EARTH BASED SPACECRAFT ORBIT AND ATTITUDE CONTROL USING A LOOK-AHEAD THRUSTER SELECTION LOGIC AND MAGNETIC TORQUERS

CONTROLE AU SOL DE L'ORBITE ET DE L'ATTITUDE D'UN VAISSEAU SPATIAL AU MOYEN D'UN CIRCUIT LOGIQUE DE SELECTION DE PROPULSEURS A ANTICIPATION ET DE GENERATEURS DE COUPLES MAGNETIQUES

English Abstract

A spacecraft ground loop controller (GLC), located on the
Earth, interfaces with a satellite ground station receiving
spacecraft telemetry from the downlink baseband equipment
and automatically sending spacecraft commands through the
command uplink baseband equipment to control the attitude
of an orbiting spacecraft and achieve partial orbit control
using commanded thruster firings and magnetic torquer
polarity and magnitude. A cooperative approach of using
all available thrusters, of both the primary and redundant
strings, provides greater fuel savings.

French Abstract

L'invention est un contrôleur de boucle de mise à la terre pour véhicule spatial; ce contrôleur, qui est installé sur la Terre, sert d'interface avec une station terrienne qui reçoit des signaux de télémétrie provenant du véhicule spatial par le biais du matériel de bande de base et transmet automatiquement des commandes au véhicule spatial via le matériel de bande de base de la liaison ascendante de commande pour contrôler l'attitude d'un véhicule spatial en orbite et contrôler l'orbite partiellement au moyen de mises à feu contrôlées en utilisant des générateurs de couples magnétiques de signe et de grandeur contrôlés. Une approche coopérative utilisant tous les propulseurs disponibles, ceux de la chaîne primaire et ceux de la chaîne redondante, permet de réaliser des économies de carburant.

Claims

8

CLAIMS

What is claimed is:

1. An Earth-based spacecraft ground loop controller
(GLC) for controlling an Earth-orbiting satellite's on-
board attitude control equipment, including momentum
wheels, momentum wheel tachometers, gyros, magnetic
torquers, and thrusters, in combination with a satellite
ground station receiving spacecraft sensor data telemetry
signals from the downlink baseband equipment and sending
spacecraft commands through the command uplink baseband
equipment, to maintain and alter the satellite's
orientation and orbit, comprising:

means for interfacing with said satellite ground
station to receive spacecraft telemetry data from the
downlink baseband equipment and to send spacecraft
commands through the command uplink baseband
equipment; and

means, coupled to said interfacing means and
responsive to telemetry data from an orbiting
spacecraft indicating momentum wheel failure, for
automatically sending spacecraft commands through the
command uplink baseband equipment to control the
attitude of said orbiting spacecraft and achieve
partial orbit control using commanded thruster firings
and magnetic torquer polarity and magnitude.

2. A controller as in claim 1 wherein said means for
automatically sending spacecraft commands comprises:

a computer running satellite control software, and
wherein said software comprises:

means for providing estimators;





9

means for providing controller algorithms;

means for providing redundancy management
routines; and

means for providing communication routines.

3. A controller as in claim 2 wherein said estimators
comprise:

means for extracting sensor data from the telemetry
data; and

means for using said sensor data to provide output
estimates of the spacecraft attitude and rates and yaw
angle.

4. A controller as in claim 1 further comprising:

means, in said spacecraft responsive to said commands
when the momentum wheels are turned off, for resetting
a first wheel tachometer such that the wheel speed
measurement is zero;

means, in said spacecraft responsive to said commands,
for using a second wheel tachometer to read the
tachometer underspeed limit measurement; and

means, in said spacecraft, for using the difference
between the first and second tachometer measurements
to cause the magnetic torquers to provide current
output to control torque and for setting the polarity
and magnitude in response to said commands.

5. A controller as in claim 1 further comprising:

means, in said spacecraft responsive to said commands,




10

for providing roll thruster torques by manual thruster
pulse commands in 10 ms quanta of pulse widths.

