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On the Development of Attitude Stabilization and Control System of a Japanese Scientific Satellite
CASE STUDIES IN SPACE
Copyright © IFAC Control Science and Technology (8th Triennial World Congress) Kyoto , Japan , 1981
ON THE DEVELOPMENT OF ATTITUDE ST ABILIZATION AND CONTROL SYSTEM OF A JAPANESE SCIENTIFIC SATELLITE K. Ninomiya Institute of Space and Aeronautical Science, University of Tokyo, Tokyo , Japan
Abstract. ISAS' experience and future plan in attitude control of scientific satellites are briefly reviewed. The mission and the attitude requirement are explained for scientific satellite EXOS-C that is scheduled for launch in February 1984 by Mu-3S vehicle. The attitude control is implemented by a bias-momentum stabilized Sun-pointing system employing magnetic torquing. The detail of the design is presented, and the development and test approach of the control system is described in relation to the Japanese scientific satellite project at ISAS, Univ. of Tokyo. Keywords. Attitude control; satellites, artificial; sensors; actuators; control system analysis. HISTORICAL BACKGROUND AND FUTURE PLAN Since the launch of Japan's first satellite "OHSUMI" in February 1970 by Lambda-4S, a simulation-test vehicle for developing Mu-4S, Institute of Space and Aeronautical Science, University of Tokyo has been responsible for developing scientific satellites and their launch vehicles (Mu-series). Table 1 lists the past and future satellites of ISAS until fiscal year 1985 (authorized) and beyond. The missions of satellites (TANSEI-I, SHINSEI, and DENPA) launched by the first generation launch vehicles (Mu-4S's) were so chosen that they did not require the attitude control capability, and the satellites were left spinning at the rate given by the last stage motor. Fluxgate-type magnetometers and a sun sensor were used to determine the spin-axis direction. The sun sensors made use of CdS photodetector and operated on potentiometer principle, and both sensors were derivatives from Sounding Rocket Project.
that were going to be applied to scientific satellites planned for launch by the vehicles of the same generation. Thus, on satellites launched by Mu-3C's (TANSEI-![, TAIYO, HAKUCHO) magnetic torquing technology (spin axis and rate control---open loop) was mastered, and together with yo-yo despinner and viscous nutation damper, successfully applied for mission accomplishments. For TAIYO, it was to keep "wheel mode" at 10 rpm spin, and for HAKUCHO it was used for the reorientations of spin axis direction and the adjustments of spin rate.
On these satellites, spin-rate decay due to eddy current was identified. An attempt had been made on DENPA to keep the spin-axis direction within the favourable range relative to Sun's ray by trimming satellite's magnetic moment by magnet, but the early failure in orbit of the satellite due to high voltage discharge did not allow to witness the effectiveness. Though nutation damper was not adopted except on DENPA, stable spin was maintained in orbit.
On these satellites, Gray-cord array of photodetector elements fabricated on one silicon -chip by integrated circuit technology was used as digital sun sensor. Horizon crossing indicators using pyloelectric PbTi03 detectors were developed and used extensively. These resulted in the improvement of performance and the attitude determination-accuracy sufficient for mission purposes, though on HAKUCHO the experiment (X-ray modulation collimator) itself could actually be used as a finer attitude sensor. On HAKUCHO (and also TANSEI-N), a very small nutations caused by the operation mode-change of the on-board data recorder and their decay by nutation damper were detected by the precise measurement of the spin-period by sun sensor. This was reflected to the modification of structural configuration in recorders to be used on HINOTORI etc. so as to minimize the effect.
Through the course of developing launch vehicles, ISAS established an approach of launching test satellite by the first vehicle of every generation and to flight-test and qualify as much as possible the satellite technologies
By using TANSEI-llI, we could develop an active magnetic stabilization system that makes an axis of satellite follow the local geomagnetic field vector in orbit. This was successfully applied to KYOKKO to take UV
2447
244 8
K. Ninomi ya
pictures of aurora borealis over the north pole. Also on TANSEI-llI a cold gas-jet system was developed and tested to master the techniques of precessing and nutation-damping of a simple spinner by axial jets. JIKIKEN had four antennas of 60 meters long each, but the attitude was not controlled. Unable to deplo y completel y equal lengths of all antenna elements, the satellite sustained complicated nutation-like motion which is yet to be anal yzed. The fourth generation of Mu; The Mu-3S's have launched 2 satellites and will launch another 2 (TANSEI-lY, HINOTORI, ASTRO-B, and EXOS-C). Thus, the objective of TANSEI-lY was to verify overall performance of Mu-3S vehicle and to perform several tests and experiments on engineering instruments and techniques to be applied for three satellites to follow. In terms of attitude requirements, HINOTORI must be offset pointed to Sun with offset angle between 0.7 0 and 1.7 0 and at about S rpm spin of minimal nutation. In case of ASTROB, driftless low-spin-rate scan (~ 0.1 rpm) of sky is required, and the spin axis must be maneuverable to arbitrary points on the celestial sphere with some restrictions. For EXOS -C mission, it can be stabilized at zero spin with the body-fixed solar cell panels facing the Sun. From this stand-point, a momentum wheel with built-in horizon scanner (SCAN WHEEL) was introduced into attitude control system of TANSEI-lY to control the spacecraft spin-rate or pitch-angle. Also, an autonomous offset-pointing control of the momentum axis with respect to the Sun's ray was implemented by relying upon the magnetic torquing technique. As an attitude sensor, a N-slit star scanner using a photomultiplier tube was newly added to our TABLE 1
Among those not included in the test by TANSEI-lY are rate integrating gyro sensor and thermal louver etc., that are being developed and qualified on the ground. Also, the schemes such as to alleviate the attitude drift of ASTRO-B due to solar radiation pressure, or to effectively damp the nutational motion of EXOS-C will have to be applied after the evaluation by analyses. The effects of air drag and gravity gradient torques and flexibility of the four 20-m antennas of EXOS-C must also be precisely evaluated. PLANET-A is going to be the first interplanetary science probe of Japan and aims at the UV imagery of the Comet Halley. MS-TS is the test spacecraft for confirming the required technology. They are both spinners having a despun antenna but will also incorporate a momentum wheel for the fulfilment of the mission. Hydrazine jet system for orbit and attitude correction is now under development. ASTRO-C is in the preliminary stage of design study, but it seems that inertial and star sensors in combination with onboard microprocessors will be exploited profitably on this spacecraft which is going to be the heaviest and most complex of ISAS' satellites that have been authorized up to now. EXOS-C SPACECRAFT AND ATTITUDE REQUIREMENT EXOS-C is planned as an activity in MAP Program and has the following mission objectives: (A) remote sounding of minor consitutents in the middle atomosphere and (B) study on the wave-particle interactions in ionospheric plasma over the Brazilian geomagnetic anomaly
Scientific and Test Satellites of ISAS
Satelli te
Launch Vehicle
Date (DIM/y)
OH SUM I TANSE1 SHINSE1 DENPA TANSEI-II TAIYO TANSEI-III KYOKKO JIK1KEN HAKUCHO TANSEI-IV HINOTORI
L-4S-5 M-4S-2 M-4S-3 M-4S-4 M-3C-1 M-3C-2 M-3H-1 M-3H-2 M-3H-3 M-3C-4 M-3S-1 M-3S-2
11/2(' 70 16/2/'71 28/9/'71 19/8/'72 16/2/'74 24/2/'75 19/2/'77 4/2/'78 16/9/'78 21/2/'79 17/2/'80 21/2/'81
ASTRO-B EXOS-C MS-T5 PLANET-A ASTRO-C MS-T6 EXOS-D
M-3S-3 M-3S-4 M-3S.kai. 1-1 M- 3S. kaL 1-2 M-3S.kaLI-3 M-3S.kaL II-1 M-3S.kaLII-2
Feb. '83 Feb. '84 Jan. '85 Aug. '85 Feb. '86 -----
----
repertoire with the purpose of using the same type on ASTRO-B.
