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Flight test of a digital controller used in a helicopter autoland system
Automatica, Vol. 23, No. 3, pp. 295 300, 1987
0005-1098/87 $3.00+0.00 Pergamon Journals Ltd. International Federation of Automatic Control
Printed in Great Britain.
Flight Test of a Digital Controller Used in a Helicopter Autoland System* DAVID R. DOWNINGI" and WAYNE H. BRYANT~
A velocity and heading controller with integral control, designed using direct digital optimal control procedures and combined with guidance and trajectory algorithms to form an autoland system, achieved successful trajectory tracking when implemented on a CH-47 helicopter and flight tested along a descending-decelerating trajectory to hover. Key Words--Aerospace computer control; direct digital control; feedforwardcontrol; optimal control; velocity control; flight test; (helicoptercontrol).
AIBtraet--This paper describes the flight test evaluation of an advanced digital helicopterflightcontrol system.The controller was designed using an optimal control design procedure for a fully coupled lateral and longitudinal vehicle model. Explicit integrals of the guidance error were included to produce a Type 1 characteristic.Gain scheduling was used to account for changes in the vehicle dynamics. The digital controller was exercised by combining it with state estimators, a trajectory generator and a closed-loop guidance algorithm to form a helicopter autoland system. A CH-47 tandem rotor helicopter was equipped with sensors, on-board digital flight computers and electro-hydraulicactuators. The system was exercised by automatically flyingstraight-in descendingdeceleratingtrajectories typical of VFR manual landing approaches. A description of the test-ground facilities, the flight hardware and software, and the velocityand position tracking performanceis included.
changes in velocity and heading guidance commands, thus providing an autotrim capability, and uses a gain scheduling technique to achieve constant response characteristics in spite of changes in the vehicle dynamics. To support the flight test, a Boeing-Vertol CH-47B tandem rotor helicopter has been equipped with an appropriate sensor complement, a generalpurpose digital computing system, and electrohydraulic actuators that permit full computer control of the vehicle. This flight system is self contained except for position data measured by a ground radar and sent to the vehicle by a digital data link. Included are descriptions of(l) the flight algorithms, (2) the flight test facilities, including both the g r o u n d and on-board hardware, and (3)the autoland system's position tracking performance and the controller's velocity tracking performance along a complex trajectory to a landing.
1. INTRODUCTION ONE PART of the National Aeronautics and Space Administration's Langley Research Center's research aimed at developing a navigation, guidance, control, display and flight management technology base, is the development of digital control design procedures and the verification of controller designs using flight test. This paper describes the flight test of an advanced digital controller. The controller is combined with a trajectory generator, a two-mode closed-loop guidance algorithm and sensor filters to form a multi-rate autoland software system. The control algorithm, operating at 10 iterations per second, achieves zero steady-state error to step
2. AUTOLAND SYSTEM ALGORITHMS The autoland system consists of a gain-scheduled velocity and heading controller, a set of estimators, a trajectory generator and a closed-loop guidance algorithm. The development of the controller and estimators is described by Stengel et al. (1978a, b, 1979). A brief description of the algorithms is given below with a discussion of the flight mechanization given later.
* Received 19 March 1986; revised 6 November 1986. The original version of this paper was not presented at any IFAC meeting.This paper was recommendedfor publicationin revised form by Associate Editor B. Friedland under the direction of Editor H. Austin Spang III. t Department of Aerospace Engineering, The University of Kansas, Lawrence,KS 66045, U.S.A. ~NASA Langley Research Center, Hampton, VA23665, U.S.A.
Control algorithm A direct digital design procedure based on quadratic synthesis modern control methods was used to simultaneously design lateral and longitudinal controllers for the present appfication. This design procedure accounts for all significant cross coupling 295
296
D.R. DOWNING and W. H. BRYANT
using the completely coupled vehicle model developed by Ostroff et al. (1976). A controller structure that explicitly includes integrals of guidance errors was used for all designs to provide an automatic trim capability. Although designed using perturbation techniques, the end controller is implemented in a form that uses only total vehicle states and produces total control signals. The control algorithm is
ALONG-TRACK 6~ ~i VELOCITY, 40 knots 20 0
VELOCIIY, ff/min
IL
0
500 r 400 ALT,TUDE,300 K
iB
J
j
LI
,,,
, ,
J I i
i
J ~
I i i i ]
~
-3 150 120
~
~ 90 rn
lOOl6000
~ 5000
100 m/mir/ '~'J 0
40(}0
30~0
2000
-13o 1000
0
RANGE TO GO, ff
FIG. 1. Nominal trajectory profiles.
