Precision Control, Knowledge and Orbit

Precision Control, Knowledge and Orbit
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SSC98-IX-3
Precision Control, Knowledge and Orbit Determination on a Small Spacecraft Bus
The OrbView-4 Attitude Control System
Dewey Adams
Orbital Science Corporation· 20301 Century Blvd, Germantown MD 20874 • 301-428-6324·
[email protected]
Dominick Bruno
Orbital Science Corporation • 20301 Century Blvd, Germantown MD 20874 • 301-428-6045·
[email protected]
Piyush Shah
Orbital Science Corporation· 20301 Century Blvd, Germantown MD 20874· 301-428-5216·
[email protected]
Dr. Brian S. Keller
Orbital Science Corporation· 20301 Century Blvd, Germantown MD 20874 • 301-428-6315 •
[email protected]
Abstract. High resolution, Earth imaging missions place stringent requirements on the spacecraft Attitude Control
System (ACS) in terms of pointing accuracy, attitude knowledge, and orbit knowledge. For OrbView-4, a body
scanner, the spacecraft must also be agile to maximize the number of targets that can be imaged each orbit. These
requirements dictate a robust, high performance design that allows rapid target acquisition and high line scan rates
for mapping while maintaining precise knowledge and control of the camera line of sight vector. The OrbView-4
ACS combines state of the art components with robust algorithm design to meet these objectives. Equipment
selection, attitude determination, attitude control, and GPS based onboard orbit determination are discussed as the
elements of an ACS design for an advanced Earth imaging spacecraft.
An overview of the OrbView-4 mission is presented along with a discussion of the requirements imposed on the
ACS. The system architecture is described in terms of its hardware and software elements. The design of the
attitude determination, orbit determination and attitude control algorithms are discussed, and the results of a system
performance analysis are summarized.
originates with customers who have a need for imagery.
In response to customer requests the OrbView
Operations Center (Ooq, after assessing customer
needs and priorities as well as spacecraft constraints
such as attitude agility, imaging rate, and downlink
capacity, uplinks imaging tasking commands to the
spacecraft. Subsequently, the collected wide band
imagery data is either downlinked directly to distributor
owned data downlink and processing centers, or stored
on board in solid state memory and down linked to one
of two OrbView ground terminals for dissemination to
the customer. The OOC maintains high latitude and
CONUS ground stations as primary downlink sites.
OrbView-4 Mission Overview
Orbview-4 is a high resolution commercial Earth
imaging satellite providing I meter panchromatic and 4
meter multispectral imaging capability. The OrbView-4
satellite also has a hyperspectral sensor with 8 meter
Ground Sample Distance (GSD) added to and
integrated with the camera to provide target spectral
information.
The OrbView-4 mission is to generate, process, and
distribute panchromatic, 4-band multispectral, and
hyperspectral imagery on a commercial basis. The
OrbView-4 payload will acquire 1 meter panchromatic
imagery between 0.45-0.9 I.l.m, 4 bands of 4 meter GSD
multispectral imagery in the 0.45 to 0.9 I.l.m spectral
band, and 280 bands of hyperspectral imagery in the
0.45 to 5.0 IJ.ffi spectral band.
Imagery tasking
D. Adams, D. Bruno, P. Shah, B. Keller
The spacecraft bus is responsible for providing all onorbit support required for the camera to obtain imagery
as commanded by the ~OC. This includes data
handling, data transmission, tasking implementation,
thermal control, attitude control, structural stability,
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12th AIAAlUSU Conference on Small Satellites
SSC98-IX-3
GPS Antenna
High Gain Antenna
System
Camera
Boresight, Zs
GPS Antenna
Fixed Solar Panel
(Cells facing -Zs)
Figure 1. OrbView-4 Imaging Spacecraft
power, clocks, and attitude knowledge required for
metric accuracy. The spacecraft gathers health and
engineering housekeeping data and downlinks it to the
OOc.
The spacecraft receives compressed and
uncompressed imagery data from the payload, formats
and stores it on board as required, and downlinks an
image data stream to the Earth ground terminals for
transmission to the data processing segment. The data
stream also contains data required to support post-pass
radiometric calibration, image motion compensation,
and geolocation. The spacecraft's mission lifetime is 5
years. The basic spacecraft design approach is singlestring, with selective redundancy applied where
necessary to achieve reliability needs. An outline of the
OrbView-4 spacecraft is shown in Figure I.
angle in any direction from nadir. The nominal orbit
period is about 94 minutes with eclipses ranging in
duration from 34 to 36 minutes.
