- Email: [email protected]
GADACS: A GPS attitude determination and control experiment on a spartan spacecraft
ControlEng.Practice,Vol.3,No.8,pp. 1125-1130,1995
1995ElsevierScienceLtd PrintedinGreatBritain 0967-0661/95$9.50+0.00
Pergamon 096%0661(95)00106-9
GADACS: A GPS ATTITUDE DETERMINATION AND CONTROL EXPERIMENT ON A SPARTAN SPACECRAFT F.H. Bauer*, E.G. Ligl~ey*, J. ~
*
,
J. O'Donmfll* and R. Sdmurr**
*NASA GoddardSpace FffglitCenter, Code 712~Greaibett MD 20771, USA **NASA GoddardSpace Flight Ontev, Code 745, Greenl~lt,.MD20771, USA
(Received October 1994; in final form March 1995)
Abstract. An attitude experimentis currently being developedto fly on a Spartan Space Shuttle deployedspacecraft. The experiment,called the GPS Attitude DeterminationAnd Control System (GADACS), is expected to answer key questions to help pave the way for GPS attitude sensing in the future. An overview of the GADACS experiment,its missiongoals, hardware description, and control law design,is presented. Key Words. Attitude control; Global PositioningSystem; control system design; satellite control; aerospace engineering
1. INTRODUCTION
tern (GADACS), is expected to answer key questions to help pave the way for GPS attitude sensing for future spacecraft. GPS attitude sensing is accomplished through interferometric frequency discrimination of the GPS carrier wave. The approach used measures minute changes in the GPS codeless beacon wavelength between two antennas separated by a fixed baseline, as shown in Fig. 1. Since the L1 1575.42 MHz GPS beacon frequency
The NAVSTAR Global Positioning System (GPS) was developed by the United States Department of Defense to provide precise navigation and time information for ground and air-based vehicles. Recent breakthroughs in application-specific integrated circuits and the development of advanced digital signal processing chips, have produced 3-pound, 4.5-watt receivers with the capability of measuring differential position to centimeter precision. This receiver has been modified by Stanford University (Cohen, 1992; Cohen el al., 1993; Lightsey et al., 1993) to enable the measurement of real-time space vehicle attitude. This new vehicle sensor provides precise time, three-axis position updates, and three-axis attitude sensing data.
planar w
to GPS s
Significant breakthroughs in the field of spacecraft attitude determination and control are expected in the next few years by using GPS technology. Attitude performance on the order of 0.5-0.1 o is expected from this technology through the use of simple, low-cost patch antennas, GPS receivers, and specialized interface electronics. This new technology should revolutionize the field of spacecraft attitude determination and control with simple, low-cost, low-weight sensors that can provide better pointing performance than many highercost, state-of-the-art Earth Sensors.
a
Antenna 1
a
t
v
e
f
~
r
o
e
l
z
- Antenna 2
~
~
Fig. 1. Attitude Sensing with Differential GPS
is used, small gained patch antennas can be installed on the spacecraft with minimal impact on the spacecraft's structural design. For three-axis attitude determination and control, a minimum of three antennas are required. For redundancy, four antennas are employed; to provide 4~r steradian field of view, a OPS receiver and set of antennas is used for each side of the spacecraft.
An attitude experiment is currently being developed to fly on a Spartan Space Shuttle deployed spacecraft. The experiment, called the GPS Attitude Determination And Control Sys1125
1126
F.H. Bauer et al. 2. GADACS MISSION OVERVIEW
The GADACS experiment will fly on the STS72 Shuttle flight as part of the Spartan "OASTFlyer" spacecraft. Spartan, in its OAST-Flyer configuration, is shown in Fig. 2. This mis-
of the GPS constellation and the changeover from one GPS receiver/antenna set to another.
3. SPACE F L I G H T E X P E R I M E N T A T I O N RATIONALE Several GPS attitude-determination experiments have been performed which have demonstrated attitude determination to 0.10 or less. These experiments include ground-based measurements made from laboratory rooftops, aircraft-based experiments, and most recently an on-orbit flight experiment (Lightsey et al., 1994). While these experiments demonstrate that the attitude-sensing concept for spacecraft appears feasible, they do not totally prove that GPS attitude sensing will be robust enough for spacecraft control operations. There are several key questions that only an onorbit free-flying experiment can answer. These include:
Fig. 2. Spartan OAST-Flyer Spacecraft sion is currently manifested to fly in November, 1995. The Spartan spacecraft is a small, rectangular, free-flying vehicle, measuring roughly 1 x 1.25 x 1.5 meters. It is released from the Shuttle and picked up after several days of conducting its experiments. In addition to the GADACS experiment, the OAST-Flyer mission will be carrying three others. These are the Return Flux Experiment (REFLEX), the Solar Exposure to Laser Ordnance Devices Experiment (SELODE), and the Shuttle Packet Radio Experiment (SPRE). The R E F L E X and SELODE experiments will be operating for approximately the first two-thirds of the OAST-Flyer's two-day mission. During this time, the Spartan attitude control system will maneuver the spacecraft into a preset series of attitudes while these experiments are conducted. The GADACS experiment will be operational the entire time the OAST-Flyer is free flying. During the R E F L E X and SELODE portion of the mission, GADACS will be collecting attitudedetermination data using its two GPS receivers. This data will be recorded on tape, along with gyro measurements that will be used to verify and calibrate the GPS attitude measurements. Throughout the mission, star tracker updates will be performed every other orbit in order to calibrate the onboard gyros. Approximately the last third of the mission is the GADACS portion. For four orbits of this time, approximately six hours, GADACS will assume control of the vehicle, using only GPS-sensed attitude for closed-loop attitude control. During this time, GADACS will control the spacecraft in a series of inertial profiles in order to test visibility