-----------
---------
Orbit h~(km)/hf(km)/i(deg)
5,142(35 1f3 1 1,108/990/30 1,870/874/32 6,566/246/31 3,237/289/31 3,136/255/32 3,813/791/66 3,976/637/65 30,051/227/31 576/545/30 605/521/39 643/577/31
Weight (kg) 24 63 66 75 56 86 129 126 92 96 185 190
500 circular/30 1 ,000/350/74 interplanetary interplanetary 500 circular/30
,,200 ,,220 .. 125 ,,125 ,,400
-------------------------------------
-------
----
Main Mission launch test engineering test ionosphere study plasma wave in space engineering test ionosphere study engineering test auroral physics magnetospheric study X-ray astronomy engineering test solar physics X-ray astronomy mid. atmosphere study engineering test Halley's comet probe X-ray astronomy engineering te at magneto spheric study
-------------------
Development of Attitude Stabilization and the auroral zone. To carry out these objectives, EXOS-C has scientific payloads; (Al) limb scanning radiometer for 1.27 ~m band airglow to measure the ozone concentration, (A2) spectrometer to measure the backscattered ultraviolet radiance from the terrestrial atmosphere to deduce global distribution of ozone density, (A3) solar image-radiometer in visible and nearinfrared bands for the limb absorption by stratospheric aerosols, ozone and nitrogen dioxide, (A4) infrared solar spectrometer for the limb absorption by stratospheric water vapor, methane, and nitrogen dioxide, (Bl) topside ionospheric plasma sounder, including a receiver for the emissions from electric power lines, (B2) plasma probes for electron density and temperature, (B3) energy spectrum -analyser for low energy precipitating particles, and (B4) flux monitor for high energy precipitating particles. EXOS-C will be launched by Mu-3S-4 in February 1984 into an elliptical orbit, as shown in Table 1. The spacecraft weighs about 220 kg, is an octagonal tube-shape of 1.0 meter wide in diagonal, and is powered by four solar panels which, when deployed, supplies the spacecraft power of 140 watts at the beginning of life. It is also mounted with four extendable and retractable antennas of Experiment (Bl) of 20 meters long each, as shown in Fig. 1. The spacecraft attitude is three-axis stabilized, with angular momentum (Z-) axis pointed to the Sun and the transverse (X- and Y-) axes inertially oriented to the commanded direction within the plane perpendicular to the Sun.
PPS EXTENDABLE ANTENNA
TO
TrX~
THERMAL LOUVER
Fig. 1.
Configuration of EXOS-C spacecraft
The Sun pointing of Z-axis is required by experiments (Al) (A3) and (A4), and also desirable from the standpoint of spacecraft power and thennal design. The pointing accuracy required is + 0.5°. Control capability of
2449
transverse axes are required by the experiment (A-2) with accuracy of +3°. Short period jitter-like motion of attitude should be minimized. Also nutation, if occurred, should be damped out as quickly as possible. The after -the-fact attitude determination accuracy shall be about +0.1° for the picture data processing of experiment (A-3). Experiments categorized as (B- )do not have the specific attitude requirements. The attitude control system must be provided with initial attitude -acquisition capability. ATTITUDE CONTROL SYSTEM DESCRIPTION Overview of The Attitude Control System: To meet the requirements above, the attitude control of EXOS-C is based on the angular momentum -bias stabilization and magnetic torquing technology. The system utilizes a momentum-wheel with built-in IR earth-scanner (SWA), magnetic torqllers (MAC and MULD) , a two dimensional fine sun-sensor (TFSS), a rate-integrating gyro-package (RIG), and a geomagnetic fieldaspect sensor (GAS) composed of 3-axis fluxgate magnetometers. A digital sun-aspect sensor (DSAS) for spinning phase, a passive nutation damper (ND) , and a yo-yo despinner supplement the system. Fig. 2 shows the overall block diagram of EXOS-C attitude control system, and Fig. 3 indicates the locations of attitude sensors and actuators. Initial Attitude Acquisition: After the deployment of yo-yo despinner and solar cellpanels in orbit, the wheel is spun-up and then the minus Z-axis of the spacecraft is maneuvered from the injected attitude toward the Sun's direction in open loop manner, where the polarity of current in the magnetic attitude control (MAC) coil is computed on the ground and stored/executed on board. When Z-axis approaches the predetermined vicinity of Sun's direction, the control is transferred to the autonomous mode. The four 20-meter antennas of Experiment (Bl) can be commanded to extend while the transverse-axes control is operated in the constant-wheel-speed mode or the ratedamping mode. Z-axis Control Loop: TFSS detects the two dimensional angular offset of Z-axis from Sun's direction. These signals, together with the geomagnetic field components from GAS, are processed in the attitude control electronics (ACE) to generate the coil current for controlling Z-axis in a bang-bang manner by either one of MAC coil (wound perpendicularly to Zaxis) or MULD coil (wound in the plane perpendicular to Y-axis). For orbit and attitude of EXOS-C, it is necessary to make use of these two coils to obtain adequate control torque by magnetic means. Transverse Axes Control Loop: Floated rateintegrating gyropackage (RIG) outputs the pulse train that corresponds to the rotational rate of the spacecraft about Z-axis. This is integrated in wheel control electronics (WCE) to give the rotation angle of the transverse
K. Ninomiya
2450
axes from a commanded reference direction. By using these angle and rate informations, the torque command for SWA is generated so as to reduce the angular deviation. The reference direction can be updated through RF command link from the ground station. The accumulated error due to RIG drift will be detected on the ground by using the data from IR earth scanner in SWA, and then compensated through RF command link. Other modes of operation are also implemented into the loop, i.e., constant-wheel-speed mode and rate-damping mode. From the mission-life point of view, the gyropackage is given a complete redundancy.
