GB2318888A - Solar panel mounted sun sensor and three-axis attitude control - Google Patents

Abstract

Precise three-axis satellite attitude control and frequent gyroscope (66) calibration during operational orbit is performed using additional processing on the outputs (60) of a simple sun sensor (46) mounted on a satellite's solar panel (48). The sensor's (46) position on the face of the solar panel (48) causes its orientation to the satellite body to vary during orbit, yielding three-axis information over the course of an entire orbit. A signal processor unit (56) operates at relatively high gain when the sun is acquired and at relatively low gain during ascent and acquisition orbital phases. Solar modeling combined with orbit and sensor geometry are used to predict and measure sun declination, satellite attitude, and solar panel alignment components for each axis independently at points in orbit when the subject modeled component is uninfluenced by other model components.

Description

SOLAR PANEL MOUNTED SUN SENSOR AND THREE-AXIS ATTITUDE DETERMINATION THEREWITH TECHNICAL FIELD OF THE INVENTION The present invention relates generally to the field of space vehicle applications, and specifically to the field of satellite attitude and orbit determination and control.

BACKGROUND OF THE INVENTION Conventionally, an expensive array of coarse and flne earth and sun sensors and precise gyroscopes has been used to place a satellite into operational orbit and maintain the satellite at desired attitudes in the orbit.

A conventional satellite may employ two banks of small photoelectric analog sun sensors mounted on the yokes of solar panels to provide imprecise outputs describing the sun's position during ascent and acquisition phases of satellite activity. Several banks of complex, large, narrow-field-of-view, digital sun sensors mounted in the body of the satellite provide precise outputs describing the sun's position during operational orbit. Several narrow-field-of- view sensors may be used to ensure that one sensor will be facing the sun whenever the satellite is not in eclipse, and a complex switching mechanism may be used to ensure that the correct sensor is active given the current orbit position. Since the sensors are located in the satellite body, they experience significant temperature cycling and thermal regulation equipment is used to manage temperature.

The body-mounted, narrow-field-of-view sun sensors provide data to calibrate the satellite's gyroscopes fairly infrequently. Because of this, more precise, complex, low-drift gyroscopes have been used than would have been required if gyroscope calibration could have been performed more often.

Conventionally, imprecise and precise infrared earth sensors have been used to acquire the earth's position and to orient the satellite correctly with respect to the earth. These infrared earth sensors suffer inherent inaccuracies due to constant fluctuation of earth temperatures over different latitudes and cloud cover over the planet in general.

Thus, conventional satellites incorporate a multitude of expensive and unnecessarily redundant equipment to maintain a routine, operational orbit. This equipment utilizes complex thermal management systems which draw power and increase the mass, volume, and complexity of the satellite. At a time of ever-increasing mission functionality demands, this multitude of orbit maintenance equipment causes an increasing amount and proportion of satellite complexity, mass, and size to be dedicated to satellite control.

Consequently, a need exists for reducing the proportion of satellite mass and volume that exists simply to maintain orbit.

BRIEF DESCRIPIION OF ThE DRAWINGS A more complete understanding of the present invention may be derived by referring to the detailed description and claims when considered in connection with the Figures, wherein like reference numbers refer to similar items throughout the Figures, and: FIG. 1 shows two satellites orbiting the earth from a first equatorial perspective; FIG. 2 shows two satellites orbiting the earth from a second equatorial perspective; FIG. 3 shows a sun sensor mounted on an articulated solar panel mounted to the body of a satellite in accordance with a preferred embodiment of the present invention; FIG. 4 shows a schematic representation of photoelectric devices in a sun sensor mounted on a solar panel from an end perspective in accordance with a preferred embodiment of the present invention; FIG. 5 shows a schematic representation of the photoelectric devices in a sun sensor mounted on a solar panel from a side perspective in accordance with a preferred embodiment of the present invention; FIG. 6 shows a block diagram of a signal processing unit in accordance with a preferred embodiment of the present invention; FIG. 7 shows a curve representing sun sensor output as a function of solar angle calculated in accordance with a preferred embodiment of the present invention; FIG. 8 shows orbital positions from the perspective of the earth's poles; and FIG. 9 shows a flow chart of a process for achieving satellite attitude control with respect to a roll axis in accordance with a preferred embodiment of the present invention.

DETAILED DESCRIPIION OF THE PREFERRED EMBODIMENTS A solar-based satellite alignment system configured in accordance with the present invention replaces precise and imprecise earth sensors, precise gyroscopes, and fine sun sensors with variable gain amplification and computer processing. Comparison of the outputs of relatively simple solar sensors with models of sun declination, spacecraft dynamics, and solar panel alignment yields detailed attitude information that can be used to control satellite attitude.

FIG. 1 shows a satellite 20 orbiting the earth 22 from a first equatorial perspective.

Satellite 20 moves in an orbital direction defined by a roll axis 24, which is tangent to the earth, by a yaw axis 26, which points toward the center of the earth, and a pitch axis 28, which is perpendicular to both roll axis 24 and yaw axis 26. Throughout an orbit 30, roll axis 24, yaw axis 26, and pitch axis 28 remain fixed relative to a satellite body 32 of satellite 20. Satellite body 32 holds a payload (not shown) and other components which are conventional in the art.

