United States Patent (19)
`Skonieczny et al.
`
`USOO52382O3A
`11) . Patent Number:
`(45) Date of Patent:
`
`5,238,203
`Aug. 24, 1993
`
`4,106,094 8/1978 Land ................................... 364/434
`54 HIGH SPEED TURN COORDINATION FOR
`4,206,891 6/1980 Perez et al.
`. . . . . ... 244/17.3
`ROTARY WING AIRCRAFT
`4,312,039 l/1982 Skutecki........................... 244/17.13
`4,313,201 1/1982 Fischer et al. ...................... 364/434
`75) Inventors: Joseph P. Skonieczny, Madison;
`4,371,939 2/1983 Adams et al. ....................... 364/434
`Donald L. Fogler, Jr., Milford;
`4,382,283 5/1983 Clelford et al. ..................... 364/434
`Phillip J. Gold, Shelton, all of Conn.;
`4,477,876. 10/1984 Wright et al........................ 364/434
`James F. Keller, Media; James B.
`4,484,283 11/1984 Verzella et al. ..................... 364/434
`Wallingf
`4,626,998 12/1986 Adams et al. ....................... 244/179
`Dy?os wallingford, both of P.
`Primary Examiner-Joseph F. Peters, Jr.
`73) Assignee: United Technologies Corporation,
`ad
`r
`Hartford, Conn.
`Assistant Examiner-Virna Lissi Mojica
`Attorney, Agent, or Firm-Patrick J. O'Shea
`21 Appl. No.: 751,431
`57
`ABSTRACT
`22 Filed:
`Aug. 28, 1991
`5ll Int. Cli.............................................. G06F 15/50 A helicopter flight control system (21) includes a model
`52 U.S. C. .................................. 244/17.13; 364/434
`following control system architecture which automati
`58 Field of Search .................. 244/17.13, 17.21, 175,
`cally provides a coordinating yaw command signal to
`244/177, 179, 181, 184; 364/432-435
`the helicopter tail rotor to coordinate helicopter flight
`References Cited
`during a banked turn. The control system processes
`information from a variety of helicopter sensors (31) in
`U.S. PATENT DOCUMENTS
`order to provide the coordinating yaw command signal
`19
`on an output line (72) to the tail rotor (20) of the heli
`3,057,584 10/1962 Bretoi..................................
`3,688,099 8/1972 Buscher ............................... 244/195
`Copter.
`4,003,532 1/1977 Adams et al
`... 24.4/17.13
`w
`4,067,517 1/1978 Barnum ............................ 244/7.3
`15 Claims, 15 Drawing Sheets
`
`(56)
`
`
`
`
`
`
`
`
`
`
`
`
`
`
`
`
`
`
`
`HG SPEED
`TURN
`COORONATION
`(HSTC)
`LOGIC
`FIGS.5-7
`
`
`
`LOW SPEED
`URN
`COORONATION
`(STC)
`OGC
`FIGS.8-9
`
`14
`
`12
`
`BODY O
`EUER
`TRANSFORM
`FG.14
`
`110
`-1
`
`AIRSPEED
`
`HEADING
`
`EUERTO
`BODY
`TRANSFORM
`FG.5
`
`AND INEGRAL
`COMPENSATOR
`FG.6
`
`OX)
`(Y
`
`478
`
`BANK PTCH
`ANGLE ATTUDE
`
`461
`
`-36
`first
`PTC
`CONTROL
`OGIC-
`
`
`
`COMMANDED
`
`COMMANDED
`PTCH
`RATE
`
`BANK PTCH
`ANGE AUDE
`
`47
`PTCHATTUDE
`
`
`
`479
`
`DJI-1006
`IPR2023-01104
`
`

`

`U.S. Patent
`
`Aug. 24, 1993
`
`Sheet 1 of 15
`
`5,238,203
`
`
`
`N (JOSNES
`
`

`

`U.S. Patent
`
`Aug. 24, 1993
`
`Sheet 2 of 15
`
`5,238,203
`
`J3
`
`N/
`
`
`
`
`
`
`
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`-? 2
`
`RATE AND
`MAGNITUDE
`
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`
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`39 -1
`
`4.
`
`87
`(R ROI RAI -RATE 88
`LATERAL ACCE.
`89
`
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`
`LONGITUDINAL GND SPEED
`
`litics.GND SPEED
`
`m
`
`
`
`l
`T
`
`AFCS
`YAW
`OGC
`MODULE
`
`FIGS 3-9
`
`J
`T
`
`-------------------- --------- --
`
`

