Spacecraft acquisition of sun pointing

A system and method for acquiring Sun pointing in a three-axis stabilized spacecraft including slewing at least one solar wing until the Sun is detected, determining an initial Sun vector, performing one or more rotations of the spacecraft body so as to bring the instantaneous Sun vector coincident with a preferred final Sun vector, at least one rotation about an optimal axis, and slewing at least one solar wing to a preferred attitude relative to the Sun. In one embodiment, the optimal axis is chosen so as to minimize the time required to achieve an optimal thermal attitude. In another embodiment, the optimal axis is chosen so as to minimize the time required to align the instantaneous Sun vector with the final desired Sun vector. In a further embodiment, the optimal axis is chosen so as to maintain Earth lock.

TECHNICAL FIELD 
The present invention relates in general to attitude control of a 
spacecraft and in particular to acquisition of Sun pointing by a 
three-axis stabilized spacecraft with solar wings. 
BACKGROUND ART 
Many three-axis stabilized spacecraft utilize solar wings to generate 
electrical energy. In order to maximize the energy output, the solar wings 
must be properly oriented towards the Sun. Spacecraft also often use 
radiators to eliminate waste heat. These radiators must not be oriented 
towards the Sun in order to properly function. Spacecraft may also have 
mission objectives that require a specific attitude relative to the Sun 
and other celestial bodies such as the Earth. 
Time and energy is expended by a spacecraft when modifying its attitude. 
Minimizing time avoids thermal problems and reduces battery drainage. 
Minimizing energy expenditure conserves the limited amount of fuel 
normally carried by a spacecraft. 
Current methods acquire a solar-pointing attitude by slewing the spacecraft 
one axis at a time. A typical solar orientation may require three phases: 
pitch search, keyhole maneuver and yaw search. Consider, for example, the 
following typical sequence of actions and associated execution times. 
The spacecraft first performs a pitch scan to search for the Sun. A 360 
degree pitch scan at a typical search rate of 0.75 degrees per second 
takes 8 minutes. During the pitch search phase, the solar wings are slewed 
to face the roll axis. If the pitch search is unsuccessful, the Sun is 
assumed to be in the sensor blind spot or keyhole. In this case, a one 
time 75 degree yaw slew at a typical search rate of 0.25 degrees per 
second takes 5 minutes. Adding 8 minutes for a second pitch search yields 
13 minutes for the keyhole maneuver. Once the spacecraft locks onto the 
Sun in pitch, the spacecraft performs a yaw scan while maintaining 
Sun-lock in pitch. A 45 degree yaw slew at a typical search rate of 0.25 
degrees per second takes 3 minutes. The total maximum time for the above 
procedures is 24 minutes. 
Modifications to the above sequence have been proposed, but none 
significantly reduce the worst-case acquisition time. 
An additional problem with existing solar acquisition methods is that the 
large search rate about the body pitch axis generally requires large 
momentum wheel storage requirements. 
A further problem occurs in Sun-nadir steered spacecraft, where Earth-lock 
can be lost during the solar acquisition sequence. 
SUMMARY OF THE INVENTION 
As such, one object of the present invention is to quickly place the 
spacecraft in a safe power state. This reduces the overall time required 
for Sun acquisition and improves chances for survival if attitude 
anomalies occur following an eclipse when the initial battery state of 
charge is low. 
Another object of the present invention is to quickly place the spacecraft 
in a safe thermal state by minimizing Sun impingement on the radiators. 
Another object of the present invention is to reduce the storage 
requirements in momentum wheels by eliminating the large search rate about 
the body pitch axis. 
A further object of the present invention is to maintain Earth lock while 
acquiring the Sun. 
In carrying out the above objects and other objects and features of the 
present invention, a method is provided to rotate a spacecraft about an 
optimal axis, thereby minimizing the time and energy required to orient 
the spacecraft to a desired Sun angle. As a consequence, the spacecraft is 
quickly placed in a safe power and thermal attitude. 