6. A controller as in claim 1 further comprising:

means, in said spacecraft responsive to said commands,
for providing pitch/yaw thruster torques by
differentially setting a commanded speed for each
wheel, such that cycling through on-orbit wheel
control mode will immediately effect a momentum unload
pulse; and

means, in said spacecraft responsive to said commands,
for setting the momentum unload pulse.

7. A controller as in claim 6 further comprising:

means for setting the momentum unload pulse by user
command.

8. A controller as in claim 1 further comprising:

means for discarding the standard roll , pitch, and yaw
independent axis control; and

means for selecting thruster firing by propagating the
spacecraft attitude forward in time for each candidate
thruster to be fired, whereby fuel use is minimized or
pointing performance is improved.

9. A controller as in claim 1 further comprising:

means for setting the current of a magnetic torquer of
the spacecraft to select the commanded magnetic
torquer polarity and magnitude; and

means for determining the magnetic torquer command




11

selection as a function of the estimated attitude of
the spacecraft.

10. A method for controlling an Earth-orbiting
satellite's on-board attitude control equipment, including
momentum wheels, momentum wheel tachometers, gyros,
magnetic torquers, and thrusters, in combination with a
satellite ground station receiving spacecraft sensor data
telemetry signals from the downlink baseband equipment and
sending spacecraft commands through the command uplink
baseband equipment, to maintain and alter the satellite's
orientation and orbit, comprising the steps of:

interfacing with said satellite ground station to
receive spacecraft-momentum-indicative telemetry data
from the downlink baseband equipment and to send
spacecraft commands through the command uplink
baseband equipment; and

in response to receiving spacecraft-momentum-
indicative telemetry data from an orbiting spacecraft
indicating momentum wheel failure, automatically
sending spacecraft commands through the command uplink
baseband equipment to control the attitude of said
orbiting spacecraft and achieve partial orbit control
using commanded thruster firings and magnetic torquer
polarity and magnitude.

11. The method as in claim 10 wherein said step of
automatically sending spacecraft commands comprises:

providing a computer running satellite control
software, and providing said software with estimators,
controller algorithms, redundancy management routines,
and communication routines.

12. The method as in claim 11 wherein said estimators




12

operate to take sensor data, as extracted from the
telemetry data, and provide output estimates of the
spacecraft attitude and rates and yaw angle.

13. The method as in claim 12 wherein said
estimators' step of providing yaw angle output estimates
comprises:

taking data, as extracted from the telemetry data, and
applying a gyrocompassing filter thereto driven by
continuously running gyros.

14. The method as in claim 12 wherein said estimators
further operate to take data from the thruster torque table
and provide thruster torque estimates.

15. The method as in claim 14 wherein said step of
providing controller algorithms comprises:

taking said estimates of the spacecraft attitude and
rates and yaw angle and thruster torque and producing
thruster firing commands to maintain the spacecraft
attitude within a defined pointing region.

16. The method as in claim 15 wherein said step of
providing controller algorithms further comprises:

taking said estimates of the spacecraft attitude and
rates and yaw angle and thruster torque and producing
torquer commands to actuate and control said magnetic
torquers to maintain the spacecraft attitude within a
defined pointing region.

17. The method as in claim 10 further comprising the
step of:

selecting thruster firings by propagating the dynamic




13

model of the spacecraft attitude forward in time for
each candidate thruster that may be fired to determine
which thruster firing or combination of thruster
firings provides minimum fuel use for a predetermined
pointing performance.

18. The method as in claim 10 wherein said commanded
thruster firings comprise the step of:

selecting said thruster firings by propagating the
spacecraft attitude forward in time for each candidate
thruster to be fired, whereby fuel use is minimized or
pointing performance is improved.

19. The method as in claim 10 wherein said commanded
magnetic torquer polarity and magnitude are selected by
setting the current of a magnetic torquer of the
spacecraft, and further comprising the step of:

determining the magnetic torquer command selection as
a function of the estimated attitude of the
spacecraft.