`

`U.S. Patent
`
`Aug. 24, 1993
`
`Sheet 3 of 15
`
`5,238,203
`
`
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`U.S. Patent
`
`Aug. 24, 1993
`
`Sheet 4 of 15
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`5,238,203
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`Aug. 24, 1993
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`Sheet 5 of 15
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`

`U.S. Patent
`
`Aug. 24, 1993
`
`sheet 6 of 15
`
`5,238,203
`
`O
`
`b
`
`AIRSPEED
`60-N
`
`263
`YAW RATE
`
`(-64
`(X-f
`265 264 Q2
`X
`
`260
`252.
`(A)
`x
`
`DESIRED AIRCRAFT 1 1 2
`YAW RATE
`(FROM PFCS)
`
`^
`
`54
`
`C
`
`1-KYCVFD
`261
`
`u-270
`
`AIRSPEED
`
`261/2,
`(X
`
`288
`
`268
`
`
`
`AIRSPEED
`
`60
`
`
`
`

`

`U.S. Patent
`
`Aug. 24, 1993
`
`Sheet 7 of 15
`
`5,238,203
`
`NER 152
`-154
`N
`
`AIRSPEED
`< 60 KNOTS?
`
`
`
`
`
`VXBTCS = 0
`
`
`
`-150
`
`SIDEARM
`CONTROLLER
`YAW
`INPUT 2
`AIRSPEED2
`80 KNOTS
`
`162
`
`
`
`
`
`N (E
`
`ENGAGE YAW CONTROL
`YAWENG s 1
`
`158
`
`= 0
`
`168
`
`PH1 < 2 DEGS
`16 N/COORDINATED TURN ?
`(i.e., IS THE TURN
`AND SLP NDICATOR
`BALL CENTERED 2)
`RSENS <
`2 DEGS/SEC 2
`
`164
`
`
`
`TEMP
`
`166
`170
`
`TWO SECOND
`TEMP
`TRANSTON DELAY
`
`TEMP = 1
`
`N
`
`176 K is )
`
`SIDEARM
`CONTROLLER ROLL
`INPUT ? (,
`
`
`
`YAWENG = 1
`TCENGHS = 1
`
`172
`
`Y
`
`
`
`
`
`
`
`174
`
`YAWENG
`TCENGHS
`
`
`
`
`
`182
`
`RETURN
`
`

`

`U.S. Patent
`
`Aug. 24, 1993
`
`Sheet 8 of 15
`
`5,238,203
`
`
`
`
`
`97
`
`03S/13
`
`STONBO1
`
`S’TÕEERHS
`
`

`

`U.S. Patent
`
`Aug. 24, 1993
`
`Sheet 9 of 15
`
`5,238,203
`
`ENTER
`
`352
`
`350
`
`358
`
`560
`
`PHIB C
`4 DEG 2
`
`N
`
`YAW RATE
`< 2 DEG/SEC 2
`
`TEMP=1
`
`564
`
`DELAY TEMP
`EDGE TRANSiTIONS
`TWO SECONDS
`
`366
`
`
`
`
`
`
`
`
`
`
`
`SIDEARM
`CONTROLLER ROLL
`INPUT 2
`
`372
`
`N
`
`DISABLE LOW SPEED
`TURN COORDINATION
`TCENGLS = 0
`
`356
`
`
`
`ENABLE LOW SPEED
`TURN COORDINATION
`TCENGS = 1
`
`

`

`Aug. 24, 1993
`
`U.S. Patent
`
`OOF
`
`
`
`NYALSILVNOLAY
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`
`
`Sheet 10 of 15
`
`5,238,203
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`
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`
`

`

`U.S. Patent
`
`Aug. 24, 1993
`
`Sheet 11 of 15
`
`5,238,203
`
`
`
`43O
`
`CREL
`
`39
`
`