• To what degree does the high orbital velocities of space flight corrupt the differential measurement of the GPS beacon signal? • Will doppler shifts impact GPS attitude sensing? * Do on-orbit thermal drifts, coupled with the above effects, impact GPS attitude sensing? • Can the GADACS stay locked on to the GPS beacons and provide accurate attitude data while the spacecraft performs major slew maneuvers? Is sensing degraded during these maneuvers? • What are the performance limitations of GPS attitude sensing? • When compared to ground testing, is multipath the same or worse on-orbit? Can this effect be effectively calibrated out, through ground testing, prior to launch? These key questions can only be answered through on-orbit tests of free-flying spacecraft. One onorbit GPS attitude experiment has flown and several others are proposed to fly over the next few years. (See Table 1.) These experiments have been planned to help the GPS attitude control community "stair step" its way through the various questions and concerns that this newtechnology attitude sensor raises. The GADACS experiments will provide two important results to the aerospace community. First, the GPS sensor data will be compared to onboard high-precision gyroscopes and a star tracker. This will provide the first ever measurement of the absolute accuracy of GPS-based attitude determination in space. The second result of the experiment is to control the spacecraft using the GPS attitude sensor exclusively. This flight will be the first time that GPS will be used in this role on a spacecraft. Thus, the proposed GADACS experiment
GADACS Control Experiment Table 1
On-Orbit GPS Attitude Experiments
Spacecraft RADCAL Crista-SPAS REX-II Spartan OAST Flyer AMSAT Phase III-D SSTI-Lewis SSTI-Clark
Date of Flight June, 1993 November, 1994 September, 1995 November, 1995 April, 1996 July, 1996 July, 1996
will provide a significant boost in the development and testing of an on-orbit GPS attitude sensor.
4. GADACS HARDWARE DESCRIPTION
1127
It provides six axes of information via three twoaxis dry tuned rotor mechanical gyros. This gyro has flown on two previous fine-pointing missions; the Solar Maximum Mission in 1980 and on the STS-35 Space Shuttle Mission in 1990. The GADACS electronics and sensor hardware are mounted on a special plate which is housed inside the Spartan spacecraft, as shown in Fig. 4.
DRIRU-I~ 2vT~NS" ' l . M f ~ * % . J g l " f
ltJ~/fl
I
The flight hardware required for the SpartanGADACS experiment is shown in the block diagram of Fig. 3. The GADACS hardware includes
",t~/
I DRIRU-III
I GPS IPatck.~tm~s (4 e,~+k)
Fig. 4. GADACS Hardware on the OAST-Flyer
+':'+"
GADA(~
Pt,',m.~ (4 ,..+ak}
II
_______1 Imterfo~ Comtlel l ]~.-4++I Emm~,k. [ ~
Plate
GAOACS
,ronlR
18V ~
Fig. 3. GADACS Hardware Block Diagram two sets of four L-band patch antennas and preamplifiers, two GPS receivers, capable of sensing position, velocity, time, and attitude, a GADACS Interface Control Electronics (G-ICE) and a Dry Rotor Inertial Reference Unit (DRIRU-II) gyroscope. The patch antennas, pre-amplifiers and GPS receivers are commercial off-the-shelf hardware developed by Trimble Navigation, Ltd. The receivers to be flown on GADACS are the Trimble TANS Vector. This receiver has been augmented with specialised orbital software developed by Stanford University and the NASA Goddard Space Flight Center. The G-ICE is a Goddard-developed electronics box which provides an interface to the various sensors, develops the attitude control signals and distributes these to the Spartan thrusters. The G-ICE includes DRIRU-II and TANS Vector interface circuitry, a Motorola 68020 microprocessor, command and telemetry processing circuitry, GADACS power supplies and system timers. The DRIRU-II is a high-precision gyroscope developed by Teledyne.
In addition to this hardware, the eight patch antennas are located on the exterior of the spacecraft. Each antenna is mounted on a 12-inch ground plane and bracket which is attached to the exterior of the Spartan structure. Four antennas are located on each of two opposite faces of the spacecraft to provide 4~r steradian field of view coverage. Figs. 2 and 5 represent two views of
GPS Anl 4thm 4 op[
SELODI
Fig. 5. Alternate View of GADACS Antennas the Spartan spacecraft. As shown, the antennas are mounted as far as is practically possible from one another without interfering with the R E F L E X payload, making the longest baseline between antennas approximately 1.2 meters. Each of the two sets of four antennas provide hemispherical cover-
1128
F.H. Bauer et al.
age to the GPS satellite constellation. At least one antenna on each side of the spacecraft is mounted out of plane with the other three. This configuration better facilitates the acquisition of full threeaxis attitude. The antenna pre-amplifiers are coupled to the antennas with a minimal cable length. These pre-amps are mounted on the inside face of the Spartan spacecraft to minimize temperature gradients.