Ua, = UA- 1 + T~Vh_1 Vk-1 = (I -- TcC2)Vk_2 - Cl(~k_ 1 -- ~k-2) --
YcC3(~k-2
-
Y.~._,) + C4(Yd. - Yd.-,)
where is the total state vector with components U, V, W, P, Q, R, ck, O, and ~b; ~t is the estimates of the state vector; U is the total control vector with components differential collective, collective, gang cyclic, and differential cyclic stick positions; V is the increment in the total control vector; Yd is the command vector and has as components the along-track, cross-track, and vertical velocities and heading; is the estimate of the vehicle's along-track, cross-track, and vertical velocities and heading; T~ is the controller update rate. X
The control gains C1, C2, C3 and C4 are updated every 2 s using algebraic functions of forward speed, vertical speed and turn rate. These algebraic expressions were derived by applying regression analysis to 12 fixed flight condition designs. These 12 designs were formed at flight conditions that covered the vehicle's flight envelope [forward speed 0-160 knots and rates of climb 2.5 to -5.1 ms -1 (+500 to - 1000fpm)]. This general structure has a digital proportionalintegral or Type 1 form; i.e. a step command will be followed with zero steady-state error. This feature provides an autotrim capability; no trim information is required. An examination of the equation for Uk shows that the most recent state estimate required to calculate the control at time k is $k- 1, the estimate based on sensor measurements approximately T~s old. In effect, this controller's structure has incorporated and accounts for in the gains a fixed T~ time lag for the state feedback loops, thus eliminating the effect of varying computation lags from cycle to cycle.
Estimator algorithm A set of estimator algorithms processes the vehicle sensors' measure of attitude, angular rate and
linear acceleration with the radar's measure of the vehicle's position to produce estimates of vehicle attitude, angular rate, position, ground velocity and accelerometer bias. The estimators use a variety of design techniques. The Euler Angle estimator is a two-step complementary filter using current measurements of the Euler angles and body angular rates. The estimate of body angular rates uses a low-pass filter to eliminate sensor noise. The vehicle position, ground speed and accelerometer bias estimator are developed using a steady-state Kalman filter which combines body-mounted accelerometer measurements with ground-based radar position measurements.
Trajectory generator and guidance algorithms The trajectory selected to exercise the autoland system is based upon the results of studies by Moen and Yenni (1975) and Moen et al. (1976) which examined manual helicopter visual flight rule approach profiles. These studies examined profiles flown by a group of pilots using several types of VTOL vehicles to determine a preferred profile. These profiles incorporated coordinated descent and deceleration maneuvers which were more demanding than constant-glide-slope constantvelocity profiles used in standard IFR approaches. The trajectory mechanized (Fig. 1) consists of three segments: a straight and level segment at constant speed and altitude, a descent-and-deceleration-to-a-hover segment and a vertical-descent segment. This trajectory was mechanized in the runway frame in the trajectory generator module using range-to-go as the independent variable. Values of nominal longitudinal velocity, vertical velocity and altitude at 25 values of range-to-go were stored with intermediate values determined using a linear interpolation algorithm. The nominal lateral velocity was zero, and the nominal heading was runway heading. The guidance algorithm operates in the runway frame and_ uses estimates of the navigation frame velocity, ~ , and position, :~e, from the estimators, and the nominal velocity, )~,, in the runway frame from the nominal trajectory generator. The guidance algorithm produces commanded runway
Helicopter digital controller flight test frame velocity, Xg, and heading, ~0, and thus y[ = [~[, ~bu]. A sideslip technique was used to correct for lateral position and velocity errors; therefore, the commanded heading was always the runway heading. The commanded velocity algorithm had two modes. The first mode operated from the beginning of the trajectory to hover. This mode modified the nominal velocity to correct for position and velocity errors using
j
t I
X, = X. +
[i.-
0
+
]
K,
K,
-'
-
/
t
#
~FPS-le'~'E~TR^CKING~
'
..................
'
FIG. 2. Flight test facilities.
Ii ° °o) 0
HIi
297
IX.
-
HI~]
Ks
where: 1~, and X, are the nominal velocity and position in runway coordinates, and H~ is the fixed transformation from the navigation frame to the runway frame. Upon arriving at hover, the position and velocity deviations from nominal values were checked; when the horizontal position was within 6.1 m (20 feet) of the pad, the altitude was within + 6.1 m (+ 20 feet) of the nominal hover altitude, the vertical velocity was within + 0.31 m s-1 (+ 1 fps), and the lateral and longitudinal velocities were each within + l . 5 m s -1 (+5fps), the second guidance mode was activated. This descent mode eliminated the position and velocity correction terms and simply commanded a fixed descent rate to touchdown; i.e. ~ = (0,0, 2,), where 2 , was fixed at 0.46ms-l (1.5fps). This guidance mode relied solely on the controller's ability to maintain zero longitudinal and lateral ground relative velocity during the descent to maintain acceptable position dispersions at touchdown. 3. FLIGHT TEST G R O U N D FACILITY
Figure 2 is a pictorial representation of the main Wallops Flight Center facilities used throughout the autoland system flight test. The Aeronautical Radar Research Complex (ARRC Systems) consists of a precision radar, a laser tracker, ground computers and a two-way digital data link. Aircraft position relative to the ARRC site is measured by the FPS-16 radar and laser tracker and converted in the ground computer to a pad-centered Cartesian coordinate system aligned with the Wallops runway selected for the day's operations. The radar tracker has a 1 sigma range error of 10 feet and azimuth
RADAR/LASER TRACKER
FIG. 3. Block diagram of flight system.