FIG. 1 shows satellite 20 in a noon orbit position 34. At noon position 34, the earth 22, satellite 20, and the sun 36 are aligned, with satellite 20 located between the earth 22 and the sun 36. Yaw axis 26 is aligned with a sun line 38. Sun line 38 is defined as a line from the center of the sun to satellite 20. Satellite 20 is also aligned with the sun and the earth when in a midnight position 34', although the sun may not be visible to satellite 20 at position 34' because the earth 22 will be located between satellite 20 and the sun 36.

Satellite 20' is also shown in a six p.m. position 40. At six p.m. position 40, roll axis 24' is parallel with sun line 38 and yaw axis 26' is perpendicular to sun line 38. Satellite 20' is not collinear with sun line 38 because satellite 20' is spaced away from sun line 38 by the radius of the earth plus the satellite's distance from the earth's surface. Satellite 20' is seen from the side, with roll axis 24' pointing from left to right in FIG. 1, tangential to the earth, yaw axis 26' pointing toward the center of the earth (i.e., into the page), and pitch axis 28' perpendicular to both roll axis 24'and yaw axis 26'.

FIGS. 1 and 2 show the sun at various points as it rises and falls relative to a hypothetical plane 42 which is co-planar with the earth's equator. Plane 42 is known as the earth equatorial plane to those skilled in the art. Plane 42 is generally perpendicular to pitch axis 28. The sun's movement relative to plane 42 throughout the year causes the earth's tilt with respect to the sun to change accordingly. The angle between satellite 20 and the sun is known as a sun declination 43, and the change in reference attitude of the sun with respect to satellite 20 is sun declination bias. When the sun is at an equinox position typical of March and September as indicated by position 36, the sun lies relatively close to plane 42 and sun declination of satellite 20 in an equatorial orbit is near zero. When the sun is at a summer solstice position typical of June as indicated by position 36", the sun is high above the earth's equatorial plane, and sun declination is near +23 . When the sun is at a winter solstice position typical of December as indicated by position 36', the sun is far below the earth's equatorial plane, and sun declination is near -23".

FIG. 2 shows satellite 20' orbiting the earth 22 from a second equatorial view. Satellite 20' is shown at six p.m. orbit position 40 from the rear, with roll axis 24' pointing tangential to the earth into the page, yaw axis 26' pointing toward the center of the earth from right to left, and with pitch axis 28' perpendicular to roll axis 24' and yaw axis 26'.

During operational orbit, satellite 20" will be oriented similarly with respect to axes 24', 26', and 28' when in a six am position 40' as well, although it will be on the opposite side of the earth from six p-rn. position 40. Satellite 20" is shown at orbit position 40' in an exemplary acquisition attitude for illustration. During acquisition, the attitude of satellite 20" fluctuates widely and substantially randomly with respect to the sun and the earth, enabling acquisition to occur regardless of the initial orientation of satellite 20" with respect to the sun 36 and the earth 22.

Satellite 20"' is shown at a nine a.m. orbit position 44, which is half-way between six am. orbit position 40' and noon orbit position 34 (FIG. 1). Satellite 20"' at position 44 is at an oblique angle with respect to the page with roll axis 24" along the satellite's orbital path 30, yaw axis 26" pointing at an angle into the page, and pitch axis 28" perpendicular to roll axis 24" and yaw axis 26".

While FIGS. 1 and 2 show satellite 20 at various orbit positions, those skilled in the art will recognize that satellite 20 may be different satellites or the same satellite at different positions in orbit 30. Similarly, those skilled in the art will recognize that the orbit positions discussed (e.g., six p.m., noon) represent places in orbit 30 of satellite 20 relative to the sun, and do not represent the time of day satellite 20 may be in any certain orbit position. For example, it is possible for satellite 20 to be in noon orbit position 34 at 3:35 pJn. in geostationary orbit and at approximate 100-minute intervals in a low earth orbit. Or, satellite 20 may be in a nonequatorial orbit such as a polar orbit.

FIG. 3 shows, from the perspective of the sun, a sun sensor 46 mounted on an articulated solar panel 48, which is mounted to satellite body 32 in accordance with a preferred embodiment of the present invention. Satellite 20 (FIG. 1) may have a second solar panel (not shown) mounted to the opposite side of satellite body 32. The second solar panel is similar to solar panel 48, and may have a sun sensor mounted to it. Because sun sensor 46 has a variable field of view (discussed below), in an altemate embodiment, sun sensor 46 could be mounted on satellite body 32 rather than on solar panel 48.

Solar panel 48 has a solar collector array 52 on one surface. Once deployed, solar panels 48 are roughly parallel to and elongated along the direction of pitch axis 28. Sun sensor 46 is mounted on the surface of solar panels 48 which contains solar collector array 52. Solar panels 48 articulate with respect to satellite body 32 throughout orbit 30 in order to keep solar collector array 52 facing the sun so that they can supply power to systems onboard satellite 20. This articulation maintains solar panels roughly normal to sun line 38 (FIG. 1) regardless of the attitude of satellite 20.