5. C O N T R O L L E R DESCRIPTION The GADACS experiment is unique because it will be the first spacecraft mission to quantify the accuracy of GPS attitude solutions. The expected in-flight accuracy of this device for a onemeter separation between GPS antennas is hoped to be approximately 0.15°; these solutions will be provided by the receiver at a 1 Hz sample rate. During four orbits of the GADACS portion of the mission, the GPS solutions will be used as sensor inputs to the attitude control system. Because the GPS attitude solutions are relatively noisy, and the exact noise characteristics that will be seen on-orbit are not well known, the controller requirements will be to achieve reasonable pointing performance without excessive actuation in the presence of noise. Furthermore, in the event of loss of GPS attitude, significant disagreement between the GPS attitude solutions and the inertially-derived measurements, or excessive spacecraft rates, control will be given back to the Spartan control system until the condition is corrected. A conceptual block diagram of the GADACS controller is shown in Fig. 6. The continuous plant ' I l~ml I~lle-Axls k~J
Euler Anglee
',
the same natural frequency as the controller, and estimate both the vehicle angular error and rate in roll, pitch, and yaw. The controller consists of three single-loop PD controllers with a bandwidth of approximately 0.01 rad/sec. The control signal drives a nonlinear gas jet actuator, a pulse-frequency modulated system with dead zone and saturation. For the purposes of simulation, the actuator was modeled as a piecewise-linear nonlinearity. A study was performed to demonstrate that this approximation is reasonably accurate when compared to the actual pulse-frequency modulation. The actuator signal is also provided as an estimator input. A three-axis high-fidelity simulation was developed to demonstrate the performance of the control system, which included gravity gradient, aerodynamic, and magnetic enviromental torques. The GPS attitude solution noise was modeled as a Gaussian random process with statistically uncorrelated samples of 0.30 RMS standard deviation. In actuality, the noise is spectrally colored by multipath. The main determinants of performance that were used in evaluating the design were controller bandwidth, steady-state error, dead-zone size, and limit-cyclefrequency. Of these, the latter proved to be the main requirement on the design; the limit-cycle frequency had to be very low, for the design to remain within its very limited fuel budget. Furthermore, the limit-cycle frequency needed to be conservatively chosen, given the uncertain noise characteristics of the sensor. During the GADACS control portion of the mission, the controller will hold the spacecraft fixed inertially. The dead zone size was designed to achieve the proper limit-cycle frequency and eliminate false actuations due to sensor noise. Because of the conservative noise characteristics used, the necessary dead zone size for this system was found to be 2 °. The system response to a large (20 °) initial error in roll axis is shown in Fig.7.The
,dl~
.
.
.
.
.
.
.
.
.
.
.
.
.
.
.
.
.
.
.
.
.
.
.
.
.
.
.
.
.
.
.
.
.
.
.
.
.
.
.
.
{dq)
.
Fig. 6. GADACS Controller Block Diagram is sampled by the GPS attitude receiver at an update rate of 1 Hz. Failure Detection and Correction (FDC) logic (described below) verifies the GPS attitude solution as compared with the inertially-derived measurement of the vehicle attitude. The verified GPS attitude quaternion is then differenced with the command quaternion and the resulting error quaternion is converted into Euler roll, pitch, and yaw angles, which are fed into three single-loop estimators. These estimators are designed to have settling qualities of
100
'
~
.
.
.
.
30G
.
.
Fig. 7. GADACS Controller Error Response
.
.
4Q O
GADACS Control Experiment roll plot shows the true roll error ( r o l l t r u e ) , the measured roll error from the GPS receiver ( r o l l e r r o r ) , and the roll-error output of the estimator ( r o l l e r r o r e a t ) . As shown, the control system captures the roll error to within the 20 dead zone within approximately 250 seconds. The saturation of the control system causes a slightly slower and more lightly damped response than in the unsaturated case, but the system performance is nonetheless satisfactory.
1129
with respect to fuel consumption and limit-cycle frequency in the presence of sensor noise. Provisions have been made to temporarily return control to the Spartan ACS in the event that the GPS measurements are unavailable or unacceptable for spacecraft attitude control.
6. FAILURE D E T E C T I O N AND CORRECTION
Fig. 8 shows the design with no input, showing
!
A Failure Detection and Correction (FDC) capability is a key component of the GADACS experiment to ensure mission safety. As the name implies, FDC ensures safe operation of the Spartan spacecraft through the detection of potential failures of different aspects of the GADACS and implementation of appropriate corrective measures. Because of the highly aggressive schedule for the system development of the GADACS experiment, the design objective was to keep the FDC system as simple as possible while making it robust enough to ensure, as much as possible, a successful experiment.
. . . . . .
.................. ~............... i............... i~i..............~ ........... i ..... jl' - :/ ~ ...... i ...... f~\ ............. / , /::\
',l^ II it~/~ .......... ~'1 ~/!\ ,tl
............ J_.............. \...:t..',A ....... d..._.l ,' ',/ k~ !/ ill
•
,
/
:
,,,:,
............ .... ' .............
f. i p. ' t~ ....... b----f-~ ...... ~!', ....
F.~._ i /
~i,,
...... :._..L.I_...,.,,. .....
V',',',,i l/!n :
'" [i~'2 ! ...... ' /1,''"~f ~]"............ Irk.
t,,
J- .....
-~
:
t : .....
•
Fig. 8. GADACS Controller Limit Cycling the limit cycles induced in roll, pitch, and yaw by the enviromental torques. This is a relatively benign limit cycle for the expected noise input profile. Fig. 9 shows the cumulative output of the jet
o:
;
rm ~
;
;
L :::::::::::::::i:::::::::::::::!:::::::::::::::i:::::::::::::::!:: ..........
