Int. Cl........................ B64C13/04; B64C13/16
`52 U.S. C. ................................ 364/434; 364/424.01
`57
`ABSTRACT
`58 Field of Search ............... 364/433,434,452,449;
`An apparatus and method for controlling the flight of a
`244/17.11, 17.13, 17.25, 237, 178, 192, 179, 180,
`helicopter, in which the pilot commands vehicle state
`191, 175, 177; 318/584-586
`changes, embodies several significant advances to the .
`References Cited
`helicopter control art. Among these are control decou
`pling to reduce undesired helicopter responses in other
`U.S. PATENT DOCUMENTS
`axes to pilot inputs in one axis, task, oriented control
`3,382,283 5/1983 Clelford et al. ..................... 364/434
`modes which are selected automatically, and automatic
`3,711,042 1/1973 Rempfer et al. ...
`... 244/9
`control of engine torque and rotor speed constraints.
`4,116,405 9/1978 Bacchi et al. ......
`. 24.4/12.4
`Novel task oriented control modes include independent
`4,150,686 4/1979 El Sherif et al.
`... 137/377
`control of heading and ground track at low speed and
`4,243,358 1/1981 Carlock et al. ....
`6/11
`2.
`MEly et al. .
`& : turn rate and flight path angle control at higher speeds.
`as was
`lay ...
`Control decoupling is achieved through a combination
`4,313,201
`/ 1982. Fischer et al. ..
`... 37/68
`4,330,829 Sv.982 Fischer et al. ...
`... 364/434
`of force/moment decoupling and open loop prediction
`235,336 10/1982 Murphy et al. .
`3/200
`methods. Vertical control, traditionally accomplished
`4,355,358 10/982 Clelford et al.
`... 364/424
`with collective only, is accomplished in this system
`4,362,085 12/1982 Venuti, Jr. ......
`... 9/86 with a combination of longitudinal cyclic and collective
`4,382,281 5/1983 Fowler et al. ..
`... 364/424
`control, increasing vertical response performance and
`4,412,280 10/1988 Murphy et al. ...
`... 364/200
`minimizing engine power change demands. These flight
`12. 3. E. et al
`y
`control automation methods were designed to enable
`i-ari
`s
`urphy et al. ....
`ww.
`4,443,853 4/1984 Maciolek et al. ..
`... 364/434
`the use of small displacement, multi aXS COtto lers for
`full envelope helicopter control; the system provides
`4,484,283 1/1984 Verzella et al. ...
`... 364/434
`dramatic pilot work load reduction for any of wide
`4,500,967 2/1985 Murphy et al.
`364/434
`3/86 range of pilot manipulator configurations.
`4,517,639 5/1985 Ferrietal.".
`4,542,506 9/1985 Oe et al. .................................. 371/9
`4,542,679 9/1985 Murphy et al.
`9/363
`21 Claims, 10 Drawing Sheets
`Aspe A
`
`WERICA
`PERFORMANCE
`constraN
`
`PRECOR
`ACCEERAION
`AE
`24.
`HEAONG
`
`ATITUDE
`MEASUREMENT
`SENSORS
`ATITLE
`PREDICTOR
`
`29
`
`RM
`
`STICK
`NP
`
`GRACEN .
`MIT
`ONTROL
`
`NEAR AccE 25
`DYNAMIC ".
`MODEL
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`27
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`33
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`39
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`ATITUDE
`REFERENCE
`GE)--G)
`4.
`43
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`ERROR
`RAE
`45 cgo
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`ACCELERATON
`SERE
`PRECTOR
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`35E to E. E."
`49
`ATTITUE
`WIBRATION
`FILER
`RATE
`CLECWE
`
`VELOCIE COMMAND
`FROM PSION HO
`FERE CONTROL SYSTEM
`
`WLOCITY
`
`WLOCITY ESTIMA
`
`37
`
`LONGINAL
`CYCLIC
`
`
`
`LATERA
`CYCC
`DIRECTONA
`
`FORCE
`MOMEN
`OECOUPLENG
`
`5.