Sun sensor 46 is desirably mounted on a yoke 54 of solar panel 48. Yoke 54 is desirably visible to the sun during the ascent phase when solar panel 48 may not be deployed as well as during acquisition and operational orbit phases. A signal processor unit 56 is also mounted on yoke 54.

Solar panel 48 electrically joins satellite body 32 by means of a slip ring junction 58. Slip ring junction 58 is an electrical connector that provides electrical connection regardless of relative rotary position of slip ring contacts. Surrounding slip ring junction 58 is an articulation point (not shown) that pivots to allow solar panel 48 to deploy and articulate. Slip ring junction 58 is subject to significant noise. This noise would normally distort lower amplitude analog signals passing through slip ring junction 58. Because signal processor unit 56 resides on the same side of slip ring junction 58 as sun sensor 46, analog sun sensor outputs are processed (e.g., converted to digital) before passing through slip ring junction 58. These digitally processed signals are not significantly influenced by the noise that is introduced at slip ring junction 58. Signal processor unit 56 sends these digitally processed signals across slip ring junction 58 to an on-board computer 62 located within satellite body 32.

A controller system 64 includes sun sensor 46, signal processor unit 56, an on-board computer 62, a gyroscope 66, a telemetry, tracking and control antenna 68, and any data signals and signal processing among these elements. Signals may be processed by either signal processor unit 56 or by on-board computer 62. In the preferred embodiment of the present invention, signal processor unit 56 resides on the surface of solar panel 48 opposite solar collector array 52, causing signal processor unit 56 to constantly face away from the sun. This protects signal processor unit 56 from variations in temperature that would result from a varied orientation to the sun as experienced within satellite body 32. This mounting location therefore provides signal processor unit 56 with inherent thermal stability. Accordingly, sun sensor 46 and signal processor unit 56 need not be equipped with a thermal stability system.

On-board computer 62 is located on the satellite body 32 side of slip ring junction 58.

On-board computer 62 couples to signal processor unit 56. Telemetry, tracking and control antenna 68 couples to signal processor unit 56 enabling communication between satellite 20 and ground control. Gyroscope 66 is also located within satellite body 32 and is in data communication with on-board computer 62. Gyroscope 66 is fixed with regard to satellite body 32, and conventionally consists of at least two momentum wheels; one rotating about roll axis 24 and another rotating about yaw axis 26.

FIG. 4 shows a schematic representation of sun sensor 46 mounted on solar panel 48 from the perspective of satellite body 32 (FIG. 3), looking along pitch axis 28 in accordance with a preferred embodiment of the present invention. Sun sensor 46 contains a roll-yaw photoelectric device 70 and a pitch photoelectric device 72. Photoelectric devices 70 and 72 are sensitive to the sun and provide electrical signals to signal processor unit 56 based on their relative orientation with respect to the sun. Pitch photoelectric device 72 is canted at approximately a 45" angle to the width of solar panel 48 as indicated by a hypothetical plane 76.

FIG. 5 shows a schematic representation of sun sensor 46 mounted on solar panel 48 from a side perspective in accordance with a preferred embodiment of the present invention.

Roll-yaw photoelectric device 70 is canted at approximately a 45" angle to the elongated surface of solar panel 48 as indicated by a hypothetical plane 74. Planes 74 and 76 (FIG. 4) are substantially normal to each other, and each is positioned at a 450 angle to solar panel 48.

Planes 74 and 76 are hypothetical and are not a part of the physical structure of sun sensor 46 but are shown for illustration only.

FIG. 6 shows a block diagram of signal processor unit 56 in accordance with a preferred embodiment of the present invention. Signal processor unit 56 is connected to photoelectric devices 70 and 72 and on-hoard computer 62 as shown in FIG. 3. Sun sensor outputs 60 and 60' are analog signals that are received by buffer amplifiers 78 and 78', then sent to summing devices 80 and 80', respectively. Summing devices 80 and 80' also receive offset signals 90 and 90' from a controller 86 via digital to analog converters 88 and 88', respectively. In a preferred embodiment, summing devices 80 and 80' take the buffered sun sensor outputs 60 and 60' adjusted by offset signals 90 and 90', and send the offset-compensated signals to a variable gain differential amplifier 82. In an alternate embodiment, controller 86 could achieve gain control at the inputs to buffer amplifiers 78, 78', rather than at summing devices 80, 80', respectively. This altemate embodiment would add flexibility and keep the inputs from saturating if they were made high gain.

Variable gain differential amplifier 82 amplifies the difference between the offsetcompensated signals, and sends an amplified difference signal to an analog to digital converter 84. Analog to digital converter 84 converts the amplified difference signal to digital, then sends the converted digital signal to controller 86. Controller 86 converts the data into a solar angle 92 (FIG. 7). Controller 64 performs satellite attitude control processing based on solar angle 92. Solar angle 92 reflects three components: the alignment of solar panel 48 (FIG. 3), the attitude of satellite 20 (FIGS. 1 and 2), and sun declination 43 (FIGS. 1 and 2). Based on processing performed on solar angle 92, controller 86 calculates new offset signals 90 and 90', and sends the new offset signals to digital to analog converters 88 and 88', which forward them to summing devices 80 and 80', respectively.