I0!I!!i! ...............i...............!............................!i!...............ii..... Fig. 9. GADACS Controller Actuator Performance actuators for this case in terms of the number of jet pulse firings ( J p u l s e s ) and the total change in velocity induced on the spacecraft (Jacc). The amount of fuel used in this case is well within the fuel budget for the GADACS control portion of the mission; additional runs were performed to demonstrate that the fuel usage remained tolerable for up to twice the expected sensor noise. In summary, the performance of the GPS attitude based controller has been shown to be satisfactory
The FDC system, as it relates to the attitude control system components of the GADACS experiment, consists of four main sections, which are described briefly below. The four main sections of the FDC concern: 1. correct operation of the GPS receivers, 2. correct operation of the DRIRU-II gyro unit being used to verify operation of the GPS receivers, 3. attitude determination from both the GPS and DRIRU-II units and verification of valid attitude solutions, and 4. attitude control of the Spartan spacecraft. The GPS receivers being used on the GADACS experiment contain a great deal of sophisticated flight software of their own. Because of this, the only FDC for the receivers is to detect a lock-up condition. If no output has been received from a receiver for a given amount of time, that receiver is reset and reinitialized. As discussed above, the DRIRU-II gyro unit consists of three gyro channels, each of which output two axes of attitude information, giving two redundant axes in each of three orthogonM directions. This section of the FDC is concerned with using the raw gyro outputs, digital pulses which represent incremental attitude changes, in order to detect for failures of a gyro channel (two-axis failure) or single axis. If a failure is detected, the gyro attitude quaternions are reconfigured to use good axes. If both axes of a redundant pair have failed, the GADACS control portion of the mission is ended.
F.H. Bauer et al.
l 130
The attitude determination section of the FDC takes the attitude solutions output by the GPS receivers and verifies them by comparing them to the two gyro quaternions (after doing the transformations necessary to put all solutions in the same reference frame). A GPS attitude solution that is within 5 o in each axis of either the prime or redundant gyro quaternion is considered to be verified. If both GPS receivers have output verified attitude solutions for the current cycle, the FDC calculates a Figure Of Merit (FOM) for each, using such information as the number of satellites each antenna of each GPS receiver has locked on to, as well as the corresponding signal-to-noise ratios. There is also a hysteresis used to prevent a "ping-ponging" situation between the two GPS receivers. Whichever attitude solution has the highest FOM is used. The final part of the FDC system deals with attitude control and is the most important, as it must ensure that system failures do not cause GADACS to endanger the Spartan spacecraft, or to put it into a state that will make it difficult for the Shuttle to retrieve. When a verified attitude solution exists, this solution is fed into the estimator, and the estimator output angular errors and rates are used in the PD control law. The resulting jet actuator command will be passed back to the Spartan. If there is no verified attitude solution present, for consecutive periods of up to 10 seconds, the estimator will be propagated open-loop, and the jet actuator commands will be zeroed. If this condition persist for longer than 10 seconds, GADACS will return control of the spacecraft back to the Spartan ACS system, and will not resume control until the GPS receivers begin to output verifiable attitude solutions. The FDC also checks to see if the specified rate limit has been exceeded for five consecutive cycles. If it has, the FDC returns control to the Spartan ACS, which will then kill the excessive rates and return the vehicle to the desired position. After this has been done, GADACS will re-assume control.
7. CONCLUSIONS The GADACS experiment represents the first calibration of GPS attitude accuracy and noise jitter in space by onboard sensors of greater measurement precision than the GPS sensor. GPS sensor sensitivity to slew maneuvers, orbital mo-
tion and vehicle multipath will be quantified during this mission. The experiment is also expected to demonstrate the ability to use attitude data derived from GPS as the sole means of controlling vehicle attitude. A GPS-based attitude control law was developed and is expected to provide pointing performance of the order of 2 ° .
8. A C K N O W L E D G E M E N T S The authors wish to gratefully acknowledge the multi-organizational team who have made GADACS GPS attitude experimention possible. In particular, they wish to acknowledge the support from NASA Headquarters Code C for their financial backing as part of the Mission Technology Infusion program and for the Spartan flight. The government and contractor support personnel team at NASA Goddard, particularly members of the Guidance and Control Branch (Code 712), the Special Payload Division (Code 740), especially Angie Russo, Howard Shore, and Peter Lansing, the software engineers, led by Dave Leucht, and the team at the Johnson Space Center, particularly Moises Montez and Penny Saunders, and the support and guidance provided by Clark Cohen and Brad Parkinson at Stanford University, Kurt Brock at Space Systems/ Loral and Marc Crotty at 'IYimble Navigation Ltd. have been crucial to the success of this program.
9. R E F E R E N C E S Cohen, C. E. (1992). Attitude determination using GPS. PhD thesis. Stanford University. Stanford, California USA. Cohen, C. E., E. G. Lightsey, W. A. Feess and B. W. Parkinson (1993). Space flight tests of attitude determination using GPS: preliminary results. In: Proceedings of ION GPS-93. Salt Lake City, Utah USA. Lightsey, E. G., C. E. Cohen and B. W. Parkinson (1993). Application of GPS attitude determination to gravity gradient stabilized spacecraft. In: Proceedings of lhe 1993 AIAA Guidance, Navigation, and Control Conference. Monterey, California USA. Lightsey, E. G., C. E. Cohen and B. W. Parkinson (1994). Attitude determination and control for spacecraft using differential GPS. In: Pro-
ceedings of ESA International Conference on Spacecraft Guidance, Navigation, and Control. Noordwijk, The Netherlands.