`(2)
`
`ATITUGE RAT
`Eff
`
`sE,
`N
`CONTRO
`53
`
`MODE
`CONTROL
`
`DJI-1014
`IPR2023-01104
`
`

`

`U.S. Patent
`
`Mar. 19, 1991
`
`Sheet 1 of 10
`
`5,001,646
`
`
`
`VERTICAL
`
`FIG. 1
`
`

`

`U.S. Patent
`
`Mar.19, 1991
`
`Sheet 2 of 10
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`U.S. Patent
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`Mar.19, 1991
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`U.S. Patent
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`Mar. 19, 1991
`
`Sheet 4 of 10
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`5,001,646
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`U.S. Patent
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`Mar.19, 1991
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`Sheet 5 of 10
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`5,001,646
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`5,001,646
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`U.S. Patent
`
`Mar.19, 1991
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`Sheet 7 of 10
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`5,001,646
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`U.S. Patent
`
`Mar. 19, 1991
`
`Sheet 8 of 10
`
`5,001,646
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`U.S. Patent
`
`Mar.19, 1991
`
`Sheet 9 of 10
`
`5,001,646
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`U.S. Patent
`
`Mar. 19, 1991
`
`Sheet 10 of 10
`
`5,001,646
`
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`1.
`
`w
`
`AUTOMATED HELICOPTER FLIGHT CONTROL
`SYSTEM
`
`SO
`
`5
`
`30
`
`BACKGROUND OF THE INVENTION
`Helicopter type aircraft are known to be aerodynami
`cally unstable. The piloting process for such vehicles
`involves the active use of both hands and both feet, as
`well as close attention to the aircraft attitude. The pro
`10
`cess is especially complicated due to the interrelated
`nature of the controls. An increase in the pitch of the
`main rotor causes a twisting of the aircraft which must
`be corrected for by an adjustment of the tail rotor. In a
`similar manner, when the pilot desires to perform a
`single action with the aircraft, he must simultaneously
`15
`perform a complex, learned series of actions on the
`controls in order to achieve the single action.
`In addition, helicopter type aircraft typically are able
`to operate closer to the limit of their material strength
`and limit of their flight envelope. An overstressing of
`20
`the aircraft can cause mechanical damage. An over
`specification of a number of flight characteristics can
`also push the aircraft beyond its limit, or cause none of
`the characteristics specified to be achieved.
`Although there are numerous examples of advance
`25
`ment of the helicopter control art, each typically deals
`with a narrow problem and its concomitant narrow
`solution, such as an engine torque limitation which
`prevents demands on the engine in excess of the torque
`limit. One reference, U.S. Pat. No. 4,420,808, E. D.
`Diamond, et al, inventors, discloses the use of a four axis
`force stick and a trim controller. The 808 patent recites
`that it has the ability to provide the capability for a pilot
`to control an aircraft in response to his perception of
`changes in attitude, altitude, speed, heading and the like,
`35
`with control inputs provided by the pilot only in the
`event that a change in aircraft response is required. The
`system of the 808' patent, then, is a trim system which
`sets the controls of a helicopter to their last position.
`This method does not take into account changing condi
`40
`tions such as wind shifts, or aircraft attitude. Under the
`system of the 808 patent, it is still the responsibility of.
`the pilot to make the minute control changes necessary
`to keep the aircraft stable, or steady with respect to a
`fixed point on the ground. The system of the 808 patent
`45
`was intended primarily to increase cockpit usable space
`by integrating the controls into a four axis stick with
`some form of control.
`SUMMARY OF THE INVENTION
`The present invention is a fully integrated fly-by-wire
`system for a helicopter utilizing a ground and inertial
`reference system and automated stability control to
`enable vectored directional control of the aircraft The
`system of the present invention corrects for external
`55
`changes and uses feedforward and feedback prediction
`to better implement pilot commands. Acceleration com
`mands enable the system to have greater compatibility
`with basic attitude stabilization loops currently in use.
`Control signals are reduced to their force and no
`ment components and then used according to a
`weighting matrix. All commands are continually modi
`fied by constraint feedbacks to thereby adapt limits of
`torque, rotor RPM and actuator authority performance
`envelopes. The built-in synchronization structures
`65
`allow transient free transition into and out of differing
`control modes. A vertical flight path energy manage
`ment system takes vertical commands to produce verti
`
`5,001,646
`2
`cal accelerations, using kinetic to potential energy con
`versions plus power adjustments for larger maneuver
`requirements.
`The system of the present invention is designed to
`command computed attitude maneuvers which maxi
`mize the response obtainable to an acceleration com
`mand. This is especially important in helicopter aircraft
`where the aircraft attitude is a major determiner of the
`speed and direction of the maneuver. All control and
`stabilization functions are decoupled so that the pilot is
`not required to apply compensating controls in any axis
`orthogonal to the axis in which commands are inserted.