and elevation errors with 1 sigma value of 1 x 10-" radians. The laser has a 1 sigma range error of 2 feet. The measured position data is transmitted to the research vehicle using a two-way digital data link that uses the radar/aircraft transponder systems for data transmission during the time intervals between radar ranging pulses. This position data is combined with on-board accelerometer data in the on-board computer to estimate aircraft position and ground-referenced velocity. To create displays for experiment initialization and subsequent realtime monitoring and evaluation, vehicle state variables are transmitted from the aircraft to the ARRC facility using the down link and subsequently from the ARRC facility to the Flight Display Research System (FDRS) using a hard-wired two-way digital data link. At the FDRS, this data is used by an Adage graphics computer system to create a display which is then transmitted by a television link to the research aircraft. 4. FLIGHT HARDWARE AND SOFTWARE
The basic vehicle system, described by Kelly et ai. (1979), includes the on-board sensors (flight control attitude/rate gyros and linear accelerometers), a Sperry 1819A flight computer, a set of analog-to-digital (A/D) and digital-to-analog (D/A) converters and a set of electro-hydraulic actuators. Added to the basic system (Fig. 3) using a specially designed two-way digital-to-digital (D/D) interface
298
D.R. DOWNING and W. H. BRYANT
TABLE 1. SUBROUTINEITERATIONRATEAND COMPUTATIONTIME VER] SPD,
Iteration Computation rate time (s- 1)
Control output/sensor input Controller Gain schedule Estimator Trajectory generator/guidance Executive Downlink formatting/digital recording
fpr~ 1500!
]4.22.2
RUN~
~
7~
I
~
A[r, •/I00
(ms)
10 10 0.5 10 2 10
20.0 12.0 24.0 11.5 3.5 2.5
10
14.5
j 500 ~
\ GF~OUNO fRACK
-~ 14 -4
i I l l l 20 I0 0 I0 2O CROSS-TRACK VELOCITY
FIG. 4. Experiment initialization display. is a ROLM 1664 flight computer. The interface of the autoland system through the Sperry 1819A computer was an experimental vehicle constraint. The ROLM 1664 was capable of implementing the total autoland system. The ROLM 1664 was added to the existing system to take advantage of the availability of a higher order language (FORTRAN) and floating-point arithmetic. The flight software consists of an executive routine and a set of subroutines to carry out the various sub-tasks associated with the autoland system operation. The subroutines include a velocity controller, controller gain scheduling, a set of estimators, a trajectory generator, a closed-loop guidance algorithm, input/output routines between the Sperry 1819A and ROLM 1664 computer, mode control and digital recording. With the exception of assembly language coding of the input/output between the Sperry and ROLM computers, all ROLM coding is done in FORTRAN and all computations use floating-point arithmetic. This capability greatly reduced the time required to develop and check out flight coding. The subroutines operate at three different iteration rates as shown in Table 1. Table 1 also presents estimates of the computation time for each module. The maximum computation time occurs during a 0.1 s cycle when all subroutines are exercised. This maximum time is 88 ms, leaving 12 ms unused. 5. FLIGHT TEST RESULTS The purpose of the flight test was to evaluate the controller design as part of an autoland system. The trajectory described earlier exercised the longitudinal and vertical velocity loops with tracking tasks and exercised the lateral velocity and heading loops as hold modes. The lateral guidance algorithm was exercised as a result of initial condition errors. A total of nine automatic landings were made in winds up to 15 knots with crosswind components up to 6 knots. The experiments were initialized by manually establishing the ground track, altitude, longitudinal velocity, lateral velocity, vertical velocity and heading at the desired values and then switching on the
autoland mode. To aid the pilot in this task, the FDRS generated the display shown in Fig. 4 and transmitted this to the vehicle. Initialization errors less than _+2 knots in longitudinal velocity, + 4 knots in lateral velocity, +0.25 m s- 1 ( + 50 fpm) in vertical velocity, + 12.1 m (+40 feet) in altitude, 2 ° in heading and +122m (+400 feet) in lateral position were achieved. The controller is evaluated as (1)a command tracker, (2)an autotrim system and (3) part of an autoland system. The digital controller command tracking performance is illustrated for a typical run in Fig. 5. The longitudinal and lateral velocity tracking is good, with errors smaller than + 2 knots, during segments with constant velocity commands. During the longitudinal deceleration segment, a hang-off error of approximately 4 knots develops. The heading tracking of the constant runway heading command is within +1 °. The hang-off during decelerations and small errors during segments with constant velocity and heading commands are characteristics of the Type 1 property predicted by theory for the controller's structure. The vertical velocity command response exhibits an undesirable low damped oscillation with peak errors of + 1.27 m s-1 (+ 250fpm). Further refinements of the vertical control and guidance algorithms are required but were not possible due to the limited flight test time. The nominal trajectory used in this evaluation (Fig. 1) requires the vehicle-controller combination to operate at and transition between the four trim conditions given in Table 2. This is automatically accomplished by the control algorithm without any knowledge of the theoretical trim values for the controls or vehicle attitudes. Figure 6 verifies the theoretical trim values and thus the model developed by Ostroff et al. (1976) used for the trim pitch attitude and collective control. An evaluation of the total differential collective control is predicted by the fact that only the incremental change about an unknown initial condition was recorded. A comparison of the incremental changes listed in
299
Helicopter digital controller flight test
80 ~ ALONGTRACK VELOCITY40 ~ COMMAND, knots 0L
J
~
~
ALONGTRACK VELOCITY 10 ~ ~ _ . . ~ . ERROR, knots 0 -10 I0 r CROSSTRACK VELOCITY , L O COMMAND, knots -I0 L_
__ ]
I
•
CROSSTRACK VELOCITY 5 ERROR, knots 0 -5 VERTICAL VELOCITY 1000 COMMAND, ft/min O -I000 250 F VERTICAL VELOCITY ERROR, ft/min 0 ' ~ v ~ V ~wvV~ V ~ X ~ v " ~ -250 10 HEADING ERROR, deg .0 I I I 1 -10 r i 0 20 40 60 80 I00 TI ME, sec
300 0 m/min -300 75
~
0 m/min -75
I
120
FIG. 5. Typical velocity and heading tracking time history. TABLE 2. C H - 4 7 B NOMINAL TRIM CONDITIONS
(s)
Longitudinal velocity (knots)
Descent Rate (m s- 1) (ft rain- 1)
0
65
0
20
65
3.21 (600)
80
0
0
100
0
0.48 (90)
Nominal time
3.50 r-
c THEORETICALTRIM VALUES
I NCREMENTAL1 75 ~DIFFERENTIAL " 0 [ ~ . _ . _. ~ . - . . . . .