OrbView-4 will be launched by a Taurus™ launch
vehicle into a sun-synchronous orbit with a 470 km
altitude, and 97.25 degree inclination. The descending
nodal crossing time will be nominally controlled to
10:30 AM, local solar time. The size of the altitude
error envelope is set to maintain a 3-day maximum
revisit time to any point on the globe within a field of
regard defined by a 50 degrees maximum elevation
Spacecraft operations vary as a function of the location
within the orbit. The typical spacecraft orbital routine
is displayed in Figure 2. On exiting eclipse the
spacecraft is oriented to align its fixed solar array
panels normal to the sun-line. This provides the
optimum attitude for recharging the batteries that were
slightly depleted during the eclipse.
During this
charging period the spacecraft is rotated about its yaw
D. Adams, D. Bruno, P. Shah, B. Keller
The spacecraft is required to support one or more
imaging windows each orbit. An imaging window
typically includes a mix of slow camera motions to scan
a target and collect its image, and fast target-to-target
slewing. This mix is highly variable. A single imaging
window may consist of a single, long scan with no
slews, or of several small-target scans with many slews
between. In practice, each imaging window contains a
variety of imaging types.
Orb View-4 Operational Sequence of Events
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SSC98-IX-3
Ground Station Pass
• Yaw slew for antenna coverage
No Imaging
Zone
Suo-point Slew
• 1 minute before eclipse exit
Pre-Image SeQuence
• Bias RWA speeds
• Slew SIC to first target
Imaging Window (10 minutes)
• Slew between targets
• Scan for image collection
Post-Image Sequence
• Return to sun-pointing
Nadir-point Slew
• At eclipse entry
No Imaging
Zone
Figure 2. ACS Events Sequence
orbit maintenance maneuvers. This orientation also
facilitates
payload
calibration
requirements.
Approximately one minute before exiting eclipse, the
attitude is reoriented to align the solar arrays with the
sun-line from thence the process repeats.
axis to provide a favorable geometry for pointing the
spacecraft high gain antenna toward a ground receiving
station. Once image tasking commands are uploaded,
the spacecraft, remains in its inertially fixed sunbathing
orientation until a few minutes before the scheduled
imaging window opens.
ACS Overview
Prior to imaging initiation the attitude control system
spins-up the reaction wheels (to avoid zero-speed
crossing during imaging) and slews the spacecraft to an
orientation that aligns the camera boresight to the first
target and yaws the spacecraft to orient the camera's
imaging array with the direction of scan. When the
imaging window opens the spacecraft attitude is
controlled in accordance with previously uploaded
open-loop parameters that control the scanning motion
during image gathering and slew commands to quickly
acquire subsequent targets.
The ACS is formulated with high performance
components to provide precise attitude and orbit
determination while achieving fine pointing control as
well as high speed slew capability to achieve rapid
target acquisition. A functional block diagram of the
ACS is shown in Figure 3.
Two star trackers are mounted, in a thermally isolated
manner, on the spacecraft optical bench perpendicular
to one another to allow precise attitude determination
about each of three orthogonal reference axes. The star
trackers provide an inertial attitude reference accurate
to within 6 arcsec (lo) while scanning at rates up to 1.5
deglsec. The star trackers are capable of maintaining
track at rates up to 6.0 deg/sec thereby facilitating rapid
target acquisition slew maneuvers.
Upon imaging completion, the spacecraft is slewed back
to the sunbathing attitude to recharge the batteries and
the reaction wheel speeds are reduced to conserve
power. On entering eclipse, the spacecraft attitude is
reoriented so that the camera boresight is nadir
pointing. This provides a favorable geometry for data
downlink and maintains a minimum drag orientation to
maximize mission life by minimizing the magnitude of
D. Adams, D. Bruno, P. Shah, B. Keller
The star trackers are complemented with the Orbital
Inertial Reference Unit (IRU) that utilizes fiber optic
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12th AIAAlUSU Conference on Small Satellites
SSC98-IX-3
gyro technology to provide highly accurate body motion
measurements. The Fiber Optic Gyros (FOG) have
been specially designed for the OrbView-4 mission.
Each FOG is comprised of two identical fiber optic
windings stacked along the axis of rotation and is
mounted to the optical bench using a single-point
attachment. This allows thermal expansion to occur
without distorting the optical bench while allowing the
FOG to sense high frequency optical bench motion and
thus preclude the need for angular displacement
sensors. The FOGs are remotely located from the IRU
electronics unit thereby removing all heat generating
elements from the optical bench. The Orbital IRU is
ideally suited for this application as it is free of dither
disturbances typically associated with ring laser or
mechanical gyros. The FOG IRU scale factor is less
that 0.04 arcseclbit, angular random walk less than
0.005 deglrt-hr, bandwidth greater than 150 Hz and its
data output frequency is 400 Hz. The IRU data is
stored onboard for subsequent downlink to facilitate
GPS Signals
image post-processing.