In one embodiment, a solar search is made by slewing the solar wings in 
opposite directions about the pitch axis until the Sun is incident on at 
least one wing. This is followed by rotating the spacecraft body about an 
axis determined by the cross product of the instantaneous Sun vector with 
a projection of the instantaneous Sun vector into the roll-yaw plane. This 
brings the spacecraft into a safe thermal attitude and moves at least one 
wing into optimal solar attitude. The spacecraft is then rotated about the 
pitch axis to bring the body into the desired solar attitude. At the same 
time, each solar wing is rotated with respect to the spacecraft body so as 
to maintain or obtain Sun lock. 
A modification of this embodiment executes the roll-yaw body rotation, the 
pitch body rotation, and final solar wing slew simultaneously. 
In another embodiment, a solar search is made by slewing the solar wings in 
opposite directions. This is followed by rotating the spacecraft body 
about an axis determined by the cross product of the initial Sun vector 
with the desired final Sun. During body rotation, the solar wings are 
rotated to maintain or obtain solar lock. 
In still another embodiment, Earth lock is maintained while acquiring the 
Sun. A solar search is made by slewing the solar wings in opposite 
directions. This is followed by rotating the spacecraft about the Earth 
vector until the Sun vector lies in the roll-yaw plane. During body 
rotation, the solar wings are rotated to maintain or obtain solar lock. 
In any of the above embodiments, solar wings that may rotate in the same 
direction may be substituted for solar wings that may rotate in opposite 
directions. 
A system is also provided in accordance with the present invention. Sensors 
on the north and south solar wings determine Sun position relative to the 
spacecraft. Using this information, a control logic determines actions for 
north and south solar wing drives, a reaction wheel system, and thrusters. 
The above objects and other objects, features, and advantages of the 
present invention are readily apparent from the following detailed 
description of the best mode for carrying out the invention when taken in 
connection with the accompanying drawings.

BEST MODE FOR CARRYING OUT THE INVENTION 
Referring now to FIG. 1, a graphical representation of a spacecraft that 
may use the present invention is shown. A spacecraft, shown generally as 
20, has an orthonormal body reference frame shown as roll axis x, pitch 
axis y, and yaw axis z. Spacecraft 20 includes a main body 21. One face of 
body 21 such as, for example, the positive x face, is to be oriented 
towards the Sun. 
A north solar wing 22 is connected to the negative y face of body 21 so as 
to rotate about an axis parallel to the pitch axis relative to body 21. 
North wing 22 is caused to rotate by north solar wing drive 23. A south 
solar wing 24 is connected to the positive y face of body 21 so as to 
rotate about an axis parallel to the pitch axis relative to body 21. South 
wing 24 is caused to rotate by south solar wing drive 25. In a preferred 
embodiment solar wings 22 and 24 can be rotated in either the same 
direction or in opposite directions simultaneously. In another embodiment, 
solar wings 22 and 24 must be rotated in the same direction. 
In a preferred embodiment, a pair of array mounted slit Sun sensors is 
used. A north Sun sensor array 26 is rotatively mounted on north solar 
wing 22. A south Sun sensor array 28 is rotatively mounted on south solar 
wing 24. Each sensor has at least one slit sensor with pitch as its 
sensitive axis and at least one slit sensor with its sensitive axis 
orthogonal to pitch. In another embodiment, the amount of electrical 
current generated from solar cells mounted to solar wings 22 and 24 is 
used to determine the direction of the Sun. The invention disclosed herein 
is not dependent on the precise method for determining the direction of 
the Sun. Various types of solar sensors such as single slit Sun sensors 
mounted on one or both wings may be used as will be recognized by one or 
ordinary skill in the art. 
Spacecraft 20 includes actuators for affecting attitude maneuvers. These 
are shown as thrusters 30 and reaction wheel system 40. Other actuators 
may be used as is known in the art. 
Spacecraft 20 may include radiators for eliminating waste heat. A north 
radiator 32 is on the positive y side of spacecraft body 21 and a south 
radiator 34 is on the negative y side of spacecraft body 21. 
Spacecraft 20 may include a desired attitude towards a celestial target. 
For example, the mission may require one face of spacecraft body 21 to be 
oriented towards Earth. For the spacecraft depicted in FIG. 1, a static 
Earth sensor 36 is mounted on the positive z face of body 21 with 
boresight parallel to the yaw axis. 