`BRIEF DESCRIPTION OF THE DRAWINGS
`The structure and method of operation of the inven
`tion, together with additional advantages thereof, will
`be best understood from the following description of
`specific embodiments when read in connection with the
`accompanying drawings, in which:
`FIG. 1 is an isometric drawing of the control stick
`which can be utilized to affect command inputs to the
`system of the present invention;
`FIG. 2 is a conceptual block diagram of one axis of
`the acceleration command control system of the present
`invention driven by the control stick of FIG. 1;
`FIG. 3 is a block diagram of the pitch control block
`incorporated into the conceptual block diagram of FIG.
`2;
`FIG. 4 is a block diagram of the pitch stabilization
`control block incorporated into the conceptual block
`diagram of FIG. 2;
`FIG. 5 is a block diagram of the lateral maneuver
`control block incorporated into the conceptual block
`diagram of FIG. 2;
`FIG. 6 is a block diagram of the roll stabilization
`control block incorporated into the conceptual block
`diagram of FIG. 2;
`FIG. 7 is a block diagram of the directional stabiliza
`tion control block;
`FIG. 8 is a block diagram of the vertical maneuvering
`command and stabilization control block;
`FIG. 9 is a block diagram of the hover position con
`trol block, without hover sensor, incorporated into the
`conceptual block diagram of FIG. 2.; and,
`FIGS. 10a and 10b are block diagrams the hover
`position control block, with hover sensor, incorporated
`into the conceptual block diagram of FIG. 2.
`DETAILED DESCRIPTION OF THE
`PREFERRED EMBOOMENTS
`Referring to FIG. 1, a control stick 11 is movably
`supported by a support 13. The indicated control move
`ments shown are for vertical, yaw, pitch, and roll, and
`are indicated by the appropriate directional arrows.
`Referring to FIG. 2, a conceptual control path block
`diagram for one axis of the system of the present inven
`tion is shown. Each of the four control axes is a specific
`realization of this more general figure. It is understood
`that the diagrammatic representations illustrated herein
`may be implemented in a digital or analog fashion Stick
`inputs from control stick 11 representing vector direc
`tion command signals are connected to a gradient and
`limit control block 21. Gradient and limit control block
`21 provides limits and modifications to signals received
`from control stick 11 because of the finite performance
`limits inherent in the control surfaces of a helicopter. In
`addition to the finite structural limitations which are
`
`

`

`10
`
`5
`
`5,001,646
`4.
`3
`The attitude reference signal from summing junction
`constant, a set of limitations due to present operating
`41 is introduced into a summing junction 43 where it is
`characteristics, known as performance constraints, are
`sumned with a signal which is the negative of the atti
`also input into gradient and limit control block 21.
`The output from gradient and limit control block 21
`tude from the attitude measurement sensors. This atti
`tude measurement signal is generated from attitude
`is connected to a non-linear dynamic model block 23. In
`measurement sensors aboard the aircraft (not shown),
`this block, the now limited vector direction command
`and is normally provided by standard aircraft attitude
`signal from the control stick 11 is used to generate three
`instrumentation. The output of summing junction 43 is
`command components. An attitude rate predictor out
`the attitude error signal.
`put from block 23 is a signal indicative of how rapidly
`The attitude error signal from summing junction 43 is
`the attitude of the aircraft has been commanded to
`change. An acceleration attitude predictor output from
`fed into a rate command gain block 45. The gains of the
`block 23 is a signal indicative of the expected attitude
`rate command gain block 45 can be varied to affect the
`desired level of aircraft performance.
`change required of the aircraft to achieve the com
`manded acceleration. The third output from block 23 is
`A vibration filter block 49 receives an attitude rate
`measurement signal. The attitude rate measurement
`an acceleration command signal
`signal is also normally provided by standard aircraft
`The acceleration command signal from block 23 is
`input into a heading transformation block 24. The head
`attitude instrumentation. The vibration filter block 49
`smooths the attitude rate measurement signal to rid it of
`ing transformation block 24 resolves aircraft body axis
`unwanted vibration to prevent overcontrol. The output
`acceleration commands into an earth referenced coordi
`of the vibration filter block 49 is negatively summed
`nate frame.
`20
`An integrator 25 receives the output from the head
`with the output of the rate command gain block 45 and
`with the attitude rate predictor signal from dynamic
`ing transformation block 24 and performs an integration
`model block 23 in a summing junction 51.