Four reaction wheels are arranged in a pyramid about
the spacecraft yaw axis. Each wheel is 65 degrees from
the yaw axis thereby providing high roll and pitch
control torque. Each reaction wheel is capable of
producing a torque of 0.3 Nm and storing 19.5 Nms of
angular momentum. The reaction wheels are divided
into three components to allow the reaction wheels to fit
inside a small spacecraft with tight volumetric
constraints and to dissipate the heat generated during
the imaging period. The reaction wheel assembly
contains the motor, rotor and bearing assemblies. The
electronics unit contains the motor drive and electrical
interface to the spacecraft. The ballast assembly
contains diodes used to dissipate, as heat, back-emf
generated power during wheel deceleration.
To
minimize jitter, each wheel rotor is balanced to better
than 0.5 gm-cm (static) and 5.0 gm-cm2 (dynamic) and
mounted to the spacecraft with vibration isolators.
Central Electronics Unit
(1)
GPS Antenna
(2 pairs)
Primary Attitude
Control Software
Time, Position, Velocity
GPS Receiver
(2)
,
Gimbal Commands --II>
Raw GPS Data
GOODS
A 'II
.-----lBC
nCI
D
---'100..
ary ata ----..-
To Antenna
Gimbals
To Solid State
Recorder
~
OJ
1------1 ..
10
o
~
Star Tracker
(2)
AT
Reaction Wheel
Ballast (4)
Attitude Power Electronics
(1 )
'"
1------"-----1 RT
Wheel Speed
-
Torque Comman d-+
Inertial Reference
Three Axis
Unit
r---Sody Rates
Reaction Wheel
Iy (4)
Safe Hold Attitude
Control Software
(1)
Coarse Sun
Sensor
Reaction Wheel
Electronics (4)
Modulated
28 Volt Power
f---Detector Currents
28 Volt Return
Three Axis
r--Field Strength
Thruster On/Off
Magnetic Torquer
(3)
(10)
Three Axis
Magnetometer
(1)
28 Volt Return
Thruster
(4)
Figure 3. ACS Block Diagram
D. Adams, D. Bruno, P. Shah, B. Keller
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SSC98-IX-3
GPS Onboard Orbit Determination Software (GOODS)
provides a spacecraft position reference in the 12000
coordinate system in each of the along track, cross track
and altitude directions accurate to within 25 meters
Ocr). The GPS antennas have been mounted on the
zenith side of the spacecraft and oriented to maximize
the contact duration with each acquired GPS space
vehicle. The GPS antennas nominally look toward the
orbit normals so that tracking duration is limited
primarily by GPS space vehicle motion and spacecraft
slews. The mean contact time of greater than 25
minutes for each tracked GPS space vehicle allows
pixel geodetic knowledge to be determined within 12
meters (Icr) during image post processing. Two GPS
receivers and two dual antennas are provided for
redundancy.
flow of payload and ancillary data to the solid state
recorder.
ACS Control Modes
The ACS utilizes six separate and distinct control
modes. The Launch, Normal, Imaging, Orbit Adjust
and Contingency control modes are executed by the
CEU processor and the Safe Hold control mode is
executed by the APE processor. At all times, either the
APE or the CEU is the attitude control computer. In
each processor, there is always one ACS control mode
active. When the APE is the control computer, the
attitude control commands from the Safe Hold mode are
supplied to the ACS actuators, while the APE inhibits
CEU attitude control commands from being supplied to
the ACS actuators. When the CEU is the control
computer, the APE inhibits the control commands from
the Safe Hold mode and delivers commands received
from the active CEU control mode to the ACS
actuators. The ACS control mode and submode can be
selected by command or autonomously following
detection of a system fault or an attitude or rate error in
excess of predefined thresholds. The attitude control
mode switching diagram is shown in Figure 4.
Ten coarse sun sensor detectors strategically positioned
about the spacecraft provide the failsafe means of
attitude determination in the Safe Hold mode. The
detectors provide near 41t steradian coverage. In the
event of an anomaly, this detector arrangement allows
optimal spacecraft slewing, in the presence of
shadowing and Earth albedo, to orient the solar array
panels toward the sun. The detector placement also
protects the camera from sun damage.
Safe Hold mode orients the spacecraft Zs-axis parallel
to the sun vector with the negative Z-axis oriented
toward the sun such that the spacecraft solar array
panels are illuminated. Once the sun orientation is
achieved a commanded rotation rate about the
spacecraft Z-axis is maintained. During eclipse the
nominal sun orientation is maintained.