1995ElsevierScienceLtd PrintedinGreatBritain 0967-0661/95$9.50+0.00
Pergamon 096%0661(95)00106-9
GADACS: A GPS ATTITUDE DETERMINATION AND CONTROL EXPERIMENT ON A SPARTAN SPACECRAFT F.H. Bauer*, E.G. Ligl~ey*, J. ~
*
,
J. O'Donmfll* and R. Sdmurr**
*NASA GoddardSpace FffglitCenter, Code 712~Greaibett MD 20771, USA **NASA GoddardSpace Flight Ontev, Code 745, Greenl~lt,.MD20771, USA
(Received October 1994; in final form March 1995)
Abstract. An attitude experimentis currently being developedto fly on a Spartan Space Shuttle deployedspacecraft. The experiment,called the GPS Attitude DeterminationAnd Control System (GADACS), is expected to answer key questions to help pave the way for GPS attitude sensing in the future. An overview of the GADACS experiment,its missiongoals, hardware description, and control law design,is presented. Key Words. Attitude control; Global PositioningSystem; control system design; satellite control; aerospace engineering
1. INTRODUCTION
tern (GADACS), is expected to answer key questions to help pave the way for GPS attitude sensing for future spacecraft. GPS attitude sensing is accomplished through interferometric frequency discrimination of the GPS carrier wave. The approach used measures minute changes in the GPS codeless beacon wavelength between two antennas separated by a fixed baseline, as shown in Fig. 1. Since the L1 1575.42 MHz GPS beacon frequency
The NAVSTAR Global Positioning System (GPS) was developed by the United States Department of Defense to provide precise navigation and time information for ground and air-based vehicles. Recent breakthroughs in application-specific integrated circuits and the development of advanced digital signal processing chips, have produced 3-pound, 4.5-watt receivers with the capability of measuring differential position to centimeter precision. This receiver has been modified by Stanford University (Cohen, 1992; Cohen el al., 1993; Lightsey et al., 1993) to enable the measurement of real-time space vehicle attitude. This new vehicle sensor provides precise time, three-axis position updates, and three-axis attitude sensing data.
planar w
to GPS s
Significant breakthroughs in the field of spacecraft attitude determination and control are expected in the next few years by using GPS technology. Attitude performance on the order of 0.5-0.1 o is expected from this technology through the use of simple, low-cost patch antennas, GPS receivers, and specialized interface electronics. This new technology should revolutionize the field of spacecraft attitude determination and control with simple, low-cost, low-weight sensors that can provide better pointing performance than many highercost, state-of-the-art Earth Sensors.
a
Antenna 1
a
t
v
e
f
~
r
o
e
l
z
- Antenna 2
~
~
Fig. 1. Attitude Sensing with Differential GPS
is used, small gained patch antennas can be installed on the spacecraft with minimal impact on the spacecraft's structural design. For three-axis attitude determination and control, a minimum of three antennas are required. For redundancy, four antennas are employed; to provide 4~r steradian field of view, a OPS receiver and set of antennas is used for each side of the spacecraft.
An attitude experiment is currently being developed to fly on a Spartan Space Shuttle deployed spacecraft. The experiment, called the GPS Attitude Determination And Control Sys1125
1126
F.H. Bauer et al. 2. GADACS MISSION OVERVIEW
The GADACS experiment will fly on the STS72 Shuttle flight as part of the Spartan "OASTFlyer" spacecraft. Spartan, in its OAST-Flyer configuration, is shown in Fig. 2. This mis-
of the GPS constellation and the changeover from one GPS receiver/antenna set to another.
3. SPACE F L I G H T E X P E R I M E N T A T I O N RATIONALE Several GPS attitude-determination experiments have been performed which have demonstrated attitude determination to 0.10 or less. These experiments include ground-based measurements made from laboratory rooftops, aircraft-based experiments, and most recently an on-orbit flight experiment (Lightsey et al., 1994). While these experiments demonstrate that the attitude-sensing concept for spacecraft appears feasible, they do not totally prove that GPS attitude sensing will be robust enough for spacecraft control operations. There are several key questions that only an onorbit free-flying experiment can answer. These include:
Fig. 2. Spartan OAST-Flyer Spacecraft sion is currently manifested to fly in November, 1995. The Spartan spacecraft is a small, rectangular, free-flying vehicle, measuring roughly 1 x 1.25 x 1.5 meters. It is released from the Shuttle and picked up after several days of conducting its experiments. In addition to the GADACS experiment, the OAST-Flyer mission will be carrying three others. These are the Return Flux Experiment (REFLEX), the Solar Exposure to Laser Ordnance Devices Experiment (SELODE), and the Shuttle Packet Radio Experiment (SPRE). The R E F L E X and SELODE experiments will be operating for approximately the first two-thirds of the OAST-Flyer's two-day mission. During this time, the Spartan attitude control system will maneuver the spacecraft into a preset series of attitudes while these experiments are conducted. The GADACS experiment will be operational the entire time the OAST-Flyer is free flying. During the R E F L E X and SELODE portion of the mission, GADACS will be collecting attitudedetermination data using its two GPS receivers. This data will be recorded on tape, along with gyro measurements that will be used to verify and calibrate the GPS attitude measurements. Throughout the mission, star tracker updates will be performed every other orbit in order to calibrate the onboard gyros. Approximately the last third of the mission is the GADACS portion. For four orbits of this time, approximately six hours, GADACS will assume control of the vehicle, using only GPS-sensed attitude for closed-loop attitude control. During this time, GADACS will control the spacecraft in a series of inertial profiles in order to test visibility