`

`U.S. Patent
`
`Aug. 24, 1993
`
`Sheet 12 of 15
`
`5,238,203
`
`ENER
`
`152
`154
`N
`
`AIRSPEED
`< 60 KNOTS?
`
`
`
`VXBTCS = 0
`
`
`
`441
`
`DELPHI < 2 DEGS 2
`16N/COORDINATED TURN ?
`(i.e., IS THE TURN
`AND SLP INDICATOR
`p
`BALL CENTERED 2)
`RSENS <
`0.025 RAD/SEC.
`
`164
`
`
`
`
`
`
`
`TEMP s 1
`
`TWO SECOND
`TEMP
`TRANSiTION DELAY
`
`
`
`(TEMP = 1)
`OR.
`(TCREL = 1)
`
`
`
`166
`
`170
`
`443
`
`174
`
`157
`N
`
`-440
`
`SIDEARM
`CONTROLLER
`YAW NPUT 2
`
`AIRSPEED >
`80 KNOTS
`
`
`
`- 446
`TCENGHS = 1
`
`s
`
`CONTROLLER ROLL
`
`SIDEARM (.
`INPUT
`K.
`
`YAWENG
`= 0
`TCENGHS
`= 0
`
`
`
`182
`
`RETURN
`
`
`
`
`
`TCENGHS = 1
`
`FIG.1
`
`

`

`U.S. Patent
`
`Aug. 24, 1993
`
`Sheet 13 of 15
`
`.
`
`5,238,203
`
`
`
`
`
`
`
`
`
`
`
`
`
`
`
`
`
`
`
`
`
`DISABLE LOW SPEED
`TURN COORDINATION
`TCENGCL =
`
`
`
`
`
`
`
`
`
`N
`DELPHI <
`4 DEG 2
`
`w
`N
`
`Y
`ENABLE LOW SPEED 454
`TURN COORDINATION
`TCENGLS = 1
`E
`
`YAW RATE
`< 2 DEG/SEC 2
`
`DELAY TEMP
`EDGE TRANSiTIONS
`TWO SECONDS
`
`SIDEARM
`CONTROLLER ROLL
`INPUT 2
`
`
`
`DELPH2
`4 DEGS. 2
`GROUND SPEED
`> 16 FEET/SEC
`ENABLE Low SPEED
`TURN COORDINATION
`TCENGLS = 1
`
`458
`
`
`
`FIG.13
`
`

`

`U.S. Patent
`
`Aug. 24, 1993
`
`Sheet 14 of 15
`
`5,238,203
`
`89+
`
`09%
`
`
`LeetJ882
`
`.INIS
`
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`

`

`U.S. Patent
`U.S. Patent
`
`Aug, 24, 1993
`Aug. 24, 1993
`
`Sheet 150f15
`Sheet 15 of 15
`
`5,238,203
`5,238,203
`
`
`
`
`
`