put from WCE. During the launch and the spin phase in orbit, DSAS also is used for attitude and spin-rate measurement.
ULO Coil Y-AXIS
I
X-AXIS
~C
Coil
~FSS
FOV of TFSS 64x64 (deg)
o o
\7 <>
Nutation Teleme-try comman[J0amper __~----~o x2
C)
Z) Yo-Yo
Oespiner
Command
Fig. 2.
Block diagram of attitude control system.
Unloading of the wheel momentum is automatically accomplished as well as manually by using magnetic unloading(MULD) coil under the control of ACE and based upon the outputs of wheel tachometer and magnetometers. Nutation Damping: EXOS-C has structural configuration stable in motion around Z-axis and is equipped with a viscous nutation damper(ND). Due to the long damping time-constant of ND, however, it is highly desirable to have active nutation control(ANC) means especially for large nutations. Initially, ANC by productof-inertia utilization was considered, in which cant-mounted RIG output-signal be used to drive SWA reducing the nutation. To get effective damping, however, each of four long antennas had to be extendable into a specified direction that finally was found unacceptable to (Bl) Experiment. As a consequence, it has been decided to adopt a magnetic ANC which uses MAC coil as actuator. ACE is so designed that it functions as both Z-axis and ANC controller. TFSS is used as detector of nutation which exceeds the resolution limit of it. Attitude Determination: Attitude determination is carried out on the ground by processing data from TFSS, IR earth scanner, and GA as well as the integrated-rate angle that is out-
Fig. 3.
Allocation of sensors and actuators.
Attitude Sensors and Actuators: The following is the brief description of the sensors and actuators in the attitude control system of EXOS-C spacecraft; RIG measures angular rate around Z-axis and is composed of two channels of rebalance electronics and single degree-of-freedom floated rate-integrating gyroscope. The rate range is + 0.44 rpm or + 1.8 rpm, scale factor of output pulse is 3-arcsec/pulse or 0.75 arcsec/pulse, cut-off frequency is 10 Hz, and non-G random drift is less than 0.01 deg./ hour. Expected operational life is greater than one year. TFSS gives two dimensional sun-angles within the field of view of 32°x 32°. The output resolution is 0.002° and the accuracy is 0.05°. DSAS not only measures the sun-aspect angle with the accuracy of 1°, but also outputs the sun-incidence pulse when sun's ray crosses the fan-beam field-of-view of l28°x 2°. GAS measures three components of the geomagnetic field with an accuracy of + 200 nT in the range of ± 50,000 nT. These-are used as inputs to Z-axis control loop, momentum unloading loop, and magnetic nutation control loop as well as for coarse attitude determination. SWA is not only for providing a biased angular momentum and controlling body angle but also for obtaining the attitude data from the built-in IR earth scanner. SWA outputs earth pulse and body-reference pulse. WCE processes these, giving the earth pulse width
Dev e lopment of Attitude Stabiliz a ti on and the phase difference between earth pulse and body-reference pulse. The nominal angular momentum provided by SWA is 18.4 N-m-s (at 2000 rpm wheel speed). Maximum reaction torque is 0.035 N-m. The IR scan-cone angle is 45°, and the accuracy we expect, when useed for attitude determination of EXOS-C, is in the order of +0.1°. MAC is an air-core coil wound in a plane perpendicular to Z-axis. This is used for precession control of Z-axis and for magnetic ANC. The maximum magnetic moment of MAC is set to 30 ampere-m 2 • MULD is wound parallel to X-Z plane. This is utilized for momentum unloading and for precession control of Z-axis. The maximum magnetic moment is 10 ampere-m 2 • ND of EXOS-C is a rectangular pipe filled with silicon oil. The calculated time constant of damping at the steady state operational phase is 1.3 hours. It is to be augmented by the magnetic active nutation control . Yo-yo despinner is the radial release type that will reduce the spin of 120 rpm given by the 3rd stage of Mu-3S to 6 rpm. After this the solar panel-deployment and the wheel spin-up will reduce the body rate to nearly zero rpm.
245 1
curves of estimated air drag torque versus mean anomaly, that is caused by a single element of four antennas. Though the values in Fig. 5 is very large in the vicinity of perigee, the net torque due to the air drag will be crossed out at least as the first order effect. This is because of the aligned configuration of antenna elements and the small distance between spacecraft's center of mass and the plane in which antennas lie. The moment of inertia values (in kg-m 2 ) used in the calculation are; Ix = I y = 118 and I z = 235; and the characteristics of an antenna element are: length 19.8 m, diameter 1.27 x 10- 2 m, mass densit y 2 . 27 X 10- 2 kg/m, bending stiffness 6.54 N-m 2 •
...
...
v7 V2
...
V'Min Disturbance Attitude
.. ...
VzMax Disturbance V:Velocity vector at the Perigee
... L.~
_D_
.
__~____________________________~~__4 N
THE DESIGN Fig.5. Disturbance Torques: Gravity gradient torque and air drag torque due to the long antennas are the significant disturbances to the control system. To evaluate the order of magnitude, Fig 4 shows a plot of magnitude of the maximum gravity gradient torque in orbit as a function of the angle s subtended by Zaxis (Sun's direction) and the direction of orbital ascending node. The orbit is assumed as a circular one having a radius equal to the average of the actual orbit.
loo OC,.. ) .. Ul." t1' ()(gM ) .. 117.7 1. ()(~ ) .. 23!5..4
'<-.,>
i a. a..-iW
lit
(;(.)