Spacecraft 20 may also include attitude determination sensors such as, for 
example, three-axis gyroscope package 38. These sensors assist in 
orienting and stabilizing spacecraft 20, as is well known in the art. 
A control logic 42, in communication with solar wing drives 23 and 25, Sun 
sensor arrays 26 and 28, thrusters 30, Earth sensor 36, gyroscope package 
38, and reaction wheel system 40, accepts input from sensors and produces 
signals to actuators and drives so as to carry out the present invention. 
Referring now to FIGS. 2, 4 and 6, flow diagrams of three embodiments of 
the present invention are shown. As will be appreciated by one of ordinary 
skill in the art, the operations illustrated in each flow diagram are not 
necessarily sequential operations. Similarly, operations may be controlled 
by software, hardware, or a combination of both. The present invention 
transcends any particular implementation and each of various embodiments 
is shown in a sequential flow chart form for ease of illustration. 
Referring now to FIGS. 3, 5 and 7, polar plots are shown. In each plot, 
spacecraft 20 is thought of as being at the center of the sphere with axes 
on the plot corresponding to the body reference frame. The y axis extends 
through the south pole and both the x and z axes extend through the 
equator. For all embodiments shown, the Sun starting position relative to 
the spacecraft body reference is shown as 70 and the initial Sun vector, 
S.sub.1, is shown as 72. 
Referring now to FIG. 2, a flow diagram, and to FIG. 3, a polar plot, an 
embodiment of the present invention for minimizing the time required to 
obtain optimal thermal attitude is shown. 
Spacecraft 20 is first stabilized, as in block 50 of FIG. 2. This results 
when angular rates about all three axes have been nulled. Methods for 
stabilizing a three-axis stabilized spacecraft are well known in the art. 
If, at the end of stabilization, solar acquisition has occurred, the method 
has completed as shown in block 52. 
If solar acquisition has not occurred, a pitch search is performed, as in 
block 54. In one search method, the solar wings are slewed about the pitch 
axis in the same direction until the Sun is detected. A 360 degree pitch 
scan at a typical solar wing slew rate of 2.45 degrees per second takes 
approximately 2.5 minutes. 
A preferred search method slews each wing about the pitch axis in an 
opposite direction. This will require at most a 180 degree pitch scan 
requiring no more than 1.25 minutes at 2.45 degrees per second before the 
Sun is acquired by at least one wing. In addition to requiring less 
worst-case time, an additional benefit of slewing the wings in opposite 
directions is that rotational momentum remains unchanged by the operation. 
Referring now to FIG. 3, a polar plot of an example of an embodiment 
minimizing the time required to obtain optimal thermal attitude is shown. 
North solar wing 22 is rotated in direction 74 while south solar wing 24 
is rotated in direction 76. Ideally, because spacecraft body 21 is not 
moved, the Sun remains in the same relative position as indicated by Sun 
vector S.sub.1. 
An additional search step may be required for Sun sensors with a blind 
region in pitch such as single slit Sun sensors, a Sun sensor array on 
only one wing, or solar cell current sensors. If the pitch search does not 
locate the Sun, a keyhole maneuver is started. In this case, a slew about 
the yaw axis followed by an additional pitch search is required. A 75 
degree yaw slew at a typical search rate of 0.25 degrees per second takes 
5 minutes. 
Once the pitch search is completed, one or both solar wings have acquired 
the Sun as best as is possible without rotating spacecraft body 21. 
Referring again to FIG. 2, if the Sun is acquired, implying that the Sun 
vector started aligned with the x axis, the method has completed as shown 
in block 56. 
If the Sun is not acquired, an axis about which to rotate the spacecraft 
body is determined as shown in block 58. This axis is designed to place 
the spacecraft in a safe thermal attitude as quickly as possible. Safe 
thermal attitude occurs if the Sun vector lies in the roll-yaw plane since 
north radiator 32 and south radiator 34 are normal to the pitch axis. The 
shortest rotation placing the Sun vector in the roll-yaw plane moves 
S.sub.1 to its projection onto the roll-yaw plane, S.sub.2. This rotation 
is about an axis in the x-z plane normal to the Sun line or, 
mathematically, is the cross product of S.sub.1 and S.sub.2. 