`

`1.
`
`5
`
`10
`
`15
`
`HIGH SPEEDTURN COORDINATION FOR
`ROTARY WING AIRCRAFT
`The Government has rights in this invention pursuant
`to a contract awarded by the Department of the Army.
`DESCRIPTION
`Technical Field
`This invention relates to flight control systems for
`rotary wing aircraft, and more particularly to such
`flight control systems which provide automatic turn
`coordination control.
`Background Art
`A coordinated turn for a rotary wing aircraft (e.g. a
`helicopter having a single main rotor) is defined as a
`banked turn where the body of the aircraft is tangential
`to a curvilinear flight path and net acceleration is nor
`20
`mal to the floor of the aircraft (no side-slip vectors).
`Control of yaw axis commands to the tail rotor is criti
`cal in this type of maneuver.
`In mechanical linkage control systems a coordinated
`turn required that the pilot simultaneously input (via
`rudder pedals) the proper amount of yaw to match the
`25
`amount of roll input provided via the cyclic stick. The
`more recent fly-by-wire flight control systems (see for
`example U.S. Pat. Nos. 4,003,532, 4,067,517, 4,206,891
`and 4,484,283, all assigned to the assignee of the present
`invention) automatically provide the matching yaw
`30
`input. The automatic flight control system (AFCS)
`commands a coordinating yaw input (typically at air
`speeds above 60 knots) based on the sensed lateral accel
`eration. The coordinating yaw signal is then used to
`modify the main and tail rotor command signals as
`necessary to drive the helicopter's lateral acceleration
`to zero. However, all these systems use conventional
`control system architecture, and hence are not compati
`ble with "model following' fly-by-wire flight control
`Systems.
`DISCLOSURE OF THE INVENTION
`An object of the present invention is to provide an
`improved model following flight control system for a
`rotary wing aircraft, which automatically compensates
`45
`for actual yaw rate and lateral acceleration during high
`speed turn coordination of the aircraft.
`According to the present invention, an improved
`model following flight control system anticipates the
`turn coordination maneuver by monitoring the roll rate
`gyro, then calculates a desired (modeled) yaw rate
`value as the product of the sine value of the aircraft
`bank angle multiplied by the aircraft acceleration due to
`gravity (g), and divided by airspeed. In further accord
`with the present invention, the aircraft's sensed lateral
`acceleration is used to modify the desired yaw rate in
`order to drive the lateral acceleration error to zero. In
`still further accord with the present invention, the de
`sired yaw rate command value is variably scaled as a
`function of airspeed so as to bring in high speed turn
`60
`coordination gradually between 60 and 80 knots air
`speed.
`The present invention reduces pilot workload since
`turn coordination is automatically provided in the flight
`control system, and improves the aircraft flight path
`65
`control due to the system's inherent ability to act as a
`heading hold system. Furthermore the pilot is no longer
`required to input a yaw command. via rudder pedals
`
`5,238,203
`2
`since the present invention automatically provides the
`required yaw input for a coordinated turn.
`These and other objects, features, and advantages of
`the present invention will become more apparent in
`light of the following best mode embodiment thereof as
`illustrated in the accompanying Drawing.
`BRIEF DESCRIPTION OF THE DRAWING
`FIG. 1 is a block diagram of the improved model
`following flight, control system of the present inven
`tion;
`FIG. 2 is a schematic illustration of one portion of the
`embodiment of FIG. 1;
`FIG. 3 is a block diagram of one embodiment of
`system components illustrated in FIG. 2;
`FIG. 4 is a schematic illustration of the functional
`elements of the component illustrated in FIG. 3;
`FIG. 5 is a schematic illustration depicting further
`functional details of FIG. 4;
`F.G. 6 is a companion schematic illustration of FIG.
`5;
`FIG. 7 is a flow chart diagram illustrating execution
`of the functions disclosed in FIGS. 5, 6, in the embodi
`ment of FIG. 3;
`FIG. 8 is a schematic illustration depicting still fur
`ther functional details of FIG. 4;
`FIG. 9 is a flow chart diagram illustrating execution
`of the functions disclosed in FIGS. 8, in the embodi
`ment of FIG. 3;
`FIG. 10 is a schematic illustration of a trim control
`system which operates in conjunction with the logic of
`FIGS. 5-9:
`FIG. 11 is a figurative illustration, partially in per
`spective and partially in schematic, depicting details of
`another component used in the embodiment of FIG. 1;
`FIG. 12 is an alternative flow chart diagram to that
`illustrated in FIG. 7 to accommodate incorporation of
`the trim control of FIG. 10 into the flight control func
`tion illustrated in FIGS. 5, 6;
`FIG. 13 is an alternative flow chart diagram to that
`illustrated in FIG. 9, for controlling the switches within
`the LSTC logic of FIG. 8 when the trim function of
`FIG. 10 is incorporated;
`FIG. 14 is a detailed illustration of the Body to Euler
`Transform of FIG. 4;
`FIG. 15 is a detailed illustration of the Euler to Body
`Transform of FIG. 4;
`FIG. 16 is an illustration of the proportional and
`integral compensator of FIG. 4; and
`FIG. 17 is a pictorial illustration of an aircraft in
`which the flight control system embodiment of FIG. 1
`may be used.
`BEST MODE FOR CARRYING OUT THE
`INVENTION
`Referring first to FIG. 17, which is a pictorial illustra
`tion of a helicopter embodiment 18 of a rotary wing
`aircraft in which the present invention may be used.
`The helicopter includes a main rotor assembly 19 and
`tail rotor assembly 20.
`Referring now to FIG. 1, the helicopter flight control
`system of the present invention 21 is a model following
`control system which shapes the pilot's sidearm control
`ler and displacement stick commands through an "in
`verse vehicle model' to produce the desired aircraft
`response. The system includes a Primary Flight Control
`System (PFCS) 22 and an Automatic Flight Control
`
`35
`
`50
`
`55
`
`