T~
1
Air-drag torque on an antenna element.
Flexibility of Antennas: To analyze the effect of the flexibility of four 20 meter antennas, the hybrid coordinate method has been employed. By modeling antennas as cantilevers, we can obtain the vibration modes in analytic forms. There are 6 first order modes. To the first order effect, QI and Q3 do not influence the attitude motion, Q2 and Q4 couple to Z-axis control system as shown in Fig.6, and Q5 and Q6 couple into transverse axes control system. Table 2 lists the hybrid modal data.
-7"]..~
.. 7l13li
c..,"') ..
rT1
i
t, e.. z.,-1iIA
..........
...
1
......... 1 I
+________________________________ ____
L~~~I
Fig. 4.
.
.
~
:
~
!
ro
IZ I
Gravity-gradient torque.
Based on a simplified model, Fig. 5 shows eST 4 _
IZ-w.,>
Fig. 6.
Vibration modes of EXOS-C antennas.
2452
K. Ninomiya
TABLE 2 modal no. 1st 2 order ~ 5 6 2nd 8 order 10 11 12 3rd 14 order 16 l 'I
18
Hybrid Modal Data angular freq. i rad.bee. ) 0 .162
1.')3
8 .11
eOURl iI1ii eoeff.
cx
ey
cz
0 0 35.56 0 0 - 35 .56 0 0 95.56 0 0 ! 9').')6 0 1 0 15.24 0 1-15.24 I ::J , o 0\15.24 0 0 15.24 0 , 0 5.44 i 0 0 -5.44 0 ! 5.44 0 i ! 0 0 5.44
Z-Axis Control Logic: In autonomous sunpointing mode of operation, the polarity of electric current in either MAC or MULD coils is swithed automatically to generate a control torque that will tend to reduce the deviation of Z-axis from the solar vector until it reaches in a predetermined small region. Fig. 7 is the block diagram for Z-axis control logic.
I
I
The nutation can be detected by using TFSS if one notices that X-component of TFSS output is expressed in the form; a = S-Ch/I) w, where S is the component of solar vector projected on to X-Z plane in the absence on nutation, h is the angular momentum of SWA, I is the transverse moment of inertia, and w is the X-component of angular velocity. In practice the d.c.-cut X-output of TFSS is fed into the logic circuit to generate the damping torque. The nutation damping rate is given by T/C nh) in terms of the magnetic torque T produced by MAC coil. For EXOS-C, this rate is estimated to be around 2°/hour. Transverse Axes Control Loop: As shown in Fig. 8, the system has three modes of operation; normal operation mode, rate-damping mode, and constant-wheel-speed mode. The rate -damping mode will have to be used in cases of gyro-digitizer or integrating-counter failures. The constant-wheel-speed mode will be employed when wheel is run up and long antennas .q re being extended or retracted. In the design of the system, the first order modes of flexibility are taken into account. In this case, the system has a gain margin greater than 20 dB, and phase margin of 40 degrees.
DYNAMICS Fig. 7.
Block diagram of magnetic torquers. Fig. 8.
In order to choose the adequate coil and to determine the current polarity, one needs to know the geomagnetic field relative to the spacecraft and also to determine the sign of the quantity; D = -SxBy + SyBx, where, Sx and Sy are respectively the X and Y components of the solar vector, and Bx and By are the X and Y components of the geomagnetic field. As a practical way of determining the sign of D without using an onboard computer, the X-Y plane is subdivided into 9 segments CA to I) as shown in Fig. 7. Segment I signifies the dead band. Corresponding to the combination of two segments in which the solar vector and the geomagnetic vector are projected respectively, the hard wired logic determines and issues the adequate control commands.
Block diagram of transverse axes control loop.
SIMULATION AND TEST To verify the design, computer simulations of the control system performance have been scheduled. In the simulation of Z-axis control, the effects of external disturbance torques and antenna flexibility will be carefully checked, as well as the performance in nutaion control. In transverse axes simulation, the gyro noise, nonlinear elements in the control electronics, and flexibility mode-couplings of higher modes will be paid attention to. Finally, coupled motion of Z-axis and transverse axes through the flexibility will be taken into account by using a simplified model for antennas.
Development of Attitude Stabilization As tests in developing the transverse axes control system, we started from the preliminary test to measure the parameters of SWA itself and tacho feed-back loop. Next step is to close the control loop by using the actual hardware except that the spacecraft dynamics is simulated in a minicomputer. As a confirmation of overall performance, the total system will be mounted on a spherical airbearing table and be operation-tested without simulating the moment of inertia nor flexibility. As for Z-axis control system, no dynamical test is conducted except for checks of the control logics and the polar ities and magnitudes of magnetic torquers. Acknowledgement. The author would like to express his gratitude to all Professors and technical staffs who have been actively taking part in the scientific satellite project at ISAS for their interest and support. Also, he wishes to acknowledge the cooperation of Messrs. T. Yamamoto, J. Aoyama, and K. Maeda of NEC in developing the attitude control system of EXOS-C. Discussion to Paper CS 7 .1 G. E. Cunningham (USA ) : Are the attitude control electronics impl e mented in hardwired logic or in software? K . Ninomiya (Japan): Th ey are implemented in h a rdwir ed logic. Mi c rop roc esso r s wi ll not be used for attitude control until 1986.
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P. Kant (Netherlands): In your servo test of EXOS-C ACS, did you measure the control torque of the scanning wheel to feed it into the simulated dynamiCS, or did you model the wheel dynamics? K. Ninomiya (Japan): We made the measurement on the wheel dynamics . In the simula tion which we call " mixed test" we modelled the wheel dynamiCS , according to the measure ment . J.W. Hursh (USA) : Is the EXO satellite a NASDA project? K. Ninomiya (Japan) : No. It is an ISAS project. ISAS manages and designs EXOS-C and other scientific satellites. F. Wojtalik (USA): Do your attitude control designs include a " fail - safe" automatic sun pOinting control feature? K. Ninomiya (Japan) : We have no provision except for an open loop mode of the switching of magnetic moment according to the program which will be commanded from the ground and stored in the program timer memory.