Variable gain differential amplifier 82 operates at variable gains to maximize the amplifier sensitivity at all phases of operation and orientation to the sun. This allows controller 64 to effectively achieve satellite attitude control during ascent, acquisition, and operational phases of satellite launch and orbit using outputs 60 from sun sensor 46. For example, variable gain amplifier 82 desirably operates at a relatively low gain state during ascent and acquisition when a wide field of view is useful to capture the sun. Referring back to FIG. 2, satellite 20" is shown at orbit position 40', which illustrates that satellite attitude may fluctuate significantly during the acquisition phase when satellite 20" is attempting to orient itself correctly with respect to the sun. After acquisition, variable gain amplifier 82 switches to a relatively high gain state under the control of controller 86 where it remains during operational orbit 30.

Signal processor 56 maintains amplifier 82 at the high gain state in order to detect slight motions of the satellite relative to the sun. The detection of such slight motions are used for satellite attitude control. Controller 86 achieves gain control by applying a signal through digital to analog converter 88" to amplifier 82.

Controller 86 varies the gain states of variable gain amplifier 82 between relatively low and high gain states autonomously, desirably based on information from on-board computer 62 describing the current orbital phase. If failure of a system onboard the satellite causes the sun to disappear from view of sun sensor 46 at a time when the satellite is not in eclipse, variable gain amplifier 82 is changed back to a relatively low gain state until the sun is re-acquired.

Variable gain amplifier 82 is then changed back to a relatively high gain state. On-board computer 62 uses an ephemeris table (not shown) to determine whether the satellite should be in eclipse.

In general, solar panel 48 tracks and is substantially normal to the sun. If solar sensor 46 is mounted on solar panel 48, the high gain output of signal processor 56 can be used in conjunction with orbit tables and solar panel resolver outputs to calculate the orientation of satellite 20. Because solar panels typically have limited ranges of motion, there are times in a satellite's orbit when the panels will not be normal to the sun.

FIG. 7 illustrates a solar activity curve 94 which represents the two half-sine sun sensor outputs 60 (FIG. 6) as a function of solar angle 92 calculated in accordance with a preferred embodiment of the present invention (outputs of a second sun sensor are shown inverted).

Referring to both FIGS. 6 and 7, signal processor 56 desirably operates within a substantially linear portion 96 of curve 94, even more desirably around a center point 98 of curve 94 to prevent saturation of variable gain amplifier 82 by small changes in sun sensor outputs 60.

Accordingly, the use of offset signal 90 places center point 98 of linear portion 96 of solar activity curve 94 within a small range of variation of solar angle 92 as adjusted by offset signal 90. The calculation of offset signal 90 along with the gain of variable amplifier 82 selects the range and slope of this linear region. Effectively, the use of offset signal 90 expands the center of linear portion 96 of solar activity curve 94, enabling controller system 64 (FIG. 3) to respond to very small changes in solar angle 92. This linearization by use of offsets can eliminate the need for complex processing of sun sensor outputs 60 when sun declination 43 (FIG. 1) is 22.50 or less by reducing the risk of saturation of variable gain amplifier 82 even when variable gain amplifier 82 is operating in the relatively high gain state.

Offset signal 90 is calculated by controller 86 based on sun declination 43 (FIG. 1).

Desirably, a periodic re-calculation of offset signal 90 maintains center point 98 of linear portion 96 of solar activity curve 94 within a small range of solar angle 92, as described above, preventing solar angle 92 from exceeding the maximum capacity of signal processor 56 and saturating variable gain amplifier 82. This ensures that signal processor 56 will be capable of determining the value of solar angle 92. Thus, use of a constantly adjusted offset signal 90 allows variable gain amplifier 82 to operate at the relatively high gain state.

Throughout orbit 30 (FIGS. 1 and 2), solar panels 48 articulate with respect to satellite body 32, causing the orientation of photoelectric devices 70 and 72 (FIGS. 4-6) with respect to roll axis 24, yaw axis 26, and pitch axis 28 to change. Since photoelectric devices 70 and 72 lie in two normal planes 74 and 76 (FIGS. 4 and 5), sun sensor 46 (FIGS. 4-6) receives information with respect to at least two axes at any given point in orbit 30 in which the sun is visible. This makes it possible to obtain detailed three-axis attitude information over the course of an entire orbit period as described below.

FIG. 8 shows the orbit positions near which various measurements are taken with respect to satellite roll axis 24, yaw axis 26, and pitch axis 28 from the perspective of the North Pole 100. As satellite 20 orbits the earth 22, one of photoelectric devices 70 and 72 (FIGS. 4-6) rotates towards the sun causing its outputs 60 (FIG. 6) to increase, while the other of photoelectric devices 70 and 72 rotates away from the sun causing its outputs 60 to decrease.