`

`10
`
`5,238,203
`3
`4.
`System (AFCS) 24. The PFCS receives displacement
`rate of change of the aircraft for each pilot commanded
`command output signals from a displacement collective
`ael Ver.
`The summing function 58 sums the commanded yaw
`stick 26 online 27 and the AFCS receives the collective
`stick discrete output signals on a line 28. The PFCS and
`rate signal online 54 (from the yaw rate model 52) with
`AFCS each receive the force output command signals
`the aircraft's actual yaw rate, received (from sensors 31,
`of a four axis sidearm controller 29, on lines 30, and the
`through lines 32 and trunk 33) as a sensed yaw rate
`aircraft's sensed parameter signals from sensors 31, on
`signal on line 64, to provide a yaw rate error signal. The
`rate error signal is amplified in a rate gain stage 67 and
`lines 32. The pilot command signals on lines 27, 28, and
`30 and the sensed parameter signals on lines 32 are
`presented on a line 65 to one input of a second summing
`junction 66. The junction 66 also receives the desired
`shown consolidated within trunk lines 33 and 34 in the
`PFCS and AFCS, respectively.
`yaw rate signal on line 62 from the inverse model 56,
`and a yaw command modifying signal on a line 68 from
`The PFCS and AFCS each contain control channel
`logic for controlling the yaw, pitch, roll and lift axes of
`a rate and magnitude limiter 70. The limiter 70, which
`the aircraft. In FIG. 1 these logic modules are shown by
`receives a nonlimited version of the yaw command
`modifying signal on a line 84 (through bus 43) from the
`blocks 35-38 for the PFCS and blocks 39-42 for the
`15
`AFCS. The PFCS provides rotor command signals and
`AFCS yaw logic module 39, limits the yaw command
`the AFCS logic provides conditioning and/or trimming
`modifying signal if magnitude and rate of change limits
`of the PFCS four axis logic functions. The PFCS and
`are exceeded. The resulting sum signal is provided on
`AFCS logic modules interconnect through bus 43.
`the output line 72 of the PFCS yaw logic module 35,
`and presented through the PFCS output trunk lines 44
`As described in detail hereinafter, the PFCS and
`20
`AFCS use a model following algorithm in each control
`to the tail rotor servos (48, FIG. 1).
`axis to provide rotor command signals on output lines
`The magnitude and rate of change of the yaw com
`44 to a main rotor mixing function 45 which commands
`mand modifying signal from the AFCS is a function of
`displacement of mechanical servos 46 and linkages 47 to
`the aircraft heading error. The aircraft heading error is
`control the tip path plane of the main rotor 19. Rotor
`the second of two feedback loops around the tail rotor
`25
`command signals are also provided on lines 44 to the
`command signal; the first being the yaw error signal on
`helicopter's tail rotor servos 48 which control the thrust
`line 65. As described in detail hereinafter, the yaw com
`of the tail rotor 20 through linkages 49. The sensed
`mand modifying signal is a calculated value provided by
`parameter signals from sensors 31, on lines 32, provide
`a model following algorithm within the AFCS, based
`the PFCS and AFCS with the aircraft's angular rate
`on the actual aircraft response to the tail rotor com
`and attitude response to the rotor command signals.
`mand signal. The yaw command modifying signal mod
`ifies the signal magnitude and rate of change of the tail
`FIG. 2 is a partial schematic section of FIG. 1, illus
`rotor command signal by cancelling that component of
`trating the functional interconnection of the PFCS 22
`and AFCS 24 yaw logic modules 35 and 39, respect
`the actual yaw rate magnitude (of the sensed yaw rate