A three-axis magnetometer and three magnetic torquer
bars are used for momentum control.
The
magnetometer sense range is -600 to +600 milliGauss
and the magnetic torquer dipole in nominally 12 ampm2• Each torquer bar is mounted on one of the three
control axes. A simple M cross B control law is
employed to maintain the momentum level to less than
0.5 Nms per control axis throughout the orbit. This
allows the spacecraft to be slewed with minimal
gyroscopic coupling induced by maintaining high levels
of stored momentum in the reaction wheels.
Contingency mode provides a means of operating the
spacecraft with the CEU computer in the event that both
star trackers are off-line or the spacecraft has an invalid
ephemeris due to a GOODS fault. Contingency mode is
also used following launch vehicle separation to provide
a power positive attitude while the star trackers and
GOODS are each initialized.
The ACS sensors and actuators are linked by the
Attitude and Power Electronics (APE). The APE acts
as a Remote Terminal (RT) on a dual redundant MilStd-1553B bus and contains a 69Rooo processor used
for Safe Hold mode processing as well as power and
thermal management. The APE is used to decode and
distribute command, control and data request signals
between the Central Electronics Unit (CEU) and
interface electronics for attitude sensors and actuators
and general purpose telemetry conditioning. The CEU
contains a RAD6000 processor which is used to
perform all the primary attitude determination and
control as well as the GOODS processing. The CEU
acts as the 1553 Bus Controller (BC) and also contains
uplink and downlink control functions and regulates the
D. Adams, D. Bruno, P. Shah, B. Keller
Normal mode provides a means of operating the
spacecraft during non-imaging and non-thrusting
operation. Three submodes are used to point the
spacecraft to the Earth, sun or an arbitrary inertial
attitude, while the fourth submode is used to provide a
means of slewing the spacecraft attitude from one
orientation to another. Typically the Earth Target
submode is used during eclipse periods; the Sun Target
mode is used either before or after imaging; and the
Inertial Target submode is used both to position the
spacecraft near its initial imaging attitude and prior to
orbit maintenance maneuvers.
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12th AIAAlUSU Conference on Small Satellites
SSC98-IX-3
CEU Resident Attitude Control Modes
Launch Vehicle Separa tion---+I
I
Launch
AWA Speed Control
AWA Torque Control
Primary
Even Backup
Odd Backup
Pulse
Sun Point
Sun Acquisition
Boot·up or Re.et-
Command
Star Tracker or
GOODS Fault
Command
1
,------Command---+I
Orbit
Adjust
Contingency
Normal
I
Sun Target
Earth Target
Inertial Target
Slew
14----Comma nd---...,
-AMCF Termination-
Maneuver Abort--:-
Imaging
_Star Tracker, IAU. AWA or_
GOODS Fault
L--_ _ _ _ _ _ _ _ _---J
Maneuver Termination-
Command
APE Resident Attitude Control Modes
Safe
Hold
Sun Point
Sun Acquisition
Figure 4. ACS Control Mode Switching Diagram
the reaction wheels with thruster augmentation, based
upon IRU sensed body rates. Four 4.45 N hydrazine
thrusters are off-pulsed for attitude control.
Imaging mode orients the camera boresight toward
ground based targets defined by an uploaded attitude
maneuver command file. The star tracker is used in
conjunction with the GPS derived orbit to provide the
attitude reference and the IRU is used to provide body
rate estimates. Reaction wheels are used to reorient to
the target attitude and to control the motion of the
spacecraft as commanded by the attitude maneuver
command file. A reaction wheel bias speed and initial
attitude are commanded before Imaging mode is
entered. The attitude is slewed to the initial imaging
attitude using the Normal mode, Slew submode prior to
the start of imaging. Ancillary data (i.e. attitude
determination and control data needed for post facto
image enhancement) processing is started before
beginning imaging and terminated once imaging is
complete.
The Launch mode is used to provide a known. benign
and controlled state of the ACS during spacecraft test as
well as during pre-launch and launch mission phases.
In Launch mode the ACS sensors are decoupled from
the ACS actuators. The ACS sensor and actuator data
outputs are provided in telemetry. Launch mode has
two submodes allowing control of either RWA speed or
torque.
Attitude Determination System
The OrbView-4 attitude determination system uses data
from the two star trackers and the IRU to estimate the
inertial-to-body attitude quaternion and body rates for
input to the attitude control system. Data from both star
trackers is used in normal operation, but the system is
also configured for single-tracker operation in the event
of a data loss from either of them. The attitude
determination logic is implemented as a conventional
The Orbit Adjust mode provides closed-loop attitude
control during thruster firing.