20. The method as in claim 10 further comprising
providing a measure of the attitude pointing performance by
the steps of:

fixing a convex pattern on the ground to define a
pointing region;

projecting a vector fixed in the spacecraft body
coordinates to the ground; and

determining if said vector intersects the ground
within the interior of said convex pattern defining
the pointing region.





14

21. The method as in claim 10 further comprising,
when the momentum wheels are turned off, the steps of:

resetting one wheel tachometer such that the wheel
speed measurement is zero;

using the other wheel tachometer to read the
tachometer underspeed limit;

using the difference between the two measured wheel
speeds to cause the magnetic torquers to provide
current output (control torque) while setting the
polarity and magnitude by user command;

providing roll thruster torques by manual thruster
pulse commands in 10 ms quanta of pulse widths;

providing pitch/yaw thruster torques by differentially
setting a commanded speed for each wheel, such that
cycling through on-orbit wheel control mode will
immediately effect a momentum unload pulse; and

setting the momentum unload pulse by user command.

22. The method as in claim 10 comprising the further
step of discarding the standard roll, pitch, and yaw
independent axis control to achieve minimum fuel use.

Description

CA 02199436 1997-09-OS
PA-96019
EARTH BASED SPACECRAFT ORBIT AND ATTITUDE CONTROL
USING A LOOK-AHEAD THRUSTER SELECTION LOGIC
AND MAGNETIC TORQUERS
BACKGROUND OF THE INVENTION
Field of the Invention:
The present invention relates to Earth satellite control
systems and more particularly to a ground-based control
system which interfaces with the satellite's ground station
to automatically continue attitude control of a spacecraft
in the event of momentum wheel failure.
Prior Art Problem to be Solved:
Various systems are known for the control of Earth-orbiting
spacecraft or satellites involving the use of hardware and
software for executing complex control algorithms that
utilize on-board attitude control equipment such as
momentum wheels, gyros, magnetic torquers, and strings of
thrusters to maintain and alter the satellite's orientation
and orbit. One example of system of this type is the
ground loop automatic control system (GLACS) of Telestar of
Canada. Pertinent features of this system are set forth in
the attached pages identified at the lower left corner as
9 - 12.
However, problems are posed with such systems when there
are failures in the proper operation of an on-board control
device. In such event, the equipment and algorithms must
be quickly adjusted to deal with the changed situation and
maintain the satellite in its desired attitude and orbit.
As the on-board control equipment and fuel are self-
contained, the required adjustments should be efficient in
execution.
Obj ects
It is accordingly an object of the present invention to


CA 02199436 1997-09-OS
PA-96019 2
provide a system that allows the continued attitude control
and simultaneous partial orbit control of a spacecraft in
the event of momentum wheel failure with the novel and
efficient use of thrusters, magnetic torquers and on-board
gyroscopes.
SUMMARY OF THE INVENTION
The present invention involves a spacecraft controller
system that is physically located on the Earth and
interfaces with the satellite ground station receiving
spacecraft telemetry from the downlink baseband equipment.
This controller system automatically sends spacecraft
commands through the command uplink baseband equipment
using a ground loop control (GLC) which controls the
attitude of an orbiting spacecraft, and achieves partial
orbit control, using commanded thruster firings. It allows
the continued attitude control of a spacecraft in the event
of momentum wheel failure with the novel and efficient use
of thrusters, magnetic torquers and on-board gyroscopes.
A cooperative approach of using all available thrusters, of
both the primary and redundant strings, provides greater
fuel savings.
BRIEF DESCRIPTION OF THE DRAWINGS
Figure 1 is a plot of the satellite's 1/2 cone angle.
Figure 2 is a plot of the satellite's yaw angle.
Figure 3 illustrates a set of earth-centered orbits with
thruster vectors and the Sun orientation to the right.
Figure 4 is an illustrative plot of Elevation (deg) versus
Azimuth (deg) for a non-circular pointing region.
DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT
An automatic control system in accordance with the
invention is composed of an Earth-based computer, including
satellite control software, that interfaces with the
satellite ground station receiving spacecraft telemetry