`fully. The PFCS yaw logic module 35 receives a yaw
`signal on line 64) which occurs during execution of a
`35
`axis command signal online 50, provided through trunk
`coordinated turn. The model following algorithm of the
`AFCS processes sensed aircraft parameters (sensors 31,
`lines 33 and lines 30, from the sidearm controller 29
`(FIG. 18). In the present embodiment the sidearm con
`FIG. 1) to generate command rate modifying signal
`troller is a four axis force stick in which the yaw axis
`characteristics which provide turn coordination at
`command signal is generated by the pilot's lateral twist
`speeds greater than 60 knots when the pilot applies roll
`ing (left or right yaw) of the sidearm controller. The
`stick inputs. The present invention lies in the shaping
`yaw command signal is presented to the input of yaw
`and conditioning of the yaw command modifying signal
`rate model circuitry 52 (e.g. a first order lag filter with
`in response to the sensed aircraft parameters.
`selected radians/sec/volt signal gain) that provides a
`As shown in FIG. 2, in addition to the commanded
`commanded yaw rate signal on a line 54 indicative of 45
`yaw rate signal received from the PFCS yaw logic
`the desired rate of change for the aircraft attitude about
`module 35 on line 54 (through trunk 43), the AFCSyaw
`the yaw axis. Selection of the yaw rate model is depen
`logic module 39 receives the following sensed aircraft
`dent on the dynamics of the aircraft and the yaw re
`parameters through trunk line 34: actual airspeed (line
`sponse desired.
`60), actual yaw rate (line 64), pitch attitude (line 86),
`The commanded yaw rate signal on line 54 is pres
`bank angle (PHI) (line 87), roll rate (line 88), lateral
`50
`ented simultaneously to: the input of a yaw-axis, vehicle
`acceleration (line 89), heading (line 90), longitudinal
`inverse model 56, a summing junction 58, and the bus 43
`ground speed (line 91), and lateral ground speed (line
`to the AFCS yaw logic module 39. The inverse model
`92). The best mode embodiment of the AFCS is as a
`56 receives the aircraft's actual airspeed from sensors
`microprocessor based electronic control system in
`31, through lines 32 and trunk 33, as a sensed airspeed
`which the algorithms of the AFCS logic modules
`55
`signal on line 60. The inverse model 56 is a Z-model
`(39–41, FIG. 1) reside in executable program listings
`transform, which may be embodied as a first order lead
`stored in memory.
`filter with instantaneous voltage gain and time constant
`FIG. 3, shows the architecture of a microprocessor
`characteristics which vary with the magnitude of the
`based AFCS 24. The commanded yaw rate signal on
`sensed airspeed signal online 60. The cascaded yaw rate
`line 54 is received from input lines 93 included within
`model 52 and inverse model 56 provide a feedforward
`the lines 43 interconnecting the AFCS and PFCS. The
`path for the sidearm control signal on line 50.
`sensed aircraft parameter signals on lines 60, 64, and
`86-92 are received from the AFCS input trunk line 34,
`The feedforward, inverse Z-model transform pro
`vides the primary control input to tail rotor 20 (FIG. 1)
`at an AFCS input port 94. Depending on the format of
`the input signals (analog or digital) the input port 94
`which causes the helicopter 18 (FIG. 17) to yaw at a
`65
`rate set by the commanded yaw rate signal on line 62.
`which may include an analog-to-digital converter, a
`This desired yaw rate signal represents the tail rotor
`frequency-to-digital convertor, and such other signal
`command necessary to achieve the desired yaw-axis
`conditioning functions known to those skilled in the art
`
`30
`
`