The ACS having
previously been oriented to the desired attitude, fires the
selected thrusters for the commanded number of jetseconds. The attitude is controlled in this mode using
D. Adams, D. Bruno, P. Shah, B. Keller
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SSC98-IX-3
Kalman filter updated with star tracker data at I Hz.
Between filter updates the attitude quaternion is
propagated with gyro data at 10Hz.
by propagating the measured attitude quaternions
forward through the delay time using the estimated rates
from the IRU processing logic.
The corrected
measurements are output to the Kalman filter logic.
The four logic components are identified on Figure 5, a
functional diagram of the attitude determination system,
along with their software execution rates. In addition to
using data from the star trackers and IRU, the logic also
uses data from the GOODS to correct the tracker data
for velocity aberration.
A six-state Kalman filter is used to estimate the
incremental corrections to the current attitude and gyro
bias estimates. Filter updates occur at I Hz. The error
covariance is propagated with the estimated rates from
the IRU processing logic. Positive definiteness and
divergence checks are performed on the covariance
matrix prior to the residual computation and covariance
state updates. For each star tracker, the corrected
measured quaternion is used together with the current
propagated attitude quaternion to form the residuals in
the star tracker frame. In normal operation, when data
from both star trackers are available, the boresight axis
residual for each tracker is set to zero so that only the
more accurate horizontal-vertical axis data from both
trackers is used to update the state estimate. However,
the boresight axis residual is used if data from only a
single tracker is available. Spurious data is rejected by
comparing the residuals to a pre-selected threshold.
The covariance and state are then updated sequentially
with the data from the two trackers. The attitude
quaternion is updated with the first three elements of the
estimated state and the gyro bias estimate is updated
with the second three elements. The resulting inertialto-body attitude quaternion is output to the attitude
control system and the gyro biases to the IRU
processing logic.
The IRU data is sampled and processed at 10Hz. For
each axis, the accumulated counts are converted to
incremental angles that are corrected for the estimated
gyro bias from the Kalman filter. These incremental
angles are used to propagate the inertial-to-body
attitude quaternion, also at 10 Hz. The IRU processing
logic also contains a discrete differentiator to compute
the body rates from the bias-corrected incremental
angles. These estimated rates are provided to the
Kalman filter for covariance propagation, and also to
the attitude control logic for rate control.
The star tracker processing logic reads the measured
attitude quaternions at 1 Hz and corrects them for
velocity aberration and transport delay. The aberration
correction is computed based on the instantaneous star
tracker line-of-sight vector in the inertial frame, and the
spacecraft velocity vector and the sun position vector
from the GOODS. The data transport delay is corrected
Attitude
Quaternions
(tracker 'rames)
Star
, Trackers
Velocity Aberration and
Corrected Attitude
Quaternions
(tracker frames)
Time Delay Corrections
Sun
I-----'Vector---~
(1 Hz)
GOODS
Spacecralt Velocily---tooi
Vector
~---.r------.....J
t--
To Conlrollogic
~
Gyro
~Data
IRU Data
Processing
(10 Hz)
I-----'_ _ Eslimated SIC Angular
Rates
Incremental
Angles
Propagate
Ouaternion
(10 Hz)
E stimated
Attitude ~
Kalman Filter (1 Hz)
• Covariance propagation
• Covariance validation
- Residual calculation
- Residual validation
- State vector update
• Ouaternion update
• Gyro bias update
.....
To
~
-to- Control
logic
Q uaternion
'--_ _ _ _ _ Estimated Inertial-to-Body Frame Attitude
Quaternion
'--_ _ _ _ _ _ _ _ _ _ _ _E,slimated Increme ntal Gyro
Biases
Figure 5. Attitude Determination System Functional Diagram
D. Adams, D. Bruno, P. Shah, B. Keller
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12th AIANUSU Conference on Small Satellites
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Figure 6 displays the result of a simulation of the
attitude determination system performance in Imaging
mode. A ten-minute imaging pass was simulated by
computing the required three-axis rate profiles for five
different scan directions. Each scan is preceded by a 45
second acceleration profile to achieve the desired initial
rates and this is included in the two minutes allotted per
scan. The top frame shows the estimated gyro rates
from the IRU processing logic and shows the rate
profiles for the five scans. The first and last 300
seconds of the simulation occur in Earth pointing mode,
with the imaging pass occurring in the middle 600
seconds.
The middle frame shows the attitude
estimation errors during the imaging pass. As shown,
these remain within ±8 arcseconds even during the
acceleration preceding each scan. This performance is
achieved through the capability of the star trackers to
maintain accuracy over the entire range of scan rates
required for the mission, and by weighting their
measurements more heavily in the computation of the
Kalman gains. The bottom frame shows the gyro bias
estimation errors. For each axis, a bias with constant
and orbit rate-varying components was modeled in the
simulation. As shown, the bias estimation errors remain
within ±O.3 arcsecondlsecond during the imaging pass.