CA 02199436 1997-09-OS
PA-96019 3
from the downlink baseband equipment. The software
includes estimators, controller algorithms, redundancy
management routines, and communication routines. The
estimators take sensor data as extracted from the telemetry
frames and provide output estimates of the spacecraft
attitude and rates (including yaw angle, for which no
direct measurement is available) . They also take data from
the thruster torque table and provide thruster torque
estimates. The yaw is determined using a gyrocompassing
filter (kinetic or dynamic) driven by continuously running
gyros. The controller algorithm takes the various
estimates and produces thruster firing commands to maintain
the spacecraft attitude within a defined pointing region.
Other actuation devices, such as magnetic torquers, may
also be used in the process, in which event, commands are
also sent to control these other actuation devices.
Appropriate control torques are provided by thrusters, and
magnetic torquers if available. If the momentum wheels are
turned off, one wheel tachometer is reset such that the
wheel speed measurement is zero. The other wheel
tachometer is not reset, and reads the tachometer
underspeed~limit. The difference between the two measured
wheel speeds allows the magnetic torquers to provide
current output (control torque) with the polarity and
magnitude being set by user command. Roll thruster torques
are provided by manual thruster pulse commands available
only in 10 ms quanta of pulse widths. Pitch/yaw thruster
torques are provided by differentially setting commanded
speed for each wheel, such that cycling through on-orbit
wheel control mode will immediately effect a momentum
unload pulse. The unload pulse is set by user command and
can be less than 10 ms.
To achieve minimum fuel use, the standard roll, pitch, and
yaw independent axis control is discarded. The attitude is
controlled to a pointing polyhedron parameterized using the


CA 02199436 1997-09-OS
PA-96019 4
region's origin offset (with respect to the nominal antenna
boresight), along with a series of (radius, polar angle)
pairs. The region defines the -3dB coverage for the
antenna beam pattern, and thus relates directly to
communications performance. The definition of the pointing
region includes as a subset the more traditional half-cone
angle region, where the origin offset is zero and the
radius is constant for all polar angles.
The spacecraft is set up in an automatic on-board thruster
control mode which acts as a safety mode if commanding from
the GLC should be interrupted. Sensor bias commands are
sent periodically to the spacecraft to center the on-board
mode deadband to the current attitude. These bias commands
also serve as a GLC "heartbeat" that can be detected at the
backup site and the absence of which indicates a problem at
the primary site.
Thrusters are fired singly or in pairs, as determined by
the thruster selection algorithm. This algorithm
propagates the dynamic model of the spacecraft attitude
forward in time to find the thruster or thruster pair that
results in the least fuel use rate before the spacecraft
reaches the edge of the attitude pointing region, with the
total thruster fuel use being weighted using the estimated
time until next firing.
An orbit model, propagating from the last Orbit
Determination result, determines the current orbit
resulting from the commanded thruster activity. Since
there will be a relatively large number of thruster firings
each day, there will be sufficient thruster activity to
allow East-West stationkeeping (EWSK), of both eccentricity
vector and radial component control, on a continuous basis.
That is, the thruster selection algorithm is biased to the
selection of the appropriate thruster to accomplish both
attitude control and an EW orbital correction in the


CA 02199436 1997-09-OS
PA-96019 5
desired direction. In this way, the need for separate EWSK
maneuvers will be greatly reduced or completely disappear,
saving overall fuel use. North-South Stationkeeping (NSSK)
requires more thruster activity than the GLC will provide,
but can still be helped by biasing the thruster selection
algorithm in the desired North/South direction depending on
the orbit inclination.
The GLC system of the invention may be made up of two fault
tolerant computers which independently run the satellite
control software. One of the computers is designated
primary, and as such is authorized to send spacecraft
attitude control commands. The secondary computer acts as
a hot backup, running the satellite control software in
parallel with the primary computer, but not actually
sending commands. If the primary computer fails for any
reason, logic in the secondary computer, or in an
independent monitor unit, will transfer commanding
authority immediately and automatically to the secondary
computer.
The primary computer constantly sends commands to the
spacecraft. Even if it is not required to send attitude
control commands, the primary computer will send the
periodic heartbeat commands which can be detected in
telemetry and will imply the state of health of the command
link. The backup computer can detect the heartbeat
commands, and can infer failure of the primary computer if
the heartbeat command disappears. Each computer also
monitors an active status line in the other computer to
detect computer failure.
The GLC system of the invention may be situated between the
system operator computer and the baseband command
transmitter so as to intercept all operator generated
spacecraft commands. If the GLC controller is sending a
command sequence, it will delay any non time-critical