As described above with respect to FIGS. 1-3, the articulation of solar panels 48 throughout orbit 30 combined with the arrangement of sun sensor 46 as shown in FIGS. 3-5 gives rise to four notable orbit positions. At these notable positions, the outputs of photoelectric devices 70 and 72 (FIGS. 4-6) change most noticeably in response to slight satellite attitude changes about one of roll axis 24 and yaw axis 26 while they do not change in response to satellite attitude changes about the other of roll axis 24 and yaw axis 26. These notable orbit positions and the intervals surrounding them are discussed below, followed by a discussion of the types of processing desirably performed on sun sensor outputs 60 during the surrounding intervals.

At noon orbit position 34 and midnight orbit position 34', yaw axis 26 is collinear with sun line 38, making roll-yaw photoelectric device 70 insensitive to yaw attitude error. In other words, sun sensor outputs 60 (FIG. 6) do not change significantly as satellite 20 rotates about yaw axis 26. However, sun sensor outputs 60 change significantly as satellite 20 rotates about roll axis 24 at noon orbit position 34. Thus, sun sensor 46 (specifically, roll-yaw photoelectric device 70) is highly sensitive to roll attitude error in orbit position 34. This is due to the geometry of sun sensor 46 (FIGS. 3-5) which causes roll-yaw photoelectric device 70 to be canted at an angle to pitch axis 28 combined with the fact that roll axis 24 is perpendicular to sun line 38 at orbit position 34. Since pitch photoelectric device 72 is canted at an angle to yaw axis 26 at this point in orbit 30 (FIG. 4), sun sensor 46 (specifically, pitch photoelectric device 72) will be relatively sensitive to changes in yaw at orbit position 34. This is especially true when sun declination is high, placing the sun at position 36" (FIGS. 1 and 2) typical of June.

In other words, during the interval approaching orbit position 34, the orientation of solar panel 48 with respect to satellite body 32 places photoelectric devices 70 and 72 at roughly 45" angles to roll axis 24. Outputs 60 from roll-yaw photoelectric device 70 increase while outputs 60' from pitch photoelectric device 72 decrease. At the interval following orbit position 34, outputs 60 from roll-yaw photoelectric device 70 decrease while outputs 60' from pitch photoelectric device 72 increase. Overall, this causes the changes in outputs 60 from each photoelectric device to be additive, increasing the sensitivity of sun sensor 46 as a whole with respect to roll axis 24.

At six p.m. orbit position 40 and six am. orbit position 40', yaw axis 26 is perpendicular to sun line 38, making roll-yaw photoelectric device 70 relatively sensitive to yaw attitude error in this orbit position. In other words, sun sensor outputs 60 change significantly as attitude of satellite 20 rotates about yaw axis 26. However, sun sensor outputs 60 do not change significantly as attitude of satellite 20 rotates about roll axis 24. Thus, sun sensor 46 (specifically, roll-yaw photoelectric device 70) will be relatively insensitive to roll attitude error in orbit positions 40 and 40'. This is due to the geometry of sun sensor 46 (FIGS. 3-5) which causes roll-yaw photoelectric device 70 to be canted at an angle to pitch axis 28 combined with the fact that roll axis 24 is roughly parallel to sun line 38 and rotation about roll axis 24 does not yield significant rotational change between photoelectric devices 70 or 72 and sun line 38.

In other words, at the intervals approaching orbit positions 40 and 40', the orientation of solar panel 48 with respect to satellite body 32 places photoelectric devices 70 and 72 at roughly 45" angles to yaw axis 26. Outputs 60 (FIGS. 6 and 7) from roll-yaw photoelectric device 70 decrease while outputs 60' from pitch photoelectric device 72 increase. At the interval following orbit position 34, outputs 60 from roll-yaw photoelectric device 70 increase while outputs 60' from pitch photoelectric device 72 decrease. Overall, this causes the changes in outputs 60 from each photoelectric device to be additive, increasing the sensitivity of sun sensor 46 as a whole with respect to yaw axis 28.

In addition, since pitch photoelectric device 72 is canted at an angle to yaw axis 26 in orbit positions 40 and 40' (FIG. 4), it will be relatively sensitive to changes in attitude about pitch axis 28.

Sun declination bias is predicted independently with respect to roll axis 24 and yaw axis 26, desirably when satellite 20 is at orbit positions where sun sensor outputs 60 do not vary in response to changes in attitude about the subject axes. This allows sun declination bias to be relatively free from any influence of satellite attitude error about the subject axes.

Accordingly, sun declination 43 (FIG. 1) with respect to roll axis 24 is measured near one or both of six p.m. orbit position 40 and six am. orbit position 40', when photoelectric device 70 is insensitive to attitude error about roll axis 24. This desirably takes place during one of a relatively small sun declination roll interval 102 and 102', respectively. Similarly, sun declination 43 with respect to yaw axis 26 is measured near noon position 34, when photoelectric device 70 is insensitive to attitude error about yaw axis 26 during a relatively small sun declination yaw interval 104.

Desirably, estimations for attitude error determination and gyroscope 66 (FIG. 6) calibration are also taken with respect to each axis independently. Preferably, these estimations are performed when satellite 20 is at orbit positions where sun sensor outputs 60 vary most significantly in response to the slightest changes in attitude about the subject axis, resulting in more precise estimates.