Orbit Determination
To meet mission pointing requirements, Orb View 4
utilizes a GPS receiver plus estimation software known
as GOODS to achieve better than 25 meter (10) realtime position knowledge. Two GPS receivers are
provided. One is active and the other is cold redundant.
The GPS receiver supports full 12 channel tracking and
on-orbit reprogramming.
Each receiver is
independently connected to its own pair of antennas,
pointed generally in opposite directions to maximize the
field of view. A simplified block diagram of the
GOODS is shown in Figure 7.
GOODS is based on NASA/Goddard's GPS Enhanced
Orbit Determination Experiment (GEODE) software.
The original GEODE software processes raw GPS
pseudorange and Doppler measurements through a
Kalman filter to estimate position, velocity, an
atmospheric drag coefficient, receiver bias, and receiver
bias drift.
Editing is performed on the GPS
measurements to reject data with high ionospheric
Gyro Rates
2.-----~-------.------~----_,,_----~------,
oG)
~OJ O~--------------~==~~
G)
"0
Attitude Estimation Errors
10.-----~.-----~._----~._----_.------~------_.
en
o
G)
en
ectI
o
-10L-----~L-
____
~L__ _ _ _~L__ _ _ _~L__ _ _ _~L__ _ _ _~
Gyro Bias Estimation Errors
oG)
~
en
15en
0.2
o
~ -0.2
o
200
400
600
800
1000
1200
Time (seconds)
Figure 6. Attitude Determination Performance in Imaging Mode
D. Adams, D. Bruno, P. Shah, B. Keller
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SSC98-IX-3
GPS Signals
Ls""vecto~
GOODS
14---E pheme ri 9-----11>1
14----Almanac:---1I>I
GPS ANTENNA
AND RECEIVER
i
Measurement i
Update
.
(30 sec)
r---Moon Vectof---tlio>
-SIC Position----.
Position Messag
._1
lr---s/C Velocity---tl>
rRange Messag
State Update
'--_ _ _ _ _----liinitialization Messag
(10Hz)
I----Time
~
Figure 7. GOODS Block Diagram
corruption.
perfonnance level.
If GOODS diverges, it is
reinitialized using the receiver state output. If the
receiver fails to acquire or loses lock at any time,
GOODS continues propagating until the receiver
acquires. If receiver re-initialization is required, the
receiver initialization procedure is followed again.
GOODS is allowed to propagate without GPS inputs for
up to _
after which Fault Detection and
Correction logic commands the spacecraft into Safe
Hold mode that does not require orbit position
knowledge.
GEODE models spacecraft accelerations with a 30x30
gravity model; Earth, sun, and moon point-mass
accelerations; a Harris-Priester based atmospheric
density model; and a solar radiation pressure model.
GEODE performs GPS raw measurement processing in
a 30 second loop. Propagation updates occur every 10
seconds. Simplified real-time propagation for control
usage nominally occurs at a 1 second rate. GEODE can
be used in a propagate-only mode if GPS measurements
GEODE was licensed from
are unavailable.
NASNGoddard and modified by Orbital to create the
GOODS.
Attitude Control
The OrbView-4 attitude control system uses the attitude
determination system for spacecraft attitude and rate
estimates, and four reaction wheels mounted in a
pyramid configuration for attitude control. The nominal
mission uses all four wheels to maximize agility along
the roll and pitch axes. However, the system can be
configured for three wheel operation. A 10Hz control
frequency is used together with a Proportional-IntegralDerivative (PID) controller to meet the stringent
pointing requirements during imaging. A control loop
bandwidth of 0.37 Hz provides sufficient robustness in
presence of parameter variations, structural modes, and
fuel slosh. The key functions of the attitude control
system, as displayed graphically in Figure 8, are
command generation, error signal generation, a simple
PID controller, and reaction wheel torque distribution.
The GOODS is integrated into the Orb View-4 flight
software which runs on a RAD6000 processor.
Simplified real-time propagation was increased to
confonn with the 10Hz control rate. GOODS utilizes
Separate
pseudorange and range rate as inputs.
interface software converts the receiver time
measurements and integrated carrier phase data into
GOODS internal
pseudorange and range rate.
quantities are also used for other calculations within the
attitude control software, including a sun vector and the
transfonnation between the WGS-84 and 12000 EarthCentered Inertial coordinate systems.