• To what degree does the high orbital velocities of space flight corrupt the differential measurement of the GPS beacon signal? • Will doppler shifts impact GPS attitude sensing? * Do on-orbit thermal drifts, coupled with the above effects, impact GPS attitude sensing? • Can the GADACS stay locked on to the GPS beacons and provide accurate attitude data while the spacecraft performs major slew maneuvers? Is sensing degraded during these maneuvers? • What are the performance limitations of GPS attitude sensing? • When compared to ground testing, is multipath the same or worse on-orbit? Can this effect be effectively calibrated out, through ground testing, prior to launch? These key questions can only be answered through on-orbit tests of free-flying spacecraft. One onorbit GPS attitude experiment has flown and several others are proposed to fly over the next few years. (See Table 1.) These experiments have been planned to help the GPS attitude control community "stair step" its way through the various questions and concerns that this newtechnology attitude sensor raises. The GADACS experiments will provide two important results to the aerospace community. First, the GPS sensor data will be compared to onboard high-precision gyroscopes and a star tracker. This will provide the first ever measurement of the absolute accuracy of GPS-based attitude determination in space. The second result of the experiment is to control the spacecraft using the GPS attitude sensor exclusively. This flight will be the first time that GPS will be used in this role on a spacecraft. Thus, the proposed GADACS experiment
GADACS Control Experiment Table 1
On-Orbit GPS Attitude Experiments
Spacecraft RADCAL Crista-SPAS REX-II Spartan OAST Flyer AMSAT Phase III-D SSTI-Lewis SSTI-Clark
Date of Flight June, 1993 November, 1994 September, 1995 November, 1995 April, 1996 July, 1996 July, 1996
will provide a significant boost in the development and testing of an on-orbit GPS attitude sensor.
4. GADACS HARDWARE DESCRIPTION
1127
It provides six axes of information via three twoaxis dry tuned rotor mechanical gyros. This gyro has flown on two previous fine-pointing missions; the Solar Maximum Mission in 1980 and on the STS-35 Space Shuttle Mission in 1990. The GADACS electronics and sensor hardware are mounted on a special plate which is housed inside the Spartan spacecraft, as shown in Fig. 4.
DRIRU-I~ 2vT~NS" ' l . M f ~ * % . J g l " f
ltJ~/fl
I
The flight hardware required for the SpartanGADACS experiment is shown in the block diagram of Fig. 3. The GADACS hardware includes
",t~/
I DRIRU-III
I GPS IPatck.~tm~s (4 e,~+k)
Fig. 4. GADACS Hardware on the OAST-Flyer
+':'+"
GADA(~
Pt,',m.~ (4 ,..+ak}
II
_______1 Imterfo~ Comtlel l ]~.-4++I Emm~,k. [ ~
Plate
GAOACS
,ronlR
18V ~
Fig. 3. GADACS Hardware Block Diagram two sets of four L-band patch antennas and preamplifiers, two GPS receivers, capable of sensing position, velocity, time, and attitude, a GADACS Interface Control Electronics (G-ICE) and a Dry Rotor Inertial Reference Unit (DRIRU-II) gyroscope. The patch antennas, pre-amplifiers and GPS receivers are commercial off-the-shelf hardware developed by Trimble Navigation, Ltd. The receivers to be flown on GADACS are the Trimble TANS Vector. This receiver has been augmented with specialised orbital software developed by Stanford University and the NASA Goddard Space Flight Center. The G-ICE is a Goddard-developed electronics box which provides an interface to the various sensors, develops the attitude control signals and distributes these to the Spartan thrusters. The G-ICE includes DRIRU-II and TANS Vector interface circuitry, a Motorola 68020 microprocessor, command and telemetry processing circuitry, GADACS power supplies and system timers. The DRIRU-II is a high-precision gyroscope developed by Teledyne.
In addition to this hardware, the eight patch antennas are located on the exterior of the spacecraft. Each antenna is mounted on a 12-inch ground plane and bracket which is attached to the exterior of the Spartan structure. Four antennas are located on each of two opposite faces of the spacecraft to provide 4~r steradian field of view coverage. Figs. 2 and 5 represent two views of
GPS Anl 4thm 4 op[
SELODI
Fig. 5. Alternate View of GADACS Antennas the Spartan spacecraft. As shown, the antennas are mounted as far as is practically possible from one another without interfering with the R E F L E X payload, making the longest baseline between antennas approximately 1.2 meters. Each of the two sets of four antennas provide hemispherical cover-
1128
F.H. Bauer et al.
age to the GPS satellite constellation. At least one antenna on each side of the spacecraft is mounted out of plane with the other three. This configuration better facilitates the acquisition of full threeaxis attitude. The antenna pre-amplifiers are coupled to the antennas with a minimal cable length. These pre-amps are mounted on the inside face of the Spartan spacecraft to minimize temperature gradients.