Copyright © IFAC Control Science and Technology (8th Triennial World Congress) Kyoto , Japan , 1981
ON THE DEVELOPMENT OF ATTITUDE ST ABILIZATION AND CONTROL SYSTEM OF A JAPANESE SCIENTIFIC SATELLITE K. Ninomiya Institute of Space and Aeronautical Science, University of Tokyo, Tokyo , Japan
Abstract. ISAS' experience and future plan in attitude control of scientific satellites are briefly reviewed. The mission and the attitude requirement are explained for scientific satellite EXOS-C that is scheduled for launch in February 1984 by Mu-3S vehicle. The attitude control is implemented by a bias-momentum stabilized Sun-pointing system employing magnetic torquing. The detail of the design is presented, and the development and test approach of the control system is described in relation to the Japanese scientific satellite project at ISAS, Univ. of Tokyo. Keywords. Attitude control; satellites, artificial; sensors; actuators; control system analysis. HISTORICAL BACKGROUND AND FUTURE PLAN Since the launch of Japan's first satellite "OHSUMI" in February 1970 by Lambda-4S, a simulation-test vehicle for developing Mu-4S, Institute of Space and Aeronautical Science, University of Tokyo has been responsible for developing scientific satellites and their launch vehicles (Mu-series). Table 1 lists the past and future satellites of ISAS until fiscal year 1985 (authorized) and beyond. The missions of satellites (TANSEI-I, SHINSEI, and DENPA) launched by the first generation launch vehicles (Mu-4S's) were so chosen that they did not require the attitude control capability, and the satellites were left spinning at the rate given by the last stage motor. Fluxgate-type magnetometers and a sun sensor were used to determine the spin-axis direction. The sun sensors made use of CdS photodetector and operated on potentiometer principle, and both sensors were derivatives from Sounding Rocket Project.
that were going to be applied to scientific satellites planned for launch by the vehicles of the same generation. Thus, on satellites launched by Mu-3C's (TANSEI-![, TAIYO, HAKUCHO) magnetic torquing technology (spin axis and rate control---open loop) was mastered, and together with yo-yo despinner and viscous nutation damper, successfully applied for mission accomplishments. For TAIYO, it was to keep "wheel mode" at 10 rpm spin, and for HAKUCHO it was used for the reorientations of spin axis direction and the adjustments of spin rate.
On these satellites, spin-rate decay due to eddy current was identified. An attempt had been made on DENPA to keep the spin-axis direction within the favourable range relative to Sun's ray by trimming satellite's magnetic moment by magnet, but the early failure in orbit of the satellite due to high voltage discharge did not allow to witness the effectiveness. Though nutation damper was not adopted except on DENPA, stable spin was maintained in orbit.
On these satellites, Gray-cord array of photodetector elements fabricated on one silicon -chip by integrated circuit technology was used as digital sun sensor. Horizon crossing indicators using pyloelectric PbTi03 detectors were developed and used extensively. These resulted in the improvement of performance and the attitude determination-accuracy sufficient for mission purposes, though on HAKUCHO the experiment (X-ray modulation collimator) itself could actually be used as a finer attitude sensor. On HAKUCHO (and also TANSEI-N), a very small nutations caused by the operation mode-change of the on-board data recorder and their decay by nutation damper were detected by the precise measurement of the spin-period by sun sensor. This was reflected to the modification of structural configuration in recorders to be used on HINOTORI etc. so as to minimize the effect.
Through the course of developing launch vehicles, ISAS established an approach of launching test satellite by the first vehicle of every generation and to flight-test and qualify as much as possible the satellite technologies
By using TANSEI-llI, we could develop an active magnetic stabilization system that makes an axis of satellite follow the local geomagnetic field vector in orbit. This was successfully applied to KYOKKO to take UV
2447
244 8
K. Ninomi ya
pictures of aurora borealis over the north pole. Also on TANSEI-llI a cold gas-jet system was developed and tested to master the techniques of precessing and nutation-damping of a simple spinner by axial jets. JIKIKEN had four antennas of 60 meters long each, but the attitude was not controlled. Unable to deplo y completel y equal lengths of all antenna elements, the satellite sustained complicated nutation-like motion which is yet to be anal yzed. The fourth generation of Mu; The Mu-3S's have launched 2 satellites and will launch another 2 (TANSEI-lY, HINOTORI, ASTRO-B, and EXOS-C). Thus, the objective of TANSEI-lY was to verify overall performance of Mu-3S vehicle and to perform several tests and experiments on engineering instruments and techniques to be applied for three satellites to follow. In terms of attitude requirements, HINOTORI must be offset pointed to Sun with offset angle between 0.7 0 and 1.7 0 and at about S rpm spin of minimal nutation. In case of ASTROB, driftless low-spin-rate scan (~ 0.1 rpm) of sky is required, and the spin axis must be maneuverable to arbitrary points on the celestial sphere with some restrictions. For EXOS -C mission, it can be stabilized at zero spin with the body-fixed solar cell panels facing the Sun. From this stand-point, a momentum wheel with built-in horizon scanner (SCAN WHEEL) was introduced into attitude control system of TANSEI-lY to control the spacecraft spin-rate or pitch-angle. Also, an autonomous offset-pointing control of the momentum axis with respect to the Sun's ray was implemented by relying upon the magnetic torquing technique. As an attitude sensor, a N-slit star scanner using a photomultiplier tube was newly added to our TABLE 1
Among those not included in the test by TANSEI-lY are rate integrating gyro sensor and thermal louver etc., that are being developed and qualified on the ground. Also, the schemes such as to alleviate the attitude drift of ASTRO-B due to solar radiation pressure, or to effectively damp the nutational motion of EXOS-C will have to be applied after the evaluation by analyses. The effects of air drag and gravity gradient torques and flexibility of the four 20-m antennas of EXOS-C must also be precisely evaluated. PLANET-A is going to be the first interplanetary science probe of Japan and aims at the UV imagery of the Comet Halley. MS-TS is the test spacecraft for confirming the required technology. They are both spinners having a despun antenna but will also incorporate a momentum wheel for the fulfilment of the mission. Hydrazine jet system for orbit and attitude correction is now under development. ASTRO-C is in the preliminary stage of design study, but it seems that inertial and star sensors in combination with onboard microprocessors will be exploited profitably on this spacecraft which is going to be the heaviest and most complex of ISAS' satellites that have been authorized up to now. EXOS-C SPACECRAFT AND ATTITUDE REQUIREMENT EXOS-C is planned as an activity in MAP Program and has the following mission objectives: (A) remote sounding of minor consitutents in the middle atomosphere and (B) study on the wave-particle interactions in ionospheric plasma over the Brazilian geomagnetic anomaly
Scientific and Test Satellites of ISAS
Satelli te
Launch Vehicle
Date (DIM/y)
OH SUM I TANSE1 SHINSE1 DENPA TANSEI-II TAIYO TANSEI-III KYOKKO JIK1KEN HAKUCHO TANSEI-IV HINOTORI
L-4S-5 M-4S-2 M-4S-3 M-4S-4 M-3C-1 M-3C-2 M-3H-1 M-3H-2 M-3H-3 M-3C-4 M-3S-1 M-3S-2
11/2(' 70 16/2/'71 28/9/'71 19/8/'72 16/2/'74 24/2/'75 19/2/'77 4/2/'78 16/9/'78 21/2/'79 17/2/'80 21/2/'81
ASTRO-B EXOS-C MS-T5 PLANET-A ASTRO-C MS-T6 EXOS-D
M-3S-3 M-3S-4 M-3S.kai. 1-1 M- 3S. kaL 1-2 M-3S.kaLI-3 M-3S.kaL II-1 M-3S.kaLII-2
Feb. '83 Feb. '84 Jan. '85 Aug. '85 Feb. '86 -----
----
repertoire with the purpose of using the same type on ASTRO-B.