Orbit detennination requires initialization of GOODS
and the receiver via a boot-strap method. GOODS is
initialized by uploading a state vector and entering
propagate mode. An initialization vector for the
receiver is then created from a GOODS-propagated
state. Once the receiver acquires and outputs valid data,
GOODS is allowed to use the measurements in state
GOODS then converges to its full
updates.
D. Adams, D. Bruno, P. Shah, B. Keller
The command generation algorithm generates
spacecraft attitude, rate, and torque feedforward
commands to point the spacecraft in the desired
direction. The attitude commands are generated in the
fonn of a quaternion which represents the desired
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12th AlAAlUSU Conference on Small Satellites
SSC98-IX-3
ATTITUDE
MANEUVER
COMMAND FILE
--1--
Uplink Imaging
Commands
10 Hz
I
r··..·I~('rifig~~:~······! I ~~r~~ea~~ed.Forward
1
L. ... ~2.1l1.Il1.!iDA~.... j
i
Quaternion
Command
NORMAL MODE
.... j
L....~2.1l1.1l1.!i.~.~.s..... j
i·····EiiUfTiiige·f····j
l......C;S!.'!l.Il1.Cl.n.~.~..... .i
j....iiiii·rHiirfar·ijeCj
L.....c;S!.'!l.Il1.ClD.~.s..... l
j""sTew'Co'iTima'ilifj
i······~ruii·Tar·gef
i
Rate
Command
Reaction Wheel
Bias Speed
Attitude
Error
Rate
ERROR SIGNAL I Error
GENERATION
Integral
10 Hz
IError
PID
CONTROL
LAW
Reaction
~ Torques
10 Hz
10 Hz
T
~
10 Hz
Estimated
Spacecraft
Attitude
Reaction
Wheel
Momentum
GYROSCOPIC
TORQUE
FEEDFORWARD
I
T
(From GOODS)
REACTION
WHEEL
TORQUE
DISTRIBUTION
Estimated Reaction
Wheel Friction Torque
L.....~.~I'l.ll.ri'lt.i.9.I1 ...1
Sun Position
SiC Position
SIC Velocity
1
I
10 Hz
REACTION
WHEEL
TACHOMETER
PROCESSING
10 Hz
Spin
-Direction
_:achometer
Pulses
'---------'
Estimated
Spacecraft
Rate
(From Attitude Determination)
Figure 8. Attitude Control System Functional Diagram
orientation of the spacecraft with respect to an inertial
frame.
The PID controller generates three axis control torques
used to orient the spacecraft in the commanded attitude.
The control torque is generated by applying
proportional, derivative, and integral control gains to
the attitude, rate, and integral errors, respectively.
Then, a gyroscopic torque feedforward term is added to
enhance pointing.
During imaging, these commands are generated from a
ground uplinked Attitude Maneuver Command File
(AMCF). The AMCF contains data to point the camera
optical boresight at the specified ground target. This
includes location of the ground target in WGS-84 Earth
Centered Earth Fixed coordinates, yaw rotation about
the camera's optical boresight, and torque feedforward
commands in spacecraft coordinates.
The reaction wheel torque distribution algorithm
converts the required control torque along spacecraft
axes to be applied to each reaction wheel. With four
wheels in use, an extra degree of freedom in distributing
a three-axis torque in spacecraft coordinates among the
four wheels exists. To provide the additional constraint,
a pseudo-inverse distribution law is used. This law
minimizes a quadratic form of the control torques (i.e.,
the sum squared of all the wheel torques). With this
approach the wheel torques tend to equalize within the
constraint of producing the requisite torque while
minimizing wheel power consumption. The general
form of the pseudo-inverse law is used to distribute
spacecraft torque among the four wheels, thus allowing
easy management of the reaction wheel bias speed.
During Normal mode operations spacecraft commands
are generated using spacecraft and sun ephemeris from
GOODS. These commands point the spacecraft at local
nadir (Earth target submode), at sun (sun target
submode), or in an inertially fixed orientation (inertial
target submode). Slews between these targets are
performed autonomously using either a ramp-ramp or a
ramp-coast-ramp rate profile.
The error signal generation algorithm computes the
attitude, rate, and integral errors by comparing the
commanded state with the estimated state. During
imaging and slews, the integral error is set to zero to
improve spacecraft pointing performance.
D. Adams, D. Bruno, P. Shah, B. Keller
The results of a simulation performed to assess attitude
control system performance during imaging are shown
in Figure 10 and Figure 9. In this scenario the
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SSC98-IX-3
AITITUDE ERROR
30~---~---~------------~----------r------------r-----------r~~---~
· Image
Turn ~ .
Around:
20
.I
I
.,
..
..
..
..
..
. . . . . 1..
. .......
10 ....
,."
•
•
..
..
..
•
...
..
..
.....
..
..
..........