A roll-yaw slew is then performed by the spacecraft as in block 60. At the 
end of this maneuver, the Sun vector is in the x-z plane. The roll-yaw 
slew is limited to 90 degrees. A typical slew rate of 0.25 degrees per 
second results in a maximum time of 6 minutes. 
If solar wings 22 and 24 were rotated in the same direction during the 
pitch search, both wings have now acquired the Sun. This places spacecraft 
20 in safe power attitude as well as safe thermal attitude. If solar wings 
22 and 24 were rotated in opposite directions, at least one solar wing has 
acquired the Sun. 
A safe thermal and power attitude has been obtained in no more than 8.5 
minutes for a typical spacecraft with solar wings 22 and 24 rotated in the 
same direction during the pitch search. 
Referring now to FIG. 3, the roll-yaw maneuver is shown on the polar plot. 
S.sub.2, shown as 78, is the projection of initial Sun vector 72 onto the 
x-z plane. The cross product of vector 72 and vector 78 is axis of 
rotation R, shown as 80. Direction of rotation about R, shown as 82, is 
determined so as to minimize the angle from S.sub.1 to S.sub.2. The path 
of the Sun vector during rotation is shown as 84. 
Referring again to FIG. 2, if the Sun is acquired on the desired spacecraft 
axis, implying that the Sun vector started in the x-y plane, the method is 
completed as shown in block 62. 
If the Sun is not acquired, a pitch slew is performed, as shown by block 
64. Spacecraft 20 is rotated about the y axis until spacecraft body 21 is 
facing in the desired attitude. This implies that the instantaneous Sun 
vector is coincident with the preferred final Sun vector. For the 
spacecraft embodiment described previously in FIG. 1, a vector normal to 
the positive x face is pointed at the Sun. 
In a preferred embodiment, solar wings 22 and 24 are rotated during pitch 
slew. If either wing has previously acquired the Sun, it is slewed to 
remain Sun-pointing while spacecraft body 21 is rotated. If either wing 
has not acquired the Sun, it is rotated until the Sun is acquired. In 
another embodiment, solar wings 22 and 24 are rotated in a separate step 
following the pitch slew. 
Referring now to FIG. 3, the pitch slew maneuver is shown on the polar 
plot. Rotation about the y axis, shown as 86, moves the Sun vector from 
S.sub.2 78 to S.sub.3 88. The path of the Sun vector during the maneuver 
is shown as 90. 
In alternate embodiments, the roll-yaw slew and the pitch slew may be 
executed sequentially as implied by paths 84 and 90 or may be executed 
simultaneously. 
Referring now to FIG. 4, a flow diagram, and to FIG. 5, a polar plot, an 
embodiment of the present invention for minimizing the time required to 
align the instantaneous Sun vector with the preferred final Sun vector is 
shown. This embodiment can be used on spacecraft with Sun sensors that 
provide two axes of information. For comparison purposes, initial Sun 
vector S.sub.1, shown as 72, and preferred final Sun vector S.sub.3, shown 
as 88, are the same in both FIG. 5 and FIG. 3. 
Referring now to FIG. 4, spacecraft 20 must be stabilized as shown in 100. 
If the Sun is acquired as a result of stabilization, the method has 
completed as shown in 102. 
A pitch search is performed, as shown in 104. This is identical to the 
pitch search associated with block 54 in FIG. 2. At the end of the pitch 
search, the current Sun vector is coincident with initial Sun vector 
S.sub.1, shown as 72. If the Sun is acquired as a result of the pitch 
search, the method has completed as shown in 106. 
If the Sun is not acquired at the end of the pitch search, a general axis 
of rotation is determined, as shown by block 108. This axis, T, is chosen 
to minimize the angular distance between initial Sun vector 72 and desired 
final Sun vector 88. The axis T is found by taking the cross product of 
S.sub.1 and S.sub.3. 