Accordingly, desired positions for estimating attitude of satellite 20 and for calibrating gyroscopes 66 with respect to roll axis 24 surround noon orbit position 34 during a gyroscope roll calibration interval 106 beginning near nine am. orbit position 44 and ending near a three p.m. orbit position 108. Similarly, desired positions for estimating the attitude of satellite 20 and calibrating gyroscopes 66 with respect to yaw axis 26 surround six p.m. and six a.m. or panel yaw alignment interval 116'. Likewise, desired positions for measuring alignment error of solar panel 48 with respect to pitch axis 28 of satellite body 32 surround six p.m. orbit position 40 during a 15-minute panel pitch alignment interval 118 and six am. orbit position 40' during a 15-minute panel pitch alignment interval 118'.

FIG. 9 illustrates a process 120 by which satellite attitude is measured and adjusted, gyroscopes calibrated, and sun declination predicted for roll axis 24 in accordance with a preferred embodiment of the present invention.

With respect to FIGS. 8 and 9, process 120 includes a query task 122 which determines whether satellite 20 is in gyroscope roll calibration interval 106. Query task 122 desirably determines orbit position of satellite 20 by consulting an ephemeris table (not shown). If query task 122 determines that satellite 20 is in gyroscope roll calibration interval 106, processing continues with a task 136 which is described below.

If query task 122 determines that satellite 20 is not in gyroscope roll calibration interval 106, then a query task 126 determines whether satellite 20 is in sun declination roll interval 102 or 102'. Query task 126 desirably determines the orbit position of satellite 20 by consulting the ephemeris table. If query task 126 determines that satellite 20 is not in sun declination roll interval 102 or 102', process flow loops back to task 122.

If query task 126 determines that satellite 20 is in sun declination roll interval 102 or 102', a task 128 converted sun sensor outputs 60 into the roll component of solar angle 92.

With reference to FIGS. 3-5 and 7, the combined geometries of the orbit and sun sensor 46 place photoelectric devices 70 and 72 at an angle such that they yield data primarily relevant to roll axis 24 of satellite 20. With reference to FIG. 6, analog to digital converter 84 converts the analog signal to digital for processing by controller 86. Task 128 may measure a difference in solar angle 92 from a previously determined solar angle 92 due to the offsets discussed above.

With reference to FIGS. 6 and 9, a task 130 may consider offset signal 90 in order to obtain solar angle 92. Solar angle 92 is an absolute figure which represents the change in sun declination since the previous estimate of sun declination. Prior sun declination is reflected in offset signal 90. With reference to FIG. 7, the use of offset signal 90 desirably maintains center point 98 of linear portion 96 of solar activity curve 94 within a small window of solar angle 92, preventing solar angle 92 from exceeding the maximum capacity of signal processor 56.

Task 130 then feeds sun declination angle 92 to a solar model 132, which models sun declination 43 (FIG. 1). Solar model 132 is described below. By comparing successive past sun declination bias measurements, solar model 132 predicts future sun declination 43 by extrapolation or other simulation or prediction techniques known to those skilled in the art.

Next, a task 134 optionally adjusts offset signal 90 in response to the newly predicted sun declination 43. Optionally, offset signal 90 could alternatively be adjusted at a point in a schedule other than when sun declination 43 roll is measured. For instance, sun declination 43 yaw and pitch attitude could be considered in calculating an appropriate offset signal 90 to apply. As discussed above with reference to FIG. 7, proper maintenance of offset signal 90 is desirable to maintain the small solar window tolerance desired to prevent saturation of variable gain differential amplifier 82 while operating at the relatively high gain state. By adjusting offset signal 90 in response to newly measured sun declination, very slight changes in sun declination bias can be detected due to the high gain operation of variable gain amplifier 82.

Since sun declination 43 is a factor in determining panel alignment and attitude, the maintenance of offset signal 90 in response to sun declination 43 enhances resolution of all readings, thereby making it possible to achieve effective attitude control. After completion of task 134, process flow loops back to task 122.

If query task 122 determines that satellite 20 is within gyroscope roll calibration interval 106, task 136 measures satellite attitude error with respect to roll axis 24. Satellite roll attitude error represents the attitude deviation of satellite 20 with respect to roll axis 24. Roll attitude error is affected by sun declination bias and panel alignment, which is discussed below with reference to a task 146.

A task 138 then adjusts the measurement obtained above in task 136 for predicted sun declination 43, obtained from solar model 132, and for predicted solar panel alignment, obtained from a solar panel alignment model 148, and feeds adjusted measurements to a satellite attitude model 140. Satellite attitude model 140 is described below and stores satellite attitude roll error measurements obtained during successive orbits. In response to prior satellite attitude roll error measurements, future satellite attitude roll error is predicted by extrapolation or other simulation or prediction techniques known to those skilled in the art.

A task 142 then uses the newly measured attitude roll error, now compensated for sun declination bias and solar panel alignment, to calibrate gyroscope 66 (FIG. 6) with respect to roll axis 24. This allows gyroscope calibrations to be performed continuously throughout gyroscope roll calibration interval 106 independently for roll axis 24 based on frequently obtained, relatively accurate sun sensor outputs 60 (FIG. 6).