COLLECTIVE, in.
CO
I,,
~I lO t
I ~
~
5 ocm
-1.75 i-
-5
-3.50 b
-I0
iVE, •
i0cm
1"20|
,
,
I
1
I
20
40
60 TIME. sec
80
100
1° 5
0
ANGLE, 1 (leg.
0
120
FIG. 6. Typical longitudinal controls and pitch time histories.
Trim differential collective CM (inches) - 1.94
(-0.725) -2.33 (-0.869) - 0.78 (-0.292) - 0.79 (-0.295)
Trim collective CM (inches) 8.58 (3.21) 6.93 (2.59) 11.82 (4.42) 11.69 (4.37)
Trim pitch angle (deg.) 2.94 3.14 6.61 6.61
Table 2 between t = 0 and t = 80 s and the changes observed in flight show the theoretical change to b ¢ approximately one-half as large as that observed in flight. This indicates an inaccuracy in the model. The use of the Type 1 structure automatically corrects for this type of uncertainty. The autoland position tracking task includes runway centerline tracking and altitude nominal profile tracking during the approach to hover and lateral and longitudinal deviation from the center of the landing pad at touchdown. Figure 7 shows the excellent position tracking performance of the autoland system during approach to hover. The initial deviation from centerline is reduced to less than _+6.1m (_+20 feet) within 30s. The
300
D . R . DOWNING and W. H. BRYANT
CROSSTRACK~ F~-- ~ ERROR
li
u ~-
.
i
.
~
=
.
-
.
~
.
=
.
~
±
]
~
~
-50 ~
0 m I
• -30
-100 60 F
15
ALTITUDE
0 rn
ERROR.(130r-601
0
k
20
I __
40
I
60
I __.
80
I
I00
--15
120
TIME, sec
FIG. 7. Autoland system position tracking capability.
maximum altitude errors of 3.66m (12 feet) occur at the initiation of the descent segment. The altitude error at hover is 2.74 m (9 feet). The mean touchdown dispersion, based on nine landings, is 2.93 m (9.6 feet). If it were necessary to reduce the touchdown dispersion, position correction guidance could easily be incorporated in the landing guidance mode. 6. CONCLUSION An advanced low-iteration-rate (10Hz) gainscheduled digital controller based on modern control theory was designed for the Boeing Vertol CH-47 tandem rotor helicopter. This controller tracked commanded longitudinal, lateral and vertical velocity and vehicle heading; when combined with dosed-loop guidance and nominal trajectory generation algorithms, it formed a helicopter autoland system. The autoland system was flight tested along a descending deceleration profile to hover followed by a vertical descent to touchdown. Based on nine automatic landings, the autoland system is shown to have excellent position tracking performance with maximum altitude and lateral tracking
errors, after initial condition errors are reduced, of 3.66m (12 feet) and 6.1 m (9.6 feet). The velocity tracking performance is good with the longitudinal and lateral velocity tracking less than + 2 knots except during longitudinal deceleration when a 4 knot hang-off error develops. Heading tracking is excellent, being less than 1°. The vertical velocity controller exhibited a lightly damped oscillation with peak values of 1.27 m s- 1 (250 fpm) that would require further refinement. The flight tests have verified the digital controller design procedure, as well as the autotrim and the Type 1 features of the longitudinal controller predicted by theory. Also demonstrated is the feasibility of implementing advanced autoland systems in state-of-the-art flight computers using a higher order language. REFERENCES Kelly,J. R., F. R. Niessen,J. R. Garren and T. S. Abbott (1979). The VTOL approach and landing technology (VALT) CH-47 research system. NASA TP-1436. Moen, G. C., D. J. DiCarlo and K. R. Yenni(1976).A parametric analysis of visual approaches for helicopters. NASA TN D-8275. Moen, G. C. and K. R. Yenni (1975). Simulation and flight studies of an approach profile indicator for VTOL aircraft. NASA TN D-8051. Ostroff,A. J., D. R. Downingand W. J. Rood (1976).A technique using a nonlinear helicoptermodel for determiningtrims and derivatives. NASA TN D-9159. Stengel, R. F., J. R. Broussard and P. W. Berry (1978a). Digital flight control design for a tandem-rotor helicopter. Automatica, 14, 301-312. Stengel, R. F., J. R. Broussard and P. W. Berry (1978b). Digital controllersfor VTOL aircraft.IEEE Trans. Aerospace Electron. Syst., AES-14, 54-63. Stengel, R. F., J. R. Broussard and P. W. Berry (1979). Modern digital flight control systemdesign for VTOL aircraft. NASA CR 159019.
0005-1098/87 $3.00+0.00 Pergamon Journals Ltd. International Federation of Automatic Control
Printed in Great Britain.
Flight Test of a Digital Controller Used in a Helicopter Autoland System* DAVID R. DOWNINGI" and WAYNE H. BRYANT~
A velocity and heading controller with integral control, designed using direct digital optimal control procedures and combined with guidance and trajectory algorithms to form an autoland system, achieved successful trajectory tracking when implemented on a CH-47 helicopter and flight tested along a descending-decelerating trajectory to hover. Key Words--Aerospace computer control; direct digital control; feedforwardcontrol; optimal control; velocity control; flight test; (helicoptercontrol).
AIBtraet--This paper describes the flight test evaluation of an advanced digital helicopterflightcontrol system.The controller was designed using an optimal control design procedure for a fully coupled lateral and longitudinal vehicle model. Explicit integrals of the guidance error were included to produce a Type 1 characteristic.Gain scheduling was used to account for changes in the vehicle dynamics. The digital controller was exercised by combining it with state estimators, a trajectory generator and a closed-loop guidance algorithm to form a helicopter autoland system. A CH-47 tandem rotor helicopter was equipped with sensors, on-board digital flight computers and electro-hydraulicactuators. The system was exercised by automatically flyingstraight-in descendingdeceleratingtrajectories typical of VFR manual landing approaches. A description of the test-ground facilities, the flight hardware and software, and the velocityand position tracking performanceis included.
changes in velocity and heading guidance commands, thus providing an autotrim capability, and uses a gain scheduling technique to achieve constant response characteristics in spite of changes in the vehicle dynamics. To support the flight test, a Boeing-Vertol CH-47B tandem rotor helicopter has been equipped with an appropriate sensor complement, a generalpurpose digital computing system, and electrohydraulic actuators that permit full computer control of the vehicle. This flight system is self contained except for position data measured by a ground radar and sent to the vehicle by a digital data link. Included are descriptions of(l) the flight algorithms, (2) the flight test facilities, including both the g r o u n d and on-board hardware, and (3)the autoland system's position tracking performance and the controller's velocity tracking performance along a complex trajectory to a landing.
1. INTRODUCTION ONE PART of the National Aeronautics and Space Administration's Langley Research Center's research aimed at developing a navigation, guidance, control, display and flight management technology base, is the development of digital control design procedures and the verification of controller designs using flight test. This paper describes the flight test of an advanced digital controller. The controller is combined with a trajectory generator, a two-mode closed-loop guidance algorithm and sensor filters to form a multi-rate autoland software system. The control algorithm, operating at 10 iterations per second, achieves zero steady-state error to step
2. AUTOLAND SYSTEM ALGORITHMS The autoland system consists of a gain-scheduled velocity and heading controller, a set of estimators, a trajectory generator and a closed-loop guidance algorithm. The development of the controller and estimators is described by Stengel et al. (1978a, b, 1979). A brief description of the algorithms is given below with a discussion of the flight mechanization given later.
* Received 19 March 1986; revised 6 November 1986. The original version of this paper was not presented at any IFAC meeting.This paper was recommendedfor publicationin revised form by Associate Editor B. Friedland under the direction of Editor H. Austin Spang III. t Department of Aerospace Engineering, The University of Kansas, Lawrence,KS 66045, U.S.A. ~NASA Langley Research Center, Hampton, VA23665, U.S.A.
Control algorithm A direct digital design procedure based on quadratic synthesis modern control methods was used to simultaneously design lateral and longitudinal controllers for the present appfication. This design procedure accounts for all significant cross coupling 295
296
D.R. DOWNING and W. H. BRYANT
using the completely coupled vehicle model developed by Ostroff et al. (1976). A controller structure that explicitly includes integrals of guidance errors was used for all designs to provide an automatic trim capability. Although designed using perturbation techniques, the end controller is implemented in a form that uses only total vehicle states and produces total control signals. The control algorithm is
ALONG-TRACK 6~ ~i VELOCITY, 40 knots 20 0
VELOCIIY, ff/min
IL
0
500 r 400 ALT,TUDE,300 K
iB
J
j
LI
,,,
, ,
J I i
i
J ~
I i i i ]
~
-3 150 120
~
~ 90 rn
lOOl6000
~ 5000
100 m/mir/ '~'J 0
40(}0
30~0
2000
-13o 1000
0
RANGE TO GO, ff
FIG. 1. Nominal trajectory profiles.