A general slew is performed, as shown in 110. The slew is about axis of 
rotation T in a direction that minimizes the angular distance between 
S.sub.1 and S.sub.3. This angular distance is limited to at most 180 
degrees. 
In a preferred embodiment, solar wings 22 and 24 are rotated during the 
general slew. If either wing has previously acquired the Sun, it is slewed 
to remain Sun-pointing while spacecraft body 21 is rotated. If either wing 
has not acquired the Sun, it is rotated until the Sun is acquired. In 
another embodiment, solar wings 22 and 24 are rotated in a separate step 
following the general slew. 
Referring now to FIG. 5, an example is illustrated. Axis for the general 
slew T is shown as vector 120. The slew direction is shown as 122. The 
path of the instantaneous Sun vector during the maneuver is shown as 124. 
Referring now to FIG. 6, a flow diagram, and to FIG. 7, a polar plot, 
another embodiment of the present invention is shown. This embodiment can 
be used to maintain a lock on a celestial body such as the Earth while 
acquiring the Sun. For comparison purposes, initial Sun vector S.sub.1, 
shown as 72, is the same in both FIG. 7 and FIG. 3. 
Referring now to FIG. 6, spacecraft 20 must be stabilized as shown in 150. 
If the Sun is acquired as a result of stabilization, the method has 
completed as shown in 152. 
A solar wing pitch search is performed, as shown in 154. This is identical 
to the pitch search associated with block 54 in FIG. 2. At the end of the 
pitch search, the current Sun vector is coincident with initial Sun vector 
S.sub.1, shown as 72. If the Sun is acquired as a result of the pitch 
search, the method has completed as shown in 156. 
In order to maintain lock, spacecraft 20 must rotate about the Earth 
vector, designated by lock axis Q. The lock slew is performed, as in block 
158, from Sun vector position S.sub.1 until the instantaneous Sun vector 
is in the x-z plane. The final Sun vector position is designated S.sub.4. 
Unless the angle between the Earth vector and Sun vector correspond with 
the angle between the Earth sensor boresight and the normal to the desired 
Sun-facing spacecraft body surface, an optimal safe thermal attitude 
cannot be obtained. 
In a preferred embodiment, solar wings 22 and 24 are rotated during the 
lock slew. If either wing had previously acquired the Sun, it is slewed to 
remain Sun-pointing while spacecraft body 21 is rotated. If either wing 
had not acquired the Sun, it is rotated until the Sun is acquired. In 
another embodiment, solar wings 22 and 24 are rotated in a separate step 
following the lock slew. 
Referring now to FIG. 7, an example of solar acquisition with Earth lock is 
shown. Lock axis Q is shown as 170. The lock slew direction is shown as 
172. The path of the instantaneous Sun vector during the maneuver is shown 
as 174. The final Sun vector position is shown as 176. 
Referring now to FIG. 8, a block diagram of a system for effecting a 
spacecraft maneuver according to the present invention is shown. North 
solar wing Sun sensor 26 and south solar wing Sun sensor 28 are operative 
to detect the position of the Sun relative to spacecraft 20. North solar 
wing position sensor 192 and south solar wing position sensor 194 are 
operable to detect the position of the respective wing relative to 
spacecraft body 21. Sun sensor and wing position sensor measurements are 
read by control logic 42, which is in communication with solar wing Sun 
sensors 26 and 28, solar wing position sensors 192 and 194, as well as 
north solar wing drive 23, south solar wing drive 25, reaction wheel 
system 40, and thrusters 30. Control logic 42 is operative to perform the 
method described with regards for FIGS. 2 through 7 above. Control logic 
42 develops solar wing drive commands for solar wing drives 23 and 25, 
wheel commands for reaction wheel system 40, and thruster commands for 
thrusters 30, so as to obtain a desired attitude for spacecraft 20 and 
solar wings 22 and 24. Solar wing angles, wheel torques and thruster 
pulses affect spacecraft attitude through spacecraft dynamics 190. 
While the best mode for carrying out the present invention has been 
described in detail, and several alternative embodiments have been 
presented, those familiar with the art to which this invention relates 
will recognize various alternative designs and embodiments within the 
scope and spirit of the present invention as defined by the following 
claims.