After task 142, a query task 144 determines whether satellite 20 is in a panel roll alignment interval 114 or 114' (FIG. 8). If query task 144 determines that satellite 20 is in panel roll alignment interval 114 or 114', task 146 uses the adjusted solar angle measurement from task 138 to determine an alignment error of solar panel 48 with respect to roll axis 24 of satellite body 32. Alignment error represents the panel alignment deviation from predicted panel alignment with respect to roll axis 24 of satellite body 32.

Task 146 compares the measurement from task 138 determined above with a previous measurement from task 138. The previous measurement was taken during a previous iteration of task 146 when satellite 20 was half an orbit away from its current orbit position. The two readings are similar to each other, but components of the two readings due to panel alignment error but not predicted by panel simulation are out of phase with respect to each other because satellite 20 was at opposite orientations to the sun when the two readings were taken. By taking the difference of the two readings, a result doubles any panel misalignment not previously accounted for by panel simulation in solar panel alignment model 148 while canceling alignment and declination components of the angle measurements. This result is fed to solar panel alignment model 148, which stores panel alignment error measurements obtained during successive orbits and predicts future panel alignment. In response to past panel alignment error measurements, solar panel alignment model 148 predicts future panel alignment by extrapolation or other simulation or prediction techniques known to those skilled in the art.

If query task 144 determines that satellite 20 is not in panel roll alignment interval 114 or 114', a query task 150 determines whether satellite 20 is in eclipse based on information in the ephemeris table.

If query task 150 determined that satellite 20 is in eclipse, a task 152 substitutes gyroscope outputs for measured attitude. Conventional conversion techniques are used to convert gyroscope outputs from rate information to obtain attitude measurements. A task 154 compares measured attitude with respect to roll axis 24 with predicted attitude stored in satellite attitude model 140. Since gyroscope calibration is desirably performed by task 142 throughout gyroscope roll calibration interval 106, gyroscope 66 (FIG. 6) will have been freshly calibrated causing gyroscope outputs to be relatively accurate.

After task 154, a task 156 evaluates error determined in task 154 to determine whether to alter orientation of satellite 20. This determination may be based on the degree of attitude error and future scheduled changes in attitude. Conventional methods of satellite attitude adjustment may be employed, including the firing of a maneuvering thruster.

While FIG. 9 illustrates a process 120 for achieving satellite attitude control with respect to roll axis 24, those skilled in the art will recognize that process 120 can be readily adapted to yield similar processes, discussed below, for achieving satellite attitude control with respect to yaw axis 26 and pitch axis 28.

A process (not shown) similar to process 120 achieves satellite attitude control with respect to yaw axis 26. Yaw satellite attitude control differs as follows: Query task 122 determines whether satellite 20 is in yaw gyroscope calibration interval 110 or 110'. Query task 126 determines whether satellite 20 is in sun declination yaw interval 104 rather than sun declination roll interval 102 or 102'. Query task 144 determines whether satellite 20 is in panel yaw alignment interval 116 or 116' rather than in panel roll alignment interval 114 or 114'.

Also with respect to yaw attitude control, tasks 128, 130, and 134 are concerned with the yaw component of sun declination bias and offset signal 90 rather than the roll component.

Similarly, tasks 136, 138, 142, and 152 are concerned with the yaw component of satellite attitude, attitude error, and gyroscope calibration rather than the roll component. Task 146 is concerned with the yaw component of solar panel alignment and alignment error rather than the roll components A process (not shown) similar to process 120 achieves satellite attitude control with respect to pitch axis 28. Pitch satellite attitude control differs as follows: Since there is no sun declination component with respect to pitch axis 28, query task 126 and tasks 128; 130, and 134 may be eliminated for pitch axis 28. Query task 122 determines whether satellite 20 is in pitch gyroscope calibration interval 112 or 112' rather than roll gyroscope calibration interval 106. Query task 144 determines whether satellite 20 is in panel pitch alignment interval 118 or 118'. Tasks 136, 138, 142, 146, and 152 are concerned with the pitch component of satellite attitude, attitude error, solar panel alignment, and gyroscope calibration rather than the roll component.

Those skilled in the art will recognize that models 132, 140, and 148 discussed above may be separate or integrated. FIG. 9 and the above discussion describe separate models 132, 140, and 148 for convenience. However, the preferred embodiment of the present invention uses an 18th order Kalman filter to model sun declination, satellite attitude, and panel alignment in an integrated fashion. While Kalman filtering is well-known, those skilled in the art will recognize that other methods of simulating, estimating, or otherwise predicting sun declination, satellite attitude, and solar panel alignment could be used as well.

Desirably, however, two of the three known model components described above are used to derive a new value for the third model component at intervals or positions in orbit when the third model component can be most accurately determined. For example, sun declination bias and satellite attitude are used to determine solar panel alignment. The process described above updates each model component in turn using newly derived data for that model component combined with previously determined values for the other two model components. Thus over an entire orbit, each newly obtained model component has been used to determine the other two model components for each axis independently of the other axes. Over successive orbits, a progressively longer string of data builds up which can be used to make progressively more accurate estimations. This approach to modeling allows a satellite's orbit attitude to be finetuned within three to seven days.