Ua, = UA- 1 + T~Vh_1 Vk-1 = (I -- TcC2)Vk_2 - Cl(~k_ 1 -- ~k-2) --
YcC3(~k-2
-
Y.~._,) + C4(Yd. - Yd.-,)
where is the total state vector with components U, V, W, P, Q, R, ck, O, and ~b; ~t is the estimates of the state vector; U is the total control vector with components differential collective, collective, gang cyclic, and differential cyclic stick positions; V is the increment in the total control vector; Yd is the command vector and has as components the along-track, cross-track, and vertical velocities and heading; is the estimate of the vehicle's along-track, cross-track, and vertical velocities and heading; T~ is the controller update rate. X
The control gains C1, C2, C3 and C4 are updated every 2 s using algebraic functions of forward speed, vertical speed and turn rate. These algebraic expressions were derived by applying regression analysis to 12 fixed flight condition designs. These 12 designs were formed at flight conditions that covered the vehicle's flight envelope [forward speed 0-160 knots and rates of climb 2.5 to -5.1 ms -1 (+500 to - 1000fpm)]. This general structure has a digital proportionalintegral or Type 1 form; i.e. a step command will be followed with zero steady-state error. This feature provides an autotrim capability; no trim information is required. An examination of the equation for Uk shows that the most recent state estimate required to calculate the control at time k is $k- 1, the estimate based on sensor measurements approximately T~s old. In effect, this controller's structure has incorporated and accounts for in the gains a fixed T~ time lag for the state feedback loops, thus eliminating the effect of varying computation lags from cycle to cycle.
Estimator algorithm A set of estimator algorithms processes the vehicle sensors' measure of attitude, angular rate and
linear acceleration with the radar's measure of the vehicle's position to produce estimates of vehicle attitude, angular rate, position, ground velocity and accelerometer bias. The estimators use a variety of design techniques. The Euler Angle estimator is a two-step complementary filter using current measurements of the Euler angles and body angular rates. The estimate of body angular rates uses a low-pass filter to eliminate sensor noise. The vehicle position, ground speed and accelerometer bias estimator are developed using a steady-state Kalman filter which combines body-mounted accelerometer measurements with ground-based radar position measurements.
Trajectory generator and guidance algorithms The trajectory selected to exercise the autoland system is based upon the results of studies by Moen and Yenni (1975) and Moen et al. (1976) which examined manual helicopter visual flight rule approach profiles. These studies examined profiles flown by a group of pilots using several types of VTOL vehicles to determine a preferred profile. These profiles incorporated coordinated descent and deceleration maneuvers which were more demanding than constant-glide-slope constantvelocity profiles used in standard IFR approaches. The trajectory mechanized (Fig. 1) consists of three segments: a straight and level segment at constant speed and altitude, a descent-and-deceleration-to-a-hover segment and a vertical-descent segment. This trajectory was mechanized in the runway frame in the trajectory generator module using range-to-go as the independent variable. Values of nominal longitudinal velocity, vertical velocity and altitude at 25 values of range-to-go were stored with intermediate values determined using a linear interpolation algorithm. The nominal lateral velocity was zero, and the nominal heading was runway heading. The guidance algorithm operates in the runway frame and_ uses estimates of the navigation frame velocity, ~ , and position, :~e, from the estimators, and the nominal velocity, )~,, in the runway frame from the nominal trajectory generator. The guidance algorithm produces commanded runway
Helicopter digital controller flight test frame velocity, Xg, and heading, ~0, and thus y[ = [~[, ~bu]. A sideslip technique was used to correct for lateral position and velocity errors; therefore, the commanded heading was always the runway heading. The commanded velocity algorithm had two modes. The first mode operated from the beginning of the trajectory to hover. This mode modified the nominal velocity to correct for position and velocity errors using
j
t I
X, = X. +
[i.-
0
+
]
K,
K,
-'
-
/
t
#
~FPS-le'~'E~TR^CKING~
'
..................
'
FIG. 2. Flight test facilities.
Ii ° °o) 0
HIi
297
IX.
-
HI~]
Ks
where: 1~, and X, are the nominal velocity and position in runway coordinates, and H~ is the fixed transformation from the navigation frame to the runway frame. Upon arriving at hover, the position and velocity deviations from nominal values were checked; when the horizontal position was within 6.1 m (20 feet) of the pad, the altitude was within + 6.1 m (+ 20 feet) of the nominal hover altitude, the vertical velocity was within + 0.31 m s-1 (+ 1 fps), and the lateral and longitudinal velocities were each within + l . 5 m s -1 (+5fps), the second guidance mode was activated. This descent mode eliminated the position and velocity correction terms and simply commanded a fixed descent rate to touchdown; i.e. ~ = (0,0, 2,), where 2 , was fixed at 0.46ms-l (1.5fps). This guidance mode relied solely on the controller's ability to maintain zero longitudinal and lateral ground relative velocity during the descent to maintain acceptable position dispersions at touchdown. 3. FLIGHT TEST G R O U N D FACILITY
Figure 2 is a pictorial representation of the main Wallops Flight Center facilities used throughout the autoland system flight test. The Aeronautical Radar Research Complex (ARRC Systems) consists of a precision radar, a laser tracker, ground computers and a two-way digital data link. Aircraft position relative to the ARRC site is measured by the FPS-16 radar and laser tracker and converted in the ground computer to a pad-centered Cartesian coordinate system aligned with the Wallops runway selected for the day's operations. The radar tracker has a 1 sigma range error of 10 feet and azimuth
RADAR/LASER TRACKER
FIG. 3. Block diagram of flight system.