`to calculate a velocity command signal. The velocity
`command signal from integrator 25 is inputtably con
`The output of summing junction 51 is fed to an atti
`nected to a second coordinate transform block 27,
`tude stabilization control block 53. A mode control
`25
`where the velocity command signal is transformed back
`input to attitude stabilization control block 53 selects
`which control laws are to treat the signal from summing
`into the aircraft body axis coordinate frame. These
`blocks 24, 25, and 27 implement one of the very unique
`junction 51. The control laws embodied within attitude
`features of the present invention, in that stick com
`stabilization control law block 53 keep track of the
`physical limits of the aircraft to the extent to which
`mands generate acceleration commands in the aircraft
`30
`coordinate frame; but the resulting velocity command is
`control can be achieved.
`The output of the attitude stabilization control law
`maintained in an earth referenced frame.
`The velocity command signal is first picked off from
`block 53 is fed into a decoupling block 55. Within de
`coupling block 55 a decoupling matrix operation is
`heading transformation block 27 and inputtably con
`performed to calculate the extent to which each output
`nected to a trim table block 29.
`35
`must be actuated in order to effect the stick inputs. The
`Trim table block 29 estimates the aircraft attitudes
`required to maintain the commanded velocity. The
`values used in the decoupling matrix at any one time are
`processing of the velocity command signal by trim table
`a function of the degree of interaction among the con
`block 29 produces an attitude predict signal.
`trolled outputs, as well as the speed and attitude of the
`aircraft. The outputs of the decoupling block 55 relate
`Beneath coordinate transform block 27 is a velocity
`to the traditional helicopter aircraft inputs of collective
`commands from position hold and fire control system
`block. This block provides a steadying input to the
`control, longitudinal cyclic control, lateral cyclic con
`system based upon the last command of the position
`trol, and directional control. Typically, the helicopter
`hold and fire control system.
`engine's fuel response to increased collective will be
`Velocity sensors on the aircraft comprise a velocity
`handled in a sub control loop not treated here.
`45
`It is well known in the helicopter control field that a
`measurement block 35. The measured velocity from
`velocity measurement block 35 is inputtably connected
`movement of one control, such as collective, not only
`produces an aircraft response in the desired axis, such as
`to a velocity estimate filter block 37 in order to reduce
`vertical, but also produces aircraft responses in other
`any noise or vibration variations in the measured veloc
`ity. The negative summing of the output of the velocity
`axes, such as yaw, pitch, and roll. In conventional heli
`SO
`copter control systems, the pilot must coordinate all
`estimate filter block 37 with a total velocity command
`signal creates a velocity error signal.
`four axes of control to produce a desired single axis
`response. A feature of the present invention performs
`The velocity command from block 27 is also summed
`that coordination automatically, in large part through
`at summing junction 33 with velocity commands from a
`position hold computation and fire control system block
`the decoupling block 55, but also through other control
`55
`31 and the output of the velocity estimate filter block 37
`feedback elements. Force and moment decoupling is
`achieved in decoupling block 55 by apportioning each
`to arrive at the total velocity error command.
`The velocity error signal from summing junction 33
`single input to block 55 into all four control surface
`is connected to the velocity control block 39 which is a
`command outputs. Each output from block 55 drives an
`proportional plus integral control element. The output
`actuator, which may be of the electrohydraulic or other
`type within the scope of the current art. The actuator, in
`of velocity control block 39 is an attitude command
`signal representing the attitude change required to re
`turn, moves the control surface to which it is attached,
`thereby effecting the commanded helicopter motion.
`duce the velocity error towards zero.
`The output from trim tables block 29, the accellera
`FIGS. 1 and 2 illustrate the overall scheme of the
`present invention. The remaining figures are more de
`tion attitude predictor from non-linear dynamic model
`tailed control logic diagrams illustrating an intermedi
`block 23, and the output of velocity control block 39 are
`ate level of control complexity and centering on the
`summed in a summing junction 41 to form an attitude
`reference signal.
`general types of control each subscheme is to control. It
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`The longitudinal acceleration command output of
`is understood that each level of control logic is imple
`acceleration limit block 111 is fed into a body orienta
`mentable either in analog or digital computer form.
`tion block 115. Body orientation block 115 also receives
`Referring to FIG. 1, the pitch axis is separate and
`distinct from the other axes pictured. Although the pilot
`a lateral acceleration command which will be shown
`later on FIG. 5. The operation carried on within body
`may activate more than one axis at a time, each axis has
`orientation block 115 is shown in the lower left half of
`its own distinct control limitation curve. Referring to
`FIG.3 as a heading rotation matrix to rotate the accel
`FIG. 3, the longitudinal control scheme, a stick shaping
`eration command into an earth referenced north/east
`block 21, similar to gradient and limit control block 21
`coordinate frame.