5. C O N T R O L L E R DESCRIPTION The GADACS experiment is unique because it will be the first spacecraft mission to quantify the accuracy of GPS attitude solutions. The expected in-flight accuracy of this device for a onemeter separation between GPS antennas is hoped to be approximately 0.15°; these solutions will be provided by the receiver at a 1 Hz sample rate. During four orbits of the GADACS portion of the mission, the GPS solutions will be used as sensor inputs to the attitude control system. Because the GPS attitude solutions are relatively noisy, and the exact noise characteristics that will be seen on-orbit are not well known, the controller requirements will be to achieve reasonable pointing performance without excessive actuation in the presence of noise. Furthermore, in the event of loss of GPS attitude, significant disagreement between the GPS attitude solutions and the inertially-derived measurements, or excessive spacecraft rates, control will be given back to the Spartan control system until the condition is corrected. A conceptual block diagram of the GADACS controller is shown in Fig. 6. The continuous plant ' I l~ml I~lle-Axls k~J
Euler Anglee
',
the same natural frequency as the controller, and estimate both the vehicle angular error and rate in roll, pitch, and yaw. The controller consists of three single-loop PD controllers with a bandwidth of approximately 0.01 rad/sec. The control signal drives a nonlinear gas jet actuator, a pulse-frequency modulated system with dead zone and saturation. For the purposes of simulation, the actuator was modeled as a piecewise-linear nonlinearity. A study was performed to demonstrate that this approximation is reasonably accurate when compared to the actual pulse-frequency modulation. The actuator signal is also provided as an estimator input. A three-axis high-fidelity simulation was developed to demonstrate the performance of the control system, which included gravity gradient, aerodynamic, and magnetic enviromental torques. The GPS attitude solution noise was modeled as a Gaussian random process with statistically uncorrelated samples of 0.30 RMS standard deviation. In actuality, the noise is spectrally colored by multipath. The main determinants of performance that were used in evaluating the design were controller bandwidth, steady-state error, dead-zone size, and limit-cyclefrequency. Of these, the latter proved to be the main requirement on the design; the limit-cycle frequency had to be very low, for the design to remain within its very limited fuel budget. Furthermore, the limit-cycle frequency needed to be conservatively chosen, given the uncertain noise characteristics of the sensor. During the GADACS control portion of the mission, the controller will hold the spacecraft fixed inertially. The dead zone size was designed to achieve the proper limit-cycle frequency and eliminate false actuations due to sensor noise. Because of the conservative noise characteristics used, the necessary dead zone size for this system was found to be 2 °. The system response to a large (20 °) initial error in roll axis is shown in Fig.7.The
,dl~
.
.
.
.
.
.
.
.
.
.
.
.
.
.
.
.
.
.
.
.
.
.
.
.
.
.
.
.
.
.
.
.
.
.
.
.
.
.
.
.
{dq)
.
Fig. 6. GADACS Controller Block Diagram is sampled by the GPS attitude receiver at an update rate of 1 Hz. Failure Detection and Correction (FDC) logic (described below) verifies the GPS attitude solution as compared with the inertially-derived measurement of the vehicle attitude. The verified GPS attitude quaternion is then differenced with the command quaternion and the resulting error quaternion is converted into Euler roll, pitch, and yaw angles, which are fed into three single-loop estimators. These estimators are designed to have settling qualities of
100
'
~
.
.
.
.
30G
.
.
Fig. 7. GADACS Controller Error Response
.
.
4Q O
GADACS Control Experiment roll plot shows the true roll error ( r o l l t r u e ) , the measured roll error from the GPS receiver ( r o l l e r r o r ) , and the roll-error output of the estimator ( r o l l e r r o r e a t ) . As shown, the control system captures the roll error to within the 20 dead zone within approximately 250 seconds. The saturation of the control system causes a slightly slower and more lightly damped response than in the unsaturated case, but the system performance is nonetheless satisfactory.
1129
with respect to fuel consumption and limit-cycle frequency in the presence of sensor noise. Provisions have been made to temporarily return control to the Spartan ACS in the event that the GPS measurements are unavailable or unacceptable for spacecraft attitude control.
6. FAILURE D E T E C T I O N AND CORRECTION
Fig. 8 shows the design with no input, showing
!
A Failure Detection and Correction (FDC) capability is a key component of the GADACS experiment to ensure mission safety. As the name implies, FDC ensures safe operation of the Spartan spacecraft through the detection of potential failures of different aspects of the GADACS and implementation of appropriate corrective measures. Because of the highly aggressive schedule for the system development of the GADACS experiment, the design objective was to keep the FDC system as simple as possible while making it robust enough to ensure, as much as possible, a successful experiment.
. . . . . .
.................. ~............... i............... i~i..............~ ........... i ..... jl' - :/ ~ ...... i ...... f~\ ............. / , /::\
',l^ II it~/~ .......... ~'1 ~/!\ ,tl
............ J_.............. \...:t..',A ....... d..._.l ,' ',/ k~ !/ ill
•
,
/
:
,,,:,
............ .... ' .............
f. i p. ' t~ ....... b----f-~ ...... ~!', ....
F.~._ i /
~i,,
...... :._..L.I_...,.,,. .....
V',',',,i l/!n :
'" [i~'2 ! ...... ' /1,''"~f ~]"............ Irk.
t,,
J- .....
-~
:
t : .....
•
Fig. 8. GADACS Controller Limit Cycling the limit cycles induced in roll, pitch, and yaw by the enviromental torques. This is a relatively benign limit cycle for the expected noise input profile. Fig. 9 shows the cumulative output of the jet
o:
;
rm ~
;
;
L :::::::::::::::i:::::::::::::::!:::::::::::::::i:::::::::::::::!:: ..........