and elevation errors with 1 sigma value of 1 x 10-" radians. The laser has a 1 sigma range error of 2 feet. The measured position data is transmitted to the research vehicle using a two-way digital data link that uses the radar/aircraft transponder systems for data transmission during the time intervals between radar ranging pulses. This position data is combined with on-board accelerometer data in the on-board computer to estimate aircraft position and ground-referenced velocity. To create displays for experiment initialization and subsequent realtime monitoring and evaluation, vehicle state variables are transmitted from the aircraft to the ARRC facility using the down link and subsequently from the ARRC facility to the Flight Display Research System (FDRS) using a hard-wired two-way digital data link. At the FDRS, this data is used by an Adage graphics computer system to create a display which is then transmitted by a television link to the research aircraft. 4. FLIGHT HARDWARE AND SOFTWARE
The basic vehicle system, described by Kelly et ai. (1979), includes the on-board sensors (flight control attitude/rate gyros and linear accelerometers), a Sperry 1819A flight computer, a set of analog-to-digital (A/D) and digital-to-analog (D/A) converters and a set of electro-hydraulic actuators. Added to the basic system (Fig. 3) using a specially designed two-way digital-to-digital (D/D) interface
298
D.R. DOWNING and W. H. BRYANT
TABLE 1. SUBROUTINEITERATIONRATEAND COMPUTATIONTIME VER] SPD,
Iteration Computation rate time (s- 1)
Control output/sensor input Controller Gain schedule Estimator Trajectory generator/guidance Executive Downlink formatting/digital recording
fpr~ 1500!
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~
7~
I
~
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(ms)
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10
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j 500 ~
\ GF~OUNO fRACK
-~ 14 -4
i I l l l 20 I0 0 I0 2O CROSS-TRACK VELOCITY
FIG. 4. Experiment initialization display. is a ROLM 1664 flight computer. The interface of the autoland system through the Sperry 1819A computer was an experimental vehicle constraint. The ROLM 1664 was capable of implementing the total autoland system. The ROLM 1664 was added to the existing system to take advantage of the availability of a higher order language (FORTRAN) and floating-point arithmetic. The flight software consists of an executive routine and a set of subroutines to carry out the various sub-tasks associated with the autoland system operation. The subroutines include a velocity controller, controller gain scheduling, a set of estimators, a trajectory generator, a closed-loop guidance algorithm, input/output routines between the Sperry 1819A and ROLM 1664 computer, mode control and digital recording. With the exception of assembly language coding of the input/output between the Sperry and ROLM computers, all ROLM coding is done in FORTRAN and all computations use floating-point arithmetic. This capability greatly reduced the time required to develop and check out flight coding. The subroutines operate at three different iteration rates as shown in Table 1. Table 1 also presents estimates of the computation time for each module. The maximum computation time occurs during a 0.1 s cycle when all subroutines are exercised. This maximum time is 88 ms, leaving 12 ms unused. 5. FLIGHT TEST RESULTS The purpose of the flight test was to evaluate the controller design as part of an autoland system. The trajectory described earlier exercised the longitudinal and vertical velocity loops with tracking tasks and exercised the lateral velocity and heading loops as hold modes. The lateral guidance algorithm was exercised as a result of initial condition errors. A total of nine automatic landings were made in winds up to 15 knots with crosswind components up to 6 knots. The experiments were initialized by manually establishing the ground track, altitude, longitudinal velocity, lateral velocity, vertical velocity and heading at the desired values and then switching on the
autoland mode. To aid the pilot in this task, the FDRS generated the display shown in Fig. 4 and transmitted this to the vehicle. Initialization errors less than _+2 knots in longitudinal velocity, + 4 knots in lateral velocity, +0.25 m s- 1 ( + 50 fpm) in vertical velocity, + 12.1 m (+40 feet) in altitude, 2 ° in heading and +122m (+400 feet) in lateral position were achieved. The controller is evaluated as (1)a command tracker, (2)an autotrim system and (3) part of an autoland system. The digital controller command tracking performance is illustrated for a typical run in Fig. 5. The longitudinal and lateral velocity tracking is good, with errors smaller than + 2 knots, during segments with constant velocity commands. During the longitudinal deceleration segment, a hang-off error of approximately 4 knots develops. The heading tracking of the constant runway heading command is within +1 °. The hang-off during decelerations and small errors during segments with constant velocity and heading commands are characteristics of the Type 1 property predicted by theory for the controller's structure. The vertical velocity command response exhibits an undesirable low damped oscillation with peak errors of + 1.27 m s-1 (+ 250fpm). Further refinements of the vertical control and guidance algorithms are required but were not possible due to the limited flight test time. The nominal trajectory used in this evaluation (Fig. 1) requires the vehicle-controller combination to operate at and transition between the four trim conditions given in Table 2. This is automatically accomplished by the control algorithm without any knowledge of the theoretical trim values for the controls or vehicle attitudes. Figure 6 verifies the theoretical trim values and thus the model developed by Ostroff et al. (1976) used for the trim pitch attitude and collective control. An evaluation of the total differential collective control is predicted by the fact that only the incremental change about an unknown initial condition was recorded. A comparison of the incremental changes listed in
299
Helicopter digital controller flight test
80 ~ ALONGTRACK VELOCITY40 ~ COMMAND, knots 0L
J
~
~
ALONGTRACK VELOCITY 10 ~ ~ _ . . ~ . ERROR, knots 0 -10 I0 r CROSSTRACK VELOCITY , L O COMMAND, knots -I0 L_
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CROSSTRACK VELOCITY 5 ERROR, knots 0 -5 VERTICAL VELOCITY 1000 COMMAND, ft/min O -I000 250 F VERTICAL VELOCITY ERROR, ft/min 0 ' ~ v ~ V ~wvV~ V ~ X ~ v " ~ -250 10 HEADING ERROR, deg .0 I I I 1 -10 r i 0 20 40 60 80 I00 TI ME, sec
300 0 m/min -300 75
~
0 m/min -75
I
120
FIG. 5. Typical velocity and heading tracking time history. TABLE 2. C H - 4 7 B NOMINAL TRIM CONDITIONS
(s)
Longitudinal velocity (knots)
Descent Rate (m s- 1) (ft rain- 1)
0
65
0
20
65
3.21 (600)
80
0
0
100
0
0.48 (90)
Nominal time
3.50 r-
c THEORETICALTRIM VALUES
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COLLECTIVE, in.