In summary, the present invention may replace an array of precise and imprecise earth sensors, precise sun sensors, precise gyroscopes, and related support systems with additional processing on the outputs provided by simple sun sensors. Allowed to operate at high gain, simple sun sensors can achieve high resolution outputs during operational orbit, replacing the need for several digital body-mounted sun sensors and supporting switching and thermal stabilization equipment. This is done without sacrificing sun sensor use during ascent and acquisition. By measuring sun declination at orbit regions uninfluenced by attitude error, highly detailed sun declination readings can be achieved Processing measures attitude error at points in orbit when the simple sun sensors are most sensitive to it. Because these precise attitude readings can be taken at many points in an orbit, they allow more frequent calibration of satellite gyroscopes, allowing substitution of lower precision and more simple gyroscopes.

Although the preferred embodiments of the invention have been illustrated and described in detail, it will be readily apparent to those skilled in the art that various modifications may be made therein without departing from the spirit of the invention or from the scope of the appended claims. For example, although the preferred embodiment has been described in terms of a sun sensor, in an alternate embodiment, the method and apparatus of the present invention could also use the moon as a reference point when the dynamic range of the sensors and the gain of the amplifiers is sufficient. Such a moon sensor could be used, for example, when the satellite is in the shadow of the earth.

Claims (10)

CLAIMS What is claimed is:

1. A satellite attitude controlling apparatus which controls satellite attitude in response to a relative orientation between a satellite and the sun, said apparatus comprising: a sun sensor mounted on said satellite so that said sun sensor is visible to the sun during acquisition and orbit phrases; and a signal processor unit coupled to said sun sensor, said signal processor unit including a controller coupled to a variable gain amplifier, said signal processor unit being configured to alter gain of said variable gain amplifier in response to orbital phase.

2. A satellite attitude controlling apparatus as claimed in claim 1 wherein: said satellite has a solar panel articulated to a satellite body; and said signal processor unit and said sun sensor are mounted on said solar panel.

3. A satellite attitude controlling apparatus as claimed in claim 1 wherein said controller is configured to predict sun declination bias using outputs intermittently obtained from said variable gain amplifier, and adjust said sun sensor outputs in response to said predicted sun declination bias to determine attitude of said satellite.

4. A satellite attitude controlling apparatus as claimed in claim 1 wherein: said apparatus additionally comprises a moon sensor mounted on said satellite so that said moon sensor is visible to the moon during eclipse phases.

5. A method for controlling attitude of a satellite using a sun sensor, said method comprising the steps of: (a) driving a solar model which predicts sun declination bias using outputs intermittently obtained from said sun sensor; (b) obtaining outputs from said sun sensor; and (c) adjusting said sun sensor outputs in response to said predicted sun declination bias to determine attitude of said satellite.

6. A method as claimed in claim 5 wherein said driving step (a) comprises the steps of: (d) measuring said sun declination bias with respect to a roll axis of said satellite at a first region in an orbit of said satellite; and (e) measuring said sun declination bias with respect to a yaw axis of said satellite at a second region in said orbit of said satellite.

7. A method as claimed in claim 5 wherein said sun sensor is mounted on an articulated solar panel which resides at an angle to a body of said satellite, and said method additionally comprises the step of: (d) adjusting said sun sensor outputs in response to alignment of said solar panel with respect to said satellite body.

8. A method as claimed in claim 5 additionally comprising the step of: (d) altering orientation of said satellite in response to said attitude determined in said adjusting step (c).

9. A method as claimed in claim 5 additionally comprising the step of: (d) predicting satellite attitude in response to outputs obtained from said sun sensor.

10. A method for controlling satellite attitude based on outputs derived from a sun sensor mounted on a solar panel of a satellite, said method comprising the steps of: coupling a signal processor unit to said sun sensor, said signal processor unit including a variable gain amplifier and a controller; operating said variable gain amplifier in a relatively low gain during an acquisition phase to acquire the sun; switching said variable gain amplifier to a relatively high gain once the sun is acquired; processing said sun sensor outputs at a satellite noon orbital position to determine sun declination bias with respect to a yaw axis; processing said sun sensor outputs at one of a satellite six p.m. orbital position and a satellite six am. orbital position to determine said sun declination bias with respect to a roll axis; processing said sun sensor outputs at one of said six p.m. position and said six a m.

position to determine a satellite attitude with respect to said yaw axis and a pitch axis; processing said sun sensor outputs at one of said noon orbital position and a midnight orbital position to determine said satellite attitude with respect to said roll axis; processing said sun sensor outputs at one of said noon and midnight orbital positions and to calibrate gyroscopes with respect to said roll axis; processing said sun sensor outputs at one of said six p.m. and six a.m. orbital positions and to calibrate gyroscopes with respect to said yaw axis; estimating alignment of said solar panel in response to said sun sensor outputs; predicting said sun declination bias in response to said sun sensor outputs; and estimating said satellite attitude error in response to said sun sensor outputs.