","..
...
..
..
...........
\,
..
..
..
..
..
..
..
•
..........................................................
.,
•
..
*
•
•
•
J
-10
-20~------~--------~--------~--------~--------~------~
o
20
40
80
60
120
100
RATE ERROR
40
o
~~~
30
u
20
...
..............
. Image
~~
J
I
Turn
~.
. . . . . . . . .: . . . . . . . . . : . . . . . . . . . :. . . . . . . . . : ...A!~u~?
......
"
. . . . . . . ." . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
".
i
.. ' . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .l
Q)
.!!!
u
Q)
10
!II
!::
CIl
0
........
-10
-20
0
::
.................................................................
20
40
60
80
~
..........
100
.,.
120
tirre (sec)
Figure 9. Control Errors During Imaging
be uplinked. Utilization of a modular development
approach for attitude determination, orbit determination
and attitude control functions allows the system to be
easily reconfigured for a wide variety of missions.
spacecraft is slewed from an initial parking attitude to
an imaging attitude. An imaging scan is then performed
starting at an initial rate of 1.5 deglsec. At the end of
the scan the spacecraft is reversed to scan in the
opposite direction.
Author Biographies
Summary
Dewey Adams is a Senior Principal Engineer in the
Attitude Determination and Control group at Orbital
Sciences Corporation. As the OrbView-4 lead ACS
engineer he is responsible for assuring that the ACS is
designed to meet performance requirements, developed
within cost and schedule constraints and tested to
guarantee successful performance once in flight. Prior
to joining Orbital, Mr. Adams was employed at
The OrbView-4 ACS combines precise attitude and
orbit determination with high performance control
actuators to meet the stringent attitude control
requirements imposed by a high resolution, body
scanning, Earth imaging spacecraft. The ACS is
completely autonomous, requiring only imaging tasking
commands generated from customer imaging requests to
D. Adams, D. Bruno, P. Shah, B. Keller
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12th AIANUSU Conference on Small Satellites
SSC98-IX-3
SPACECRAFT PITCH ATTITUDE AND RATE COMMAND
20.--------,-.----------~----------~--------~----------_.----------,2
§
<=>
- -20
~
~ -40
-60L-------~~----------~----------L---------~----------~----------~-2
o
20
40
60
80
100
120
Time (Seconds)
SPACECRAFT PITCH ACCELERATION COMMAND
0.4r--=~~~-r--------_.r_--------_r--------_.r_--------_r--------~
0.2
20
40
60
80
100
120
Time (Seconds)
Figure 10. Commanded Spacecraft Motion During Imaging
Lockheed Martin Astro Space where he was the lead
ACS engineer on a number of low Earth orbiting
imaging
and
geosynchronous
communications
spacecraft.
Mr. Adams holds a B.S. degree in
Aeronautical and Astronautical Engineering from the
Pennsylvania State University, and a M.S. degree in
Mechanical and Aerospace Engineering from Rutgers
University.
Piyush Shah is a Principal Engineer in the Attitude
Determination and Control group at Orbital Sciences
Corporation. He is responsible for the design of the
attitude control system, and also participated in the
initial overall subsystem definition. Prior to joining
Orbital, Mr. Shah was an attitude control subsystem
engineer at Hughes Space and Communications. He has
also worked at TRW. Mr. Shah holds a B.S. degree in
Electrical Engineering from Ohio State University, and
a M.S. degree in Electrical Engineering from University
of Michigan.
Dominick Bruno is a Senior Staff Engineer in the
Attitude Determination and Control group at Orbital
Sciences Corporation. He is responsible for the design
of the attitude detennination system, and also
participated in the initial overall subsystem definition.
Prior to joining Orbital, Mr. Bruno was an attitude
control systems engineer at Lockheed Martin Astro
Space.
He has also worked at the Aerospace
Corporation, the Charles Stark Draper Laboratory, and
Grumman Aerospace. Mr. Bruno holds a S.B. and
Engineer degrees in Aeronautics and Astronautics from
M.I.T. and a M.S. degree in Aeronautics and
Astronautics from Polytechnic University. He is a
Senior Member of the AlAA.
D. Adams, D. Bruno, P. Shah, B. Keller
Dr. Brian S. Keller received his B.S.E. in Aerospace
Engineering from the University of Michigan (1986)
and his M.S. (1987) and Ph.D. (1993) in Aeronautics
and Astronautics from Stanford University. He has held
engineering positions with JPL, Ford Aerospace, and
Lockheed-Martin working on various satellites
including Galileo, Landsat 7, DMSP, and TIROS. Dr.
Keller started at Orbital in 1997 and is currently the
ACS lead for the OrbView-3 spacecraft.
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