-----------
---------
Orbit h~(km)/hf(km)/i(deg)
5,142(35 1f3 1 1,108/990/30 1,870/874/32 6,566/246/31 3,237/289/31 3,136/255/32 3,813/791/66 3,976/637/65 30,051/227/31 576/545/30 605/521/39 643/577/31
Weight (kg) 24 63 66 75 56 86 129 126 92 96 185 190
500 circular/30 1 ,000/350/74 interplanetary interplanetary 500 circular/30
,,200 ,,220 .. 125 ,,125 ,,400
-------------------------------------
-------
----
Main Mission launch test engineering test ionosphere study plasma wave in space engineering test ionosphere study engineering test auroral physics magnetospheric study X-ray astronomy engineering test solar physics X-ray astronomy mid. atmosphere study engineering test Halley's comet probe X-ray astronomy engineering te at magneto spheric study
-------------------
Development of Attitude Stabilization and the auroral zone. To carry out these objectives, EXOS-C has scientific payloads; (Al) limb scanning radiometer for 1.27 ~m band airglow to measure the ozone concentration, (A2) spectrometer to measure the backscattered ultraviolet radiance from the terrestrial atmosphere to deduce global distribution of ozone density, (A3) solar image-radiometer in visible and nearinfrared bands for the limb absorption by stratospheric aerosols, ozone and nitrogen dioxide, (A4) infrared solar spectrometer for the limb absorption by stratospheric water vapor, methane, and nitrogen dioxide, (Bl) topside ionospheric plasma sounder, including a receiver for the emissions from electric power lines, (B2) plasma probes for electron density and temperature, (B3) energy spectrum -analyser for low energy precipitating particles, and (B4) flux monitor for high energy precipitating particles. EXOS-C will be launched by Mu-3S-4 in February 1984 into an elliptical orbit, as shown in Table 1. The spacecraft weighs about 220 kg, is an octagonal tube-shape of 1.0 meter wide in diagonal, and is powered by four solar panels which, when deployed, supplies the spacecraft power of 140 watts at the beginning of life. It is also mounted with four extendable and retractable antennas of Experiment (Bl) of 20 meters long each, as shown in Fig. 1. The spacecraft attitude is three-axis stabilized, with angular momentum (Z-) axis pointed to the Sun and the transverse (X- and Y-) axes inertially oriented to the commanded direction within the plane perpendicular to the Sun.
PPS EXTENDABLE ANTENNA
TO
TrX~
THERMAL LOUVER
Fig. 1.
Configuration of EXOS-C spacecraft
The Sun pointing of Z-axis is required by experiments (Al) (A3) and (A4), and also desirable from the standpoint of spacecraft power and thennal design. The pointing accuracy required is + 0.5°. Control capability of
2449
transverse axes are required by the experiment (A-2) with accuracy of +3°. Short period jitter-like motion of attitude should be minimized. Also nutation, if occurred, should be damped out as quickly as possible. The after -the-fact attitude determination accuracy shall be about +0.1° for the picture data processing of experiment (A-3). Experiments categorized as (B- )do not have the specific attitude requirements. The attitude control system must be provided with initial attitude -acquisition capability. ATTITUDE CONTROL SYSTEM DESCRIPTION Overview of The Attitude Control System: To meet the requirements above, the attitude control of EXOS-C is based on the angular momentum -bias stabilization and magnetic torquing technology. The system utilizes a momentum-wheel with built-in IR earth-scanner (SWA), magnetic torqllers (MAC and MULD) , a two dimensional fine sun-sensor (TFSS), a rate-integrating gyro-package (RIG), and a geomagnetic fieldaspect sensor (GAS) composed of 3-axis fluxgate magnetometers. A digital sun-aspect sensor (DSAS) for spinning phase, a passive nutation damper (ND) , and a yo-yo despinner supplement the system. Fig. 2 shows the overall block diagram of EXOS-C attitude control system, and Fig. 3 indicates the locations of attitude sensors and actuators. Initial Attitude Acquisition: After the deployment of yo-yo despinner and solar cellpanels in orbit, the wheel is spun-up and then the minus Z-axis of the spacecraft is maneuvered from the injected attitude toward the Sun's direction in open loop manner, where the polarity of current in the magnetic attitude control (MAC) coil is computed on the ground and stored/executed on board. When Z-axis approaches the predetermined vicinity of Sun's direction, the control is transferred to the autonomous mode. The four 20-meter antennas of Experiment (Bl) can be commanded to extend while the transverse-axes control is operated in the constant-wheel-speed mode or the ratedamping mode. Z-axis Control Loop: TFSS detects the two dimensional angular offset of Z-axis from Sun's direction. These signals, together with the geomagnetic field components from GAS, are processed in the attitude control electronics (ACE) to generate the coil current for controlling Z-axis in a bang-bang manner by either one of MAC coil (wound perpendicularly to Zaxis) or MULD coil (wound in the plane perpendicular to Y-axis). For orbit and attitude of EXOS-C, it is necessary to make use of these two coils to obtain adequate control torque by magnetic means. Transverse Axes Control Loop: Floated rateintegrating gyropackage (RIG) outputs the pulse train that corresponds to the rotational rate of the spacecraft about Z-axis. This is integrated in wheel control electronics (WCE) to give the rotation angle of the transverse
K. Ninomiya
2450
axes from a commanded reference direction. By using these angle and rate informations, the torque command for SWA is generated so as to reduce the angular deviation. The reference direction can be updated through RF command link from the ground station. The accumulated error due to RIG drift will be detected on the ground by using the data from IR earth scanner in SWA, and then compensated through RF command link. Other modes of operation are also implemented into the loop, i.e., constant-wheel-speed mode and rate-damping mode. From the mission-life point of view, the gyropackage is given a complete redundancy.