CA 02199436 1997-09-OS
PA-96019 6
operator generated command for the duration of that
sequence. The GLC computers communicate with the operator
computers to receive parameter updates and send logger,
warning, or alarm messages.
The GLC computers control switches in the baseband
telemetry and command strings to switch to backup baseband
equipment strings if a failure is detected in any baseband
equipment unit. In this way, even though individual system
components can fail, system redundancy is achieved and the
GLC has a very high system availability.
If a backup ground station is available, a backup GLC unit
can be located there and can monitor the primary site using
a combination of the telemetered command carrier receiver
signal strength and the telemetered command execute history
of the heartbeat commands.
Attached are lists of items regarding features of the
invention relating to and entitled CONTROL LAW, FUEL
OPTIMAL THRUSTER SELECTION, YAW ESTIMATOR, ORBIT ESTIMATION
AND CONTROL, and EAST/WEST ~V SCHEDULE. These items
illustrate the following features of the invention.
CONTROL LAW:
There are two methods of choosing thruster commands, i.e.,
fuel optimal and backup. The fuel optimal approach
involves the steps of generating a list of feasible
single/double fire commands and choosing the best command
based on fuel efficiency and meeting the ~V schedule
generated by orbit control. A non-circular pointing region
can be used. An illustrative plot of Elevation (deg)
versus Azimuth (deg) for a non-circular pointing region is
shown in Figure 4. The backup approach is only used if
there is no fuel optimal solution and involves generating
a list of single/double/triple fire commands based on
rate/attitude switching curves, and choosing the command
closest to meeting the 0V schedule. And, Bang-Bang roll


CA 02199436 1997-09-OS
PA-96019 7
and yaw magnetic torquer commands are based on the
rate/attitude switching curve.
FUEL OPTIMAL THRUSTER SELECTION:
A list of feasible thruster commands are generated, i.e.,
the 1/2 cone angle and yaw attitude is brought in bounds.
Solutions are ranked by ratio of time in bounds and fuel
use. The solution that best meets the ~V schedule from
orbit control is chosen. Illustrative 1/2 cone angle and
yaw angle plots are shown in Figures 1 and 2.
YAW ESTIMATOR:
Yaw estimation is required in the absence of a yaw sensor.
A ground based gyrocompassing yaw estimator may be used.
The yaw attitude determination algorithm involves a 4th
order Kalman filter using DSS, roll ES, and roll/yaw DIRA,
and a 4th order Kalman filter ususing roll ES and roll/yaw
DIRA. Transitions between the two filters are determined
depending on Sun presence.
ORBIT ESTIMATION AND CONTROL:
An internal orbit model based on the last orbit
determination and thruster commands issued by the control
law is propogated. Using the daily averages of the
internal orbit state, it is assessed as to how ~V should be
directed for the next day. Desired ~V accumulations are
scheduled for the control law to meet. Orbit control can
be set to orbit neutral if no ~V accumulation is desired.
EAST/WEST ~V SCHEDULE:
Figure 3 illustrates a set of earth-centered orbits along
with thruster vectors and the Sun orientation to the right.
The thruster vectors are used to correct the satellite's
eccentricity with respect to the Sun as shown.
It will be seen from the foregoing description that a
ground loop automatic contol system is disclosed that
provides an efficient and novel thruster selection and
control logic and simultaneous east-west orbit and
spacecraft attitude control. Additionally, yaw estimates
are obtained more simply and more accurately.

Representative Drawing