`of FIG. 2, is shown in the upper left hand corner of
`The outputs of the body orientation block 115 include
`FIG. 3.
`a north acceleration signal and an east acceleration
`The stickshaping block 21 shown in FIG.3 is specific
`signal The vertical north acceleration signal is fed into
`to the longitudinal control function. The graph in the
`an integrator 117. The output of integrator 117 is a
`stick shaping block 21 illustrates the signal output verti
`north velocity command which results from the inte
`cally as a function of stick position input horizontally.
`gration of the north acceleration signal The east accel
`15
`The LQ line represents the output limit of the stick. The
`eration signal is fed into an integrator 119. The output of
`DQ zone just to the right of the vertical axis represents
`integrator 119 is an east velocity command which re
`the stick deadzone, the range over which the stick may
`sults from the integration of the east acceleration signal
`be moved forward and still produce no output. The
`The east velocity command from integrator 119 and
`point BQ on the horizontal axis represents the transition
`the north velocity command from integrator 117 are
`20
`from a second order output between DQ and BQ to a
`rotated back into the body axis frame through an in
`linear output between BQ and LQ. At BQ, the slope of
`verse heading rotation matrix block 121. The outputs of
`the second order segment exactly equals the slope of the
`block 121 are the longitudinal velocity command and
`linear segment. The purpose of the stick shaping func
`the lateral velocity command.
`tion is to provide a low stick gain in the region of neu
`This rotation, integration, and inverse rotation
`25
`tral stick for precise controllability and a high stick gain
`method allows stick inputs to command body axis accel
`in the region of large stick forces for maximum aircraft
`erations, but the resultant velocity commands are main
`response. The deadzone region DO is needed to reject
`tained in the earth reference frame. For a helicopter this
`inadvertant pilot inputs which may occur when the
`provides a feature where the aircraft can be turned from
`pilot unintentionally applies a small force in an axis
`its direction of flight to fire at a target, for example,
`30
`orthogonal to the axis that he intends to command.
`without changing its flight path across the ground.
`The output of stick shaping block 21 is input into a
`Above and to the left of the inverse heading rotation
`feedback summing junction 101. Summing junction 101
`block 121 is the position computation block 123. Posi
`and blocks 103, 105, and 107 together form a rate lim
`tion computation block 123 receives data from any
`ited first order log filter on the shaped stick output of
`velocity or position currently available in the instru
`35
`block 21. Filter bandwidth block 103 controls the band
`mentation sensing existing art. Position computation
`width of this low pass filter while stick rate limit block
`block 123 outputs a longitudinal velocity command
`105 controls the maximum rate of change of the shaped
`which is summined with the longitudinal velocity com
`pitch stick output of filter bandwidth block 103. The
`mand from the fire control system block 31, from FIG.
`output of maneuver rate command limit block 105 is fed
`2, in a summing junction 125. The fire control system
`into integrator block 107. One output of stick rate limit
`block 31 was referred to in FIG. 2, and will not be
`block 105 will be subsequently applied to an input
`discussed in greater detail.
`shown later. An output from integrator 107 is also
`The longitudinal velocity is introduced near the bot
`picked off and applied to feedback summing junction
`tom right of FIG. 3 and is typically derived from stan
`101 to complete the filter implementation.
`dard inertial, doppler, or other sensors found on air
`The shaped pitch stick from the output of the integra
`craft. The longitudinal velocity measurement, the out
`tor 107 is also fed into a maneuver command scale fac
`put of summing junction 125, and the longitudinal ve
`tor block 109. The output from the maneuver command
`locity command velocity commands output of heading
`scale factor block 109 is fed into an acceleration limit
`rotation block 121 are summed in a summing junction
`block 111. The limitations of acceleration limit block
`127 to form a signal indicative of the longitudinal veloc
`50
`111 placed upon the signal from maneuver command
`ity error.
`scale factor block 109 are a function of the limitations of
`The output of summing junction 127 is fed into the
`the aircraft based upon its current velocity and current
`input of a displacement velocity loop gain block 129.
`pitch position.
`The output of summing junction 127 is also fed into the
`The output of acceleration limit block 111 is the lon
`input of an integral velocity loop gain block 131. The
`55
`gitudinal acceleration command, and is fed into an ac
`output of integral velocity loop gain block 131 is fed
`into the input of an integrator 133. The output of inte
`celeration feed forward block 113. The acceleration
`feed forward block has a gain of the 57.3 degrees to
`grator 133 and the output of displacement velocity loop
`radian conversion divided by 32.2 feet/sec2, the accel
`gain block 129 are summed in a summing junction 135
`eration of gravity and represents the required pitch
`to form 0CV, the pitch command for velocity loop con
`angle change to achieve the commanded acceleration
`trol.