put from WCE. During the launch and the spin phase in orbit, DSAS also is used for attitude and spin-rate measurement.
ULO Coil Y-AXIS
I
X-AXIS
~C
Coil
~FSS
FOV of TFSS 64x64 (deg)
o o
\7 <>
Nutation Teleme-try comman[J0amper __~----~o x2
C)
Z) Yo-Yo
Oespiner
Command
Fig. 2.
Block diagram of attitude control system.
Unloading of the wheel momentum is automatically accomplished as well as manually by using magnetic unloading(MULD) coil under the control of ACE and based upon the outputs of wheel tachometer and magnetometers. Nutation Damping: EXOS-C has structural configuration stable in motion around Z-axis and is equipped with a viscous nutation damper(ND). Due to the long damping time-constant of ND, however, it is highly desirable to have active nutation control(ANC) means especially for large nutations. Initially, ANC by productof-inertia utilization was considered, in which cant-mounted RIG output-signal be used to drive SWA reducing the nutation. To get effective damping, however, each of four long antennas had to be extendable into a specified direction that finally was found unacceptable to (Bl) Experiment. As a consequence, it has been decided to adopt a magnetic ANC which uses MAC coil as actuator. ACE is so designed that it functions as both Z-axis and ANC controller. TFSS is used as detector of nutation which exceeds the resolution limit of it. Attitude Determination: Attitude determination is carried out on the ground by processing data from TFSS, IR earth scanner, and GA as well as the integrated-rate angle that is out-
Fig. 3.
Allocation of sensors and actuators.
Attitude Sensors and Actuators: The following is the brief description of the sensors and actuators in the attitude control system of EXOS-C spacecraft; RIG measures angular rate around Z-axis and is composed of two channels of rebalance electronics and single degree-of-freedom floated rate-integrating gyroscope. The rate range is + 0.44 rpm or + 1.8 rpm, scale factor of output pulse is 3-arcsec/pulse or 0.75 arcsec/pulse, cut-off frequency is 10 Hz, and non-G random drift is less than 0.01 deg./ hour. Expected operational life is greater than one year. TFSS gives two dimensional sun-angles within the field of view of 32°x 32°. The output resolution is 0.002° and the accuracy is 0.05°. DSAS not only measures the sun-aspect angle with the accuracy of 1°, but also outputs the sun-incidence pulse when sun's ray crosses the fan-beam field-of-view of l28°x 2°. GAS measures three components of the geomagnetic field with an accuracy of + 200 nT in the range of ± 50,000 nT. These-are used as inputs to Z-axis control loop, momentum unloading loop, and magnetic nutation control loop as well as for coarse attitude determination. SWA is not only for providing a biased angular momentum and controlling body angle but also for obtaining the attitude data from the built-in IR earth scanner. SWA outputs earth pulse and body-reference pulse. WCE processes these, giving the earth pulse width
Dev e lopment of Attitude Stabiliz a ti on and the phase difference between earth pulse and body-reference pulse. The nominal angular momentum provided by SWA is 18.4 N-m-s (at 2000 rpm wheel speed). Maximum reaction torque is 0.035 N-m. The IR scan-cone angle is 45°, and the accuracy we expect, when useed for attitude determination of EXOS-C, is in the order of +0.1°. MAC is an air-core coil wound in a plane perpendicular to Z-axis. This is used for precession control of Z-axis and for magnetic ANC. The maximum magnetic moment of MAC is set to 30 ampere-m 2 • MULD is wound parallel to X-Z plane. This is utilized for momentum unloading and for precession control of Z-axis. The maximum magnetic moment is 10 ampere-m 2 • ND of EXOS-C is a rectangular pipe filled with silicon oil. The calculated time constant of damping at the steady state operational phase is 1.3 hours. It is to be augmented by the magnetic active nutation control . Yo-yo despinner is the radial release type that will reduce the spin of 120 rpm given by the 3rd stage of Mu-3S to 6 rpm. After this the solar panel-deployment and the wheel spin-up will reduce the body rate to nearly zero rpm.
245 1
curves of estimated air drag torque versus mean anomaly, that is caused by a single element of four antennas. Though the values in Fig. 5 is very large in the vicinity of perigee, the net torque due to the air drag will be crossed out at least as the first order effect. This is because of the aligned configuration of antenna elements and the small distance between spacecraft's center of mass and the plane in which antennas lie. The moment of inertia values (in kg-m 2 ) used in the calculation are; Ix = I y = 118 and I z = 235; and the characteristics of an antenna element are: length 19.8 m, diameter 1.27 x 10- 2 m, mass densit y 2 . 27 X 10- 2 kg/m, bending stiffness 6.54 N-m 2 •
...
...
v7 V2
...
V'Min Disturbance Attitude
.. ...
VzMax Disturbance V:Velocity vector at the Perigee
... L.~
_D_
.
__~____________________________~~__4 N
THE DESIGN Fig.5. Disturbance Torques: Gravity gradient torque and air drag torque due to the long antennas are the significant disturbances to the control system. To evaluate the order of magnitude, Fig 4 shows a plot of magnitude of the maximum gravity gradient torque in orbit as a function of the angle s subtended by Zaxis (Sun's direction) and the direction of orbital ascending node. The orbit is assumed as a circular one having a radius equal to the average of the actual orbit.
loo OC,.. ) .. Ul." t1' ()(gM ) .. 117.7 1. ()(~ ) .. 23!5..4
'<-.,>
i a. a..-iW
lit
(;(.)