`

`5,238,203
`5
`6
`as being required to transform the input signals to digi
`WENG, and a high speed turn coordination engaged
`tal signal format.
`signal, TCENGHS.
`The input port is connected through an address/data
`FIG. 7 is a flow chart diagram of a routine 150 for
`bus 95 to a microprocessor 96 (e.g., Intel 80286, Motor
`controlling TCENGHS, and YAWENG. The routine is
`ola 68020), memory means 98 (including RAM, UV
`entered at step 152, where a first test 154 determines if
`the airspeed is less then or equal to 60 knots. If it is, the
`PROM, EEPROM), and an output port 100. The output
`port may comprise a digital-to-analog converter, a par
`signal VXBTCS is cleared by a step 156. Otherwise, a
`test 157 determines if a manual yaw input is being pro
`allel-to-serial convertor, a discrete output driver, and
`vided from the sidearm controller, and if it is a step 158
`such other signal conversion functions known to those
`engages manual yaw control by setting the discrete
`skilled in the art as being required to transform the
`YAWENG. If the pilot is not providing a yaw input via
`AFCS digital signal format to that required by the con
`the sidearm controller, a test 159 determines if the air
`trol system (21, FIG. 1). The output port lines, includ
`speed is greater than or equal to 80 knots, and if it is the
`ing the line 84 to the PFCS yaw logic module 35, are
`signal VXBTCS is set in a step 160. These steps and
`presented through lines 101 to the interconnecting lines
`tests 154,159 implement a hysteresis function which sets
`43
`VXBTCS when airspeed is above 80 knots, and clears
`FIG. 4 is a block diagram of the AFCS yaw logic 110
`VXBTCS when airspeed goes below 60 knots, assuming
`resident in the memory 98, and executable in the micro
`no yaw input is received. Although specific values are
`processor 96. The logic comprises both High Speed
`often times recited in the specification, they are used by
`Turn Coordination (HSTC) logic 112 and Low Speed
`way of example to facilitate an understanding of the
`20
`Turn Coordination (LSTC) logic 114. The HSTC and
`invention, and not by way of limitation on the inven
`the LSTC logics will be discussed separately, along
`tion. As may be understood by those skilled in the art,
`with a turn coordination trim systern to be discussed
`the actual values will depend on the requirements of
`hereinafter. Note, the HSTClogic includes the logic for
`each aircraft.
`conditioning pilot commanded yaw inputs via the side
`A test 162 is performed next to determine if the mag
`25
`arm controller at speeds above 60 knots.
`nitude of the bank angle, PHI is less than 2 degs. If it is,
`a test 163 is performed to determine if the aircraft is in
`HIGH SPEED TURN COORONATION
`a coordinated turn (i.e., is the pilot's turn and slip indica
`As shown in FIG. 5, the HSTClogic 112 receives the
`tor ball centered?). If test 163 determines the aircraft is
`bank angle signal, the airspeed signal, the lateral accel
`in a coordinated turn a test 164 is performed to deter
`30
`eration signal, and the roll rate signal. The bank angle
`mine if the magnitude of the yaw rate RSENS is less
`signal on the line 87 is input to logic 116 which imple
`than 2 degs/sec, and if it is a step 166 sets the variable
`ments the equation:
`TEMP. If either of tests 162-164 are false a step 168
`clearSTEMP.
`A subroutine 170 is performed next to delay any
`transitions of the variable TEMP for two seconds. It
`should be understood that the two second delay has
`been placed in the logic 150 in the interest of clarity, but
`clearly in a real time control system such as a flight
`control system of the present invention, delaying pro
`cessor functions for two seconds is unacceptable.
`Therefore, during the two second delay of the subrou
`tine, the microprocessor 96 (FIG. 3) is performing other
`tasks until the two seconds has elapsed, as well known
`to those skilled in the art.
`A test 172 is performed next to determine if TEMP is
`set, and if it is both automatic turn coordination and
`manual yaw control are disengaged (i.e., both
`TCENGHS and YAWENG are cleared) in a step 174.
`Otherwise, a test 176 determines if the magnitude of the
`bank angle, PHI is greater than or equal to two degrees,
`and if it is a test 178 is performed next. The test 178
`examines if the pilot is inputting a roll signal via the
`sidearm controller by checking the signal on the line 58
`(FIG. 2), and if he is inputting a roll signal, a test 179 is
`performed next to determine if VXBTCS is set and if so
`turn coordination is engaged by a step 180. A return
`step 182 is the final step in the routine 150.
`It should be noted in the logic illustrated in FIG. 7,
`that airspeed is only used as a criteria to engage the
`HSTC logic, it is not used as a criteria for disengaging
`the HSTC logic. The two criteria for disengaging the
`HSTC logic are bank angle PHI, and yaw rate RSENS.
`The test 163 is primarily used for the yaw manual con
`trol where yaw input is being received from the sidearm
`controller. If the magnitude of both PHI and RSENS
`falls below their respective predetermined minimum,
`and the aircraft is coordinated, both the HSTC and
`
`where
`rhstc= yaw rate necessary for a coordinated turn
`g=32.174 feet/sec2
`PHI=the bank angle signal on the line 87.
`The airspeed signal in knots on line 60 is input to again
`118 which converts the signal to feet/second and which
`provides a signal to a limiting function 120 to prevent a
`limited output signal on a line 122 from dropping below
`a certain value (e.g., 16 ft/sec) to prevent dividing by
`45
`zero. The limited output signal is input to a lag filter 124
`which provides the conditioned airspeed signal on a line
`126 to the logic 116.
`The logic 116 provides the signal on a line 128 to a
`switch 132 the operation of which shall be discussed in
`detail hereinafter. The signal rhsc is indicative of the
`amount of yaw required for a coordinated turn at a
`particular aircraft bank angle and airspeed.
`The lateral acceleration signal on the line 89 is input
`to a magnitude limiter 140, which provides a signal to a
`gain 142 that is controlled as a function of the airspeed
`signal. As airspeed increases from 60 knots to 80 knots,
`the magnitude of the gain 142 also increases from zero
`to a non-zero value, resulting in a larger product on a
`line 144. Similarly, as the airspeed decreases the magni
`tude of the gain decreases, resulting in a smaller product
`being output on the line 144. When airspeed is less than
`60 knots, the gain of gain function 142 is zero, providing
`a signal equal to zero on line 144. The product signal on
`the line 144 is input to a two pole switch 145 whose
`65
`operation is controlled by a discrete signal on a line 146
`produced by an OR gate 147. The OR gate is responsive
`to two discrete signals: pilot yaw input engaged YA
`
`rhtc= (g sin(PHI)/(conditioned airspeed)
`
`(Eq. 1)
`
`35
`
`O
`
`5
`
`55
`
`