CO
I,,
~I lO t
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~
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-5
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,
,
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20
40
60 TIME. sec
80
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1° 5
0
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0
120
FIG. 6. Typical longitudinal controls and pitch time histories.
Trim differential collective CM (inches) - 1.94
(-0.725) -2.33 (-0.869) - 0.78 (-0.292) - 0.79 (-0.295)
Trim collective CM (inches) 8.58 (3.21) 6.93 (2.59) 11.82 (4.42) 11.69 (4.37)
Trim pitch angle (deg.) 2.94 3.14 6.61 6.61
Table 2 between t = 0 and t = 80 s and the changes observed in flight show the theoretical change to b ¢ approximately one-half as large as that observed in flight. This indicates an inaccuracy in the model. The use of the Type 1 structure automatically corrects for this type of uncertainty. The autoland position tracking task includes runway centerline tracking and altitude nominal profile tracking during the approach to hover and lateral and longitudinal deviation from the center of the landing pad at touchdown. Figure 7 shows the excellent position tracking performance of the autoland system during approach to hover. The initial deviation from centerline is reduced to less than _+6.1m (_+20 feet) within 30s. The
300
D . R . DOWNING and W. H. BRYANT
CROSSTRACK~ F~-- ~ ERROR
li
u ~-
.
i
.
~
=
.
-
.
~
.
=
.
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TIME, sec
FIG. 7. Autoland system position tracking capability.
maximum altitude errors of 3.66m (12 feet) occur at the initiation of the descent segment. The altitude error at hover is 2.74 m (9 feet). The mean touchdown dispersion, based on nine landings, is 2.93 m (9.6 feet). If it were necessary to reduce the touchdown dispersion, position correction guidance could easily be incorporated in the landing guidance mode. 6. CONCLUSION An advanced low-iteration-rate (10Hz) gainscheduled digital controller based on modern control theory was designed for the Boeing Vertol CH-47 tandem rotor helicopter. This controller tracked commanded longitudinal, lateral and vertical velocity and vehicle heading; when combined with dosed-loop guidance and nominal trajectory generation algorithms, it formed a helicopter autoland system. The autoland system was flight tested along a descending deceleration profile to hover followed by a vertical descent to touchdown. Based on nine automatic landings, the autoland system is shown to have excellent position tracking performance with maximum altitude and lateral tracking
errors, after initial condition errors are reduced, of 3.66m (12 feet) and 6.1 m (9.6 feet). The velocity tracking performance is good with the longitudinal and lateral velocity tracking less than + 2 knots except during longitudinal deceleration when a 4 knot hang-off error develops. Heading tracking is excellent, being less than 1°. The vertical velocity controller exhibited a lightly damped oscillation with peak values of 1.27 m s- 1 (250 fpm) that would require further refinement. The flight tests have verified the digital controller design procedure, as well as the autotrim and the Type 1 features of the longitudinal controller predicted by theory. Also demonstrated is the feasibility of implementing advanced autoland systems in state-of-the-art flight computers using a higher order language. REFERENCES Kelly,J. R., F. R. Niessen,J. R. Garren and T. S. Abbott (1979). The VTOL approach and landing technology (VALT) CH-47 research system. NASA TP-1436. Moen, G. C., D. J. DiCarlo and K. R. Yenni(1976).A parametric analysis of visual approaches for helicopters. NASA TN D-8275. Moen, G. C. and K. R. Yenni (1975). Simulation and flight studies of an approach profile indicator for VTOL aircraft. NASA TN D-8051. Ostroff,A. J., D. R. Downingand W. J. Rood (1976).A technique using a nonlinear helicoptermodel for determiningtrims and derivatives. NASA TN D-9159. Stengel, R. F., J. R. Broussard and P. W. Berry (1978a). Digital flight control design for a tandem-rotor helicopter. Automatica, 14, 301-312. Stengel, R. F., J. R. Broussard and P. W. Berry (1978b). Digital controllersfor VTOL aircraft.IEEE Trans. Aerospace Electron. Syst., AES-14, 54-63. Stengel, R. F., J. R. Broussard and P. W. Berry (1979). Modern digital flight control systemdesign for VTOL aircraft. NASA CR 159019.