T~
1
Air-drag torque on an antenna element.
Flexibility of Antennas: To analyze the effect of the flexibility of four 20 meter antennas, the hybrid coordinate method has been employed. By modeling antennas as cantilevers, we can obtain the vibration modes in analytic forms. There are 6 first order modes. To the first order effect, QI and Q3 do not influence the attitude motion, Q2 and Q4 couple to Z-axis control system as shown in Fig.6, and Q5 and Q6 couple into transverse axes control system. Table 2 lists the hybrid modal data.
-7"]..~
.. 7l13li
c..,"') ..
rT1
i
t, e.. z.,-1iIA
..........
...
1
......... 1 I
+________________________________ ____
L~~~I
Fig. 4.
.
.
~
:
~
!
ro
IZ I
Gravity-gradient torque.
Based on a simplified model, Fig. 5 shows eST 4 _
IZ-w.,>
Fig. 6.
Vibration modes of EXOS-C antennas.
2452
K. Ninomiya
TABLE 2 modal no. 1st 2 order ~ 5 6 2nd 8 order 10 11 12 3rd 14 order 16 l 'I
18
Hybrid Modal Data angular freq. i rad.bee. ) 0 .162
1.')3
8 .11
eOURl iI1ii eoeff.
cx
ey
cz
0 0 35.56 0 0 - 35 .56 0 0 95.56 0 0 ! 9').')6 0 1 0 15.24 0 1-15.24 I ::J , o 0\15.24 0 0 15.24 0 , 0 5.44 i 0 0 -5.44 0 ! 5.44 0 i ! 0 0 5.44
Z-Axis Control Logic: In autonomous sunpointing mode of operation, the polarity of electric current in either MAC or MULD coils is swithed automatically to generate a control torque that will tend to reduce the deviation of Z-axis from the solar vector until it reaches in a predetermined small region. Fig. 7 is the block diagram for Z-axis control logic.
I
I
The nutation can be detected by using TFSS if one notices that X-component of TFSS output is expressed in the form; a = S-Ch/I) w, where S is the component of solar vector projected on to X-Z plane in the absence on nutation, h is the angular momentum of SWA, I is the transverse moment of inertia, and w is the X-component of angular velocity. In practice the d.c.-cut X-output of TFSS is fed into the logic circuit to generate the damping torque. The nutation damping rate is given by T/C nh) in terms of the magnetic torque T produced by MAC coil. For EXOS-C, this rate is estimated to be around 2°/hour. Transverse Axes Control Loop: As shown in Fig. 8, the system has three modes of operation; normal operation mode, rate-damping mode, and constant-wheel-speed mode. The rate -damping mode will have to be used in cases of gyro-digitizer or integrating-counter failures. The constant-wheel-speed mode will be employed when wheel is run up and long antennas .q re being extended or retracted. In the design of the system, the first order modes of flexibility are taken into account. In this case, the system has a gain margin greater than 20 dB, and phase margin of 40 degrees.
DYNAMICS Fig. 7.
Block diagram of magnetic torquers. Fig. 8.
In order to choose the adequate coil and to determine the current polarity, one needs to know the geomagnetic field relative to the spacecraft and also to determine the sign of the quantity; D = -SxBy + SyBx, where, Sx and Sy are respectively the X and Y components of the solar vector, and Bx and By are the X and Y components of the geomagnetic field. As a practical way of determining the sign of D without using an onboard computer, the X-Y plane is subdivided into 9 segments CA to I) as shown in Fig. 7. Segment I signifies the dead band. Corresponding to the combination of two segments in which the solar vector and the geomagnetic vector are projected respectively, the hard wired logic determines and issues the adequate control commands.
Block diagram of transverse axes control loop.
SIMULATION AND TEST To verify the design, computer simulations of the control system performance have been scheduled. In the simulation of Z-axis control, the effects of external disturbance torques and antenna flexibility will be carefully checked, as well as the performance in nutaion control. In transverse axes simulation, the gyro noise, nonlinear elements in the control electronics, and flexibility mode-couplings of higher modes will be paid attention to. Finally, coupled motion of Z-axis and transverse axes through the flexibility will be taken into account by using a simplified model for antennas.
Development of Attitude Stabilization As tests in developing the transverse axes control system, we started from the preliminary test to measure the parameters of SWA itself and tacho feed-back loop. Next step is to close the control loop by using the actual hardware except that the spacecraft dynamics is simulated in a minicomputer. As a confirmation of overall performance, the total system will be mounted on a spherical airbearing table and be operation-tested without simulating the moment of inertia nor flexibility. As for Z-axis control system, no dynamical test is conducted except for checks of the control logics and the polar ities and magnitudes of magnetic torquers. Acknowledgement. The author would like to express his gratitude to all Professors and technical staffs who have been actively taking part in the scientific satellite project at ISAS for their interest and support. Also, he wishes to acknowledge the cooperation of Messrs. T. Yamamoto, J. Aoyama, and K. Maeda of NEC in developing the attitude control system of EXOS-C. Discussion to Paper CS 7 .1 G. E. Cunningham (USA ) : Are the attitude control electronics impl e mented in hardwired logic or in software? K . Ninomiya (Japan): Th ey are implemented in h a rdwir ed logic. Mi c rop roc esso r s wi ll not be used for attitude control until 1986.
2453
P. Kant (Netherlands): In your servo test of EXOS-C ACS, did you measure the control torque of the scanning wheel to feed it into the simulated dynamiCS, or did you model the wheel dynamics? K. Ninomiya (Japan): We made the measurement on the wheel dynamics . In the simula tion which we call " mixed test" we modelled the wheel dynamiCS , according to the measure ment . J.W. Hursh (USA) : Is the EXO satellite a NASDA project? K. Ninomiya (Japan) : No. It is an ISAS project. ISAS manages and designs EXOS-C and other scientific satellites. F. Wojtalik (USA): Do your attitude control designs include a " fail - safe" automatic sun pOinting control feature? K. Ninomiya (Japan) : We have no provision except for an open loop mode of the switching of magnetic moment according to the program which will be commanded from the ground and stored in the program timer memory.