I0!I!!i! ...............i...............!............................!i!...............ii..... Fig. 9. GADACS Controller Actuator Performance actuators for this case in terms of the number of jet pulse firings ( J p u l s e s ) and the total change in velocity induced on the spacecraft (Jacc). The amount of fuel used in this case is well within the fuel budget for the GADACS control portion of the mission; additional runs were performed to demonstrate that the fuel usage remained tolerable for up to twice the expected sensor noise. In summary, the performance of the GPS attitude based controller has been shown to be satisfactory
The FDC system, as it relates to the attitude control system components of the GADACS experiment, consists of four main sections, which are described briefly below. The four main sections of the FDC concern: 1. correct operation of the GPS receivers, 2. correct operation of the DRIRU-II gyro unit being used to verify operation of the GPS receivers, 3. attitude determination from both the GPS and DRIRU-II units and verification of valid attitude solutions, and 4. attitude control of the Spartan spacecraft. The GPS receivers being used on the GADACS experiment contain a great deal of sophisticated flight software of their own. Because of this, the only FDC for the receivers is to detect a lock-up condition. If no output has been received from a receiver for a given amount of time, that receiver is reset and reinitialized. As discussed above, the DRIRU-II gyro unit consists of three gyro channels, each of which output two axes of attitude information, giving two redundant axes in each of three orthogonM directions. This section of the FDC is concerned with using the raw gyro outputs, digital pulses which represent incremental attitude changes, in order to detect for failures of a gyro channel (two-axis failure) or single axis. If a failure is detected, the gyro attitude quaternions are reconfigured to use good axes. If both axes of a redundant pair have failed, the GADACS control portion of the mission is ended.
F.H. Bauer et al.
l 130
The attitude determination section of the FDC takes the attitude solutions output by the GPS receivers and verifies them by comparing them to the two gyro quaternions (after doing the transformations necessary to put all solutions in the same reference frame). A GPS attitude solution that is within 5 o in each axis of either the prime or redundant gyro quaternion is considered to be verified. If both GPS receivers have output verified attitude solutions for the current cycle, the FDC calculates a Figure Of Merit (FOM) for each, using such information as the number of satellites each antenna of each GPS receiver has locked on to, as well as the corresponding signal-to-noise ratios. There is also a hysteresis used to prevent a "ping-ponging" situation between the two GPS receivers. Whichever attitude solution has the highest FOM is used. The final part of the FDC system deals with attitude control and is the most important, as it must ensure that system failures do not cause GADACS to endanger the Spartan spacecraft, or to put it into a state that will make it difficult for the Shuttle to retrieve. When a verified attitude solution exists, this solution is fed into the estimator, and the estimator output angular errors and rates are used in the PD control law. The resulting jet actuator command will be passed back to the Spartan. If there is no verified attitude solution present, for consecutive periods of up to 10 seconds, the estimator will be propagated open-loop, and the jet actuator commands will be zeroed. If this condition persist for longer than 10 seconds, GADACS will return control of the spacecraft back to the Spartan ACS system, and will not resume control until the GPS receivers begin to output verifiable attitude solutions. The FDC also checks to see if the specified rate limit has been exceeded for five consecutive cycles. If it has, the FDC returns control to the Spartan ACS, which will then kill the excessive rates and return the vehicle to the desired position. After this has been done, GADACS will re-assume control.
7. CONCLUSIONS The GADACS experiment represents the first calibration of GPS attitude accuracy and noise jitter in space by onboard sensors of greater measurement precision than the GPS sensor. GPS sensor sensitivity to slew maneuvers, orbital mo-
tion and vehicle multipath will be quantified during this mission. The experiment is also expected to demonstrate the ability to use attitude data derived from GPS as the sole means of controlling vehicle attitude. A GPS-based attitude control law was developed and is expected to provide pointing performance of the order of 2 ° .
8. A C K N O W L E D G E M E N T S The authors wish to gratefully acknowledge the multi-organizational team who have made GADACS GPS attitude experimention possible. In particular, they wish to acknowledge the support from NASA Headquarters Code C for their financial backing as part of the Mission Technology Infusion program and for the Spartan flight. The government and contractor support personnel team at NASA Goddard, particularly members of the Guidance and Control Branch (Code 712), the Special Payload Division (Code 740), especially Angie Russo, Howard Shore, and Peter Lansing, the software engineers, led by Dave Leucht, and the team at the Johnson Space Center, particularly Moises Montez and Penny Saunders, and the support and guidance provided by Clark Cohen and Brad Parkinson at Stanford University, Kurt Brock at Space Systems/ Loral and Marc Crotty at 'IYimble Navigation Ltd. have been crucial to the success of this program.
9. R E F E R E N C E S Cohen, C. E. (1992). Attitude determination using GPS. PhD thesis. Stanford University. Stanford, California USA. Cohen, C. E., E. G. Lightsey, W. A. Feess and B. W. Parkinson (1993). Space flight tests of attitude determination using GPS: preliminary results. In: Proceedings of ION GPS-93. Salt Lake City, Utah USA. Lightsey, E. G., C. E. Cohen and B. W. Parkinson (1993). Application of GPS attitude determination to gravity gradient stabilized spacecraft. In: Proceedings of lhe 1993 AIAA Guidance, Navigation, and Control Conference. Monterey, California USA. Lightsey, E. G., C. E. Cohen and B. W. Parkinson (1994). Attitude determination and control for spacecraft using differential GPS. In: Pro-
ceedings of ESA International Conference on Spacecraft Guidance, Navigation, and Control. Noordwijk, The Netherlands.