Spacecraft Attitude Reacquisition Using Sun Sensor Alignment
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Solution Overview
Problem
Existing methods for reacquiring a three-axis attitude in spacecraft are time-consuming, costly, and require extensive operator intervention, often relying on expensive hardware and complex algorithms, while failing to efficiently align the spacecraft with both the Earth and the sun.
Innovation Solution
A method utilizing existing spacecraft resources, including control electronics, rate and attitude sensors, and torque generators, uses orbital ephemeris data and epoch time to autonomously or semi-autonomously align the spacecraft's attitude with the sun and Earth by determining the desired sun direction and rotating the spacecraft to coincide with it, then aligning with the Earth using a line between the sun and the spacecraft.
Engineering Contradictions & Design Principles
Engineering Contradiction Analysis
1Reliability
If traditional methods are used to reacquire Earth-pointing attitude, then the spacecraft can be realigned to the desired attitude, but the process is time-consuming and requires extensive operator intervention
Solution Approach 1:
The method performs preliminary alignment of the spacecraft roll axis with the sun line before executing the main attitude reacquisition sequence. This preliminary action simplifies subsequent maneuvers and reduces the total time required for full attitude recovery, as the spacecraft is already partially oriented correctly when the main reacquisition algorithm begins execution
Solution Approach 2:
The attitude reacquisition system uses the spacecraft's own sun sensor and onboard computer to autonomously determine the sun's position and calculate the necessary maneuver commands. The system serves itself by using its existing sensors and computational resources rather than requiring external ground control intervention, significantly reducing reacquisition time and operator burden
2Reliability
If multiple discrete maneuvers are performed to align the spacecraft, then the desired attitude is achieved, but the process requires more time and attitude control propellant
Solution Approach 1:
The method combines multiple alignment objectives into a single integrated maneuver sequence. Instead of executing separate maneuvers for roll alignment, pitch alignment, and yaw alignment as independent steps, the algorithm computes a unified set of quaternion commands that achieve all three alignments simultaneously, reducing both the number of discrete maneuvers and the associated propellant consumption
Solution Approach 2:
The algorithm rapidly transitions the spacecraft through intermediate attitudes by computing optimal quaternion-based maneuver profiles that minimize time and propellant expenditure. Rather than making slow, incremental adjustments through multiple small maneuvers, the system rushes through the attitude transition space using mathematically optimized paths that achieve the desired orientation with minimal propellant loss
3Measurement precision
If expensive hardware such as star trackers is used for attitude determination, then precise inertial attitude can be achieved, but the cost and complexity increase significantly
Solution Approach 1:
The method enables the spacecraft's sun sensor to perform multiple functions: it determines both the sun's position for power generation alignment and simultaneously provides the reference information needed for complete three-axis attitude reacquisition. This multi-functional use of the sun sensor eliminates the need for dedicated expensive instruments like star trackers, reducing hardware complexity while maintaining sufficient precision for attitude determination
Solution Approach 2:
The algorithm creates a computational model (quaternion representation) of the desired attitude state based on sun sensor measurements and orbital ephemeris data. This mathematical copy of the attitude information allows the spacecraft to determine its orientation and compute correction maneuvers without requiring physical star tracker hardware, achieving the necessary measurement precision through software-based attitude estimation
4Reliability
If conventional reacquisition methods are used, then Earth-pointing attitude can be restored, but extensive ground controller intervention is required
Solution Approach 1:
The attitude reacquisition system uses the spacecraft's own sun sensor and onboard computer to autonomously determine the sun's position and calculate the necessary maneuver commands. The system serves itself by using its existing sensors and computational resources rather than requiring external ground control intervention, significantly reducing reacquisition time and operator burden
Solution Approach 2:
The method implements a feedback loop where the spacecraft continuously monitors its attitude using rate sensors and sun sensor measurements, compares the current state with the desired attitude, and automatically adjusts the maneuver commands accