Data processing: vehicles – navigation – and relative location – Vehicle control – guidance – operation – or indication – Aeronautical vehicle
Reexamination Certificate
2002-04-01
2003-05-27
Camby, Richard M. (Department: 3661)
Data processing: vehicles, navigation, and relative location
Vehicle control, guidance, operation, or indication
Aeronautical vehicle
C244S171000
Reexamination Certificate
active
06571156
ABSTRACT:
BACKGROUND OF THE INVENTION
The present invention generally relates to power acquisition for spacecraft and, more particularly, to algorithms and methods of power acquisition for spacecraft in solar panel wing-deployed configuration.
Prior art spacecraft typically acquire the sun for power safety using sun sensor assemblies, for example, either wide field of view (WFOV) sun sensor or narrow field of view (NFOV) slit sun sensor. Prior art methods of sun, or power, acquisition using sun sensor assemblies often serve dual purposes of acquiring the sun to achieve an accurate spacecraft body-to-sun attitude knowledge and placing the solar wings normal, i.e., orthogonal, to the sun-line, i.e., a line from the spacecraft to the sun, to maximize solar power. A prior art method of power acquisition is disclosed in U.S. Pat. No. 5,255,879, issued Oct. 26, 1993, entitled “Three Axes Stabilized Spacecraft and Method of Sun Acquisition”, and assigned to the assignee of the present invention. Methods, however, that use sun sensor assemblies for power acquisition can require expensive electronic hardware, such as buffer channel hardware, and can require precise maneuvering to maintain a perfect slew along a spacecraft axis that is not very close to one of the principal axes. In a spacecraft configuration with solar wing or wings deployed where accurate spacecraft body-to-sun attitude knowledge is not needed, but rather it is only needed to keep the spacecraft power safe and thermally safe, a simple and robust algorithm for power safety, which, for example, avoids the difficulties just described with sun sensor buffers and precise maneuvering, would be preferable.
For increased dependability and safety, modern spacecraft may rely on more than one spacecraft control processor (SCP) to control the functions and attitude of the spacecraft. A situation where the spacecraft switches control from one SCP to another is referred to as “toggle”. An example toggle situation occurs when a first spacecraft control processor, SCP
1
and its connected sensors, is in active control of the spacecraft and a second spacecraft control processor, SCP
2
and its connected sensors, is in standby, i.e., is powered off. SCP
1
encounters some problems detected by a monitor and decides to turn off SCP
1
and turn on SCP
2
. SCP
2
then takes over control of the spacecraft.
With the advent of high power-dissipation payloads in spacecraft, comes a large radiator to maintain payload temperature below the high operating limit. There are various situations in the operation of the spacecraft when power to the spacecraft payload may be turned off. After a situation where the spacecraft toggles from one spacecraft control processor to another, power to the spacecraft payload may be turned off. In post-toggle, i.e., after a toggle has occurred, with the spacecraft in payload-off configuration, the radiator continues to dissipate a large amount of power and, thus, requires more power for the heater to maintain the spacecraft above survival temperature compared to spacecraft with lower power-dissipation payloads. If the payload remains on, the spacecraft needs even more power to survive and the situation is worse than payload off. A power acquisition algorithm, referred to as safe-hold algorithm, may be used post-toggle to ensure that the spacecraft achieves power, thermal, and momentum safety. Due, for example, to the increased post-toggle power requirements for the heater with high power-dissipation payloads, the previously used safe-hold algorithm is no longer adequate to ensure power, thermal, and momentum safety of the spacecraft under all conditions or scenarios.
For example, a previously used safe-hold algorithm for post-toggle safety rotates the spacecraft body along an axis in the xz plane (perpendicular to the wing axis) and places the solar wing in sun searching mode. In the worst case, referred to as “paddle wheel scenario” (or cases close to this scenario), the received wing power is a sinusoidal curve with theoretical average wing power of 2/&pgr;, which is approximately 64%. The paddle wheel scenario occurs when the sun line and the spacecraft spin axis are orthogonal to each other, resulting in the sun being in the “keyhole”, i.e., sun in wing axis or the body y-axis, twice per revolution. Due to bus shadowing effect and the fact that the wing will take more than 360 seconds to rotate 180 degrees to search for the sun every time the sun passes the keyhole position, the received wing power is lower than the theoretical value. The useful wing power is further limited by the battery charge rate so that some of the power is discarded because the battery cannot accept all received power for charging. Because the spacecraft requires power higher than this useful wing power for safety, the previously used safe-hold algorithm no longer renders power, thermal, and momentum safety for the spacecraft.
FIG. 1
shows a phase transition block diagram of a previously used power acquisition algorithm. Algorithm
100
, shown in
FIG. 1
, uses wide field of view (WFOV) sun sensor assembly (SSA), and is applicable to both wing stowed or wing deployed spacecraft configuration. Algorithm
100
includes an initialization phase
102
, null rate phase
104
, pitch search phase
106
, keyhole search phase
108
, steer to null phase
110
, sun hold phase
112
, and fault hold phase
114
. The arrows shown in
FIG. 1
indicate the control logic, or flow of control, between phases of algorithm
100
. For example, from initialization phase
102
control may pass either to null rate phase
104
under a “normal” state of affairs, or to fault hold phase
114
under abnormal conditions, such as algorithm
100
“timing out” before power acquisition has been achieved. As seen in
FIG. 1
, pitch search phase
106
may be followed by keyhole search phase
108
or may proceed directly to steer to null phase
110
. As described above, algorithm
100
may leave the spacecraft in the paddle wheel scenario where useful wing power is inadequate to place the spacecraft in power, thermal, and momentum safety. Thus, algorithm
100
represents a complicated algorithm for power acquisition, which is not very robust, i.e., is prone to failure, or entering the fault hold phase
118
state, under many conditions.
As can be seen, there is a need for a simple, robust algorithm for power acquisition for power, thermal, and momentum safety for high heater-power spacecraft, and especially for post-toggle power, thermal, and momentum safety for spacecraft with high heater-power demand in payload-off configuration. There is also a need for an algorithm for spacecraft power acquisition for power, thermal, and momentum safety that can rely on direct power measurement from the solar wings rather than the indirect power measurement provided by the use of sun sensor assemblies such as WFOV sun sensors.
SUMMARY OF THE INVENTION
The present invention provides a simple, robust algorithm for power acquisition for power, thermal, and momentum safety for high heater-power spacecraft, and especially for post-toggle power, thermal, and momentum safety for spacecraft with high heater-power demand in payload-off configuration. The present invention also provides an algorithm for spacecraft power acquisition for power, thermal, and momentum safety that can rely on direct power measurement from the solar wings rather than the indirect power measurement provided by the use of sun sensor assemblies such as WFOV sun sensors.
In one aspect of the present invention, an algorithm for a spacecraft includes a wing sun search phase, an xz slew phase, and a safe hold phase. In the wing sun search phase, a solar wing current is monitored against a low current threshold while a solar wing sun search is performed, the xz slew phase is entered when the solar wing current stays below the low current threshold until a solar wing current persistence timer expires, and the safe hold phase is entered when the solar wing current does not stay below the low current threshold until the solar wing current p
Ketkar Umesh
Mansour Sadek W.
Sun Sheng N.
Wang Grant
Camby Richard M.
DiPinto & Shimokaji P.C.
The Boeing Company
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