`Referring to FIG. 4, the pitch stabilization control
`The sign of the acceleration feed forward block 113 is
`negative for longitudinal to account for positive (for
`block diagram is shown. A trim attitude table 201 ac
`cepts a calibrated airspeed input. This input is normally
`ward) acceleration requiring negative (down) pitch
`available from air data equipment typically available on
`attitude.
`The output of acceleration feed forward block 113 is
`aircraft. The output of trim attitude table block 201 is
`designated 6FF1 for reintroduction into another con
`the second feedforward pitch signal 6 FF2. Also shown
`trol sequence shown later.
`on FIG. 4 are the first feedforward pitch 0FF signal
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`command scale factor block 311 is fed into an accelera
`from acceleration feed forward block 113 of FIG.3 and
`the third feedforward pitch signal 6 FF3 which emanates
`tion limit block 313. The limitations of acceleration limit
`block 313 placed upon the signal from maneuver com
`from a control loop to be hereinafter discussed in FIG.
`8. The first, second, and third feedforward pitch com
`mand scale factor block 311 are a function of the limita
`tions of the aircraft based upon its current velocity and
`mands, 8FF1, 8FF2, and 8FF3, respectively are summed in
`a summing junction 203.
`current pitch position.
`Still referring to FIG. 4, the output of summing junc
`The lateral acceleration command output of accelera
`tion 203 is summed with 8 cv from FIG. 3 at summing
`tion limit block 313 is fed into a body orientation block
`junction 209. Further, the measured pitch attitude is
`315. Body orientation block 315 also receives the longi
`negatively summed with the total pitch command at
`tudinal acceleration command which was illustrated on
`summing junction 209 to form a pitch error signal. The
`FIG. 3. The operation carried on within body orienta
`measured pitch signal is normally obtained from aircraft
`tion block 315 was shown in the lower left half of FIG.
`flight data systems typically present on aircraft. The
`3 as a heading rotation matrix to rotate the acceleration
`pitch error signal output of summing junction 209 is fed
`command into an earth referenced north/east coordi
`into a pitch error gain block 211.
`nate frame.
`Near the bottom of FIG. 4, the shaped pitch stick rate
`The outputs of the body orientation block 315 include
`quantity from FIG. 3 appears and is fed into a stick rate
`a north acceleration signal and an east acceleration
`gain block 213.
`signal. The north acceleration signal is fed into an inte
`The shaped pitch stick quantity from FIG. 3 also
`grator 317. The output of integrator 317 is a north ve
`appears and is fed into a pitch rate maneuvering mode
`locity command which results from the integration of
`20
`gain block 215. The output of the pitch maneuvering
`the north acceleration signal. The east acceleration
`mode gain block 215 is summed with the output of the
`signal is fed into an integrator 319. The output of inte
`stick rate gain block 213 in a summing junction 217 to
`grator 319 is an east velocity command which results
`form the total pitch rate command.
`from the integration of the east acceleration signal.
`The measured pitch rate quantity, normally obtained
`The east velocity command from integrator 319 and
`25
`from aircraft flight data systems typically present on
`the north velocity command from integrator 317 are
`rotated back into the body axis frame through an in
`aircraft, is fed into a vibration filter block 219. The
`negative of the output of the vibration filter block 219,
`verse heading rotation matrix block 321. The outputs of
`the total pitch rate command output of the summing
`block 321 are the longitudinal velocity command and
`junction 217, and the output of the pitch error gain
`the lateral velocity command.
`30
`block 211 are summed in a summing junction 221, to
`This rotation, integration, and inverse rotation
`produce a pitch rate error signal.
`method allows stickinputs to command body axis accel
`The output of the summing junction 221 is fed into
`erations, but the resultant velocity commands are main
`the input of a pitch rate loop gain block 223. The output
`tained in the earth reference frame. This provides a
`of the summing junction 221 is also fed into the input of
`feature where the aircraft can be turned from its direc
`35
`a pitch rate loop integral gain block 225. The output of
`tion of flight to fire at a target, for example, without
`changing its flight path across the ground.
`the pitch rate loop integral gain block 225 is fed to an
`integrator 227. The output from integrator 227 and the
`To the right of the inverse headin