`

`15
`
`5,238,203
`7
`8
`manual yaw control are disengaged. This allows the
`234 provides a signal on a line 250, and the summing
`HSTClogic to remain engaged throughout a decelerat
`function 240 provides a signal on a line 252.
`ing turn in which the airspeed drops below 60 knots, but
`Referring to FIG. 6, the remainder of the HSTC
`either yaw rate or bank angle remains above its respec
`logic 112 receives the signals on the lines 250,252 which
`tive predetermined minimum threshold defined in tests
`are provided to summing functions 254,256 respec
`tively. The logic also receives the aircraft yaw rate
`162,164.
`Referring back to FIG. 5 with the knowledge of how
`signal on the line 70 which is input to a summing func
`the HSTC logic is engaged/disengaged, the signal on
`tion 260. The yaw rate command signal on the line 62
`the line 144 is input to the switch 145. When the HSTC
`from the PFCS is input to both a gain function 261, and
`logic or manual yaw control are engaged (i.e.,
`to a multiplier 262 which receives an airspeed depen
`dant gain from a scheduler function 263. The multiplier
`TCENGHS=1 or YAWENG=1) the switch 145
`closes, allowing the signal on the line 144 to pass onto a
`function 262 provides a signal on a line 264 to the sum
`line 190 which goes to both a lag filter 192, and a sum
`ming functions 256 and 260. The summing function 260
`provides a signal on a line 265 to a multiplier 266 which
`ming function 194. Note, the switch 145 is illustrated in
`the open position. Making the switch responsive to both
`also receives an airspeed gain signal on a line 267 from
`a scheduler 268. The scheduler 268 provides again of
`TCENGHS and YAWENG, engages the lateral accel
`eration path (89,144,190) when either the high speed
`zero belows 60 knots, and ramps the gain from 60-80
`knots. The multiplier 266 provides a signal on a line 269
`turn coordination conditions are met, or the pilot is
`providing a manual input via the sidearm controller.
`to the summing function 254. Note the inverse symme
`The roll rate signal on the line 88 is input to a lag filter
`try of gain versus airspeed in the schedulers 263,268.
`20
`198 which attenuates high frequency noise, and pro
`The gain function 261 provides a signal on a line 270
`vides a signal on line 200 to a first limiting function 202,
`to limit functions 274,276. Limit function 274 provides a
`and to a second limiting function 204. The limits of the
`signal to the summing function 254, which in turn pro
`first and second limiting functions 202, 204, respec
`vides an output signal on a line 280. The limit function
`tively, are designed to provide a signal on a line 206
`276 provides a signal to a summing function 284, which
`25
`when the aircraft is rolled in one direction (e.g., 0" to
`also receives a signal on a line 286, and provides the sum
`179), and similarly to provide a signal on a line 208
`of these two signals on a line 288.
`when the aircraft is rolled in the other direction (e.g., 0
`FIG. 4-7 presents a detailed illustration of the control
`to - 179). Two gain functions 210,212 are also pro
`laws for the HSTC logic and the manual yaw control
`vided. The magnitude of the first gain 210 is less than
`logic. Nevertheless, the present invention lies in the
`30
`the magnitude of the second gain 212, to account for the
`HSTC computing a yaw signal necessary to provide a
`asymmetrical forces associated with rolling the aircraft
`coordinated turn at a particular aircraft bank angle and
`airspeed, and shaping/conditioning this desired yaw
`to the left, rather than the right, which arise as a result
`of the main rotor gyroscopic forces and other aircraft
`signal as a further function of airspeed, lateral accelera
`asymmetries. The outputs from the first and second gain
`tion, and roll rate to provide signals which operate in
`35
`210,212 are both input to a summing function 214 which
`the aircraft yaw control logic to provide for an im
`provides a turn anticipation signal on a line 216 to again
`proved high speed automatic turn coordination control
`system. By providing two different yaw rate command
`218.
`The gain 218, which is a function of the airspeed
`signals on the lines 280,288, the HSTC achieves in
`signal on the line 63 (similar to the gain 142), provides a
`proved automatic turn coordination, in part due to the
`signal to a limiting function 220 which in turn provides
`additional shaping/conditioning associated with having
`a signal on a line 222 to a switch 224. The state of the
`two separate lateral acceleration paths, one with a lag
`switch 224 is controlled by the signal TCENGHS, such
`filter 192, and the other without.
`that when TCENGHS is true (i.e., set) the switch is
`The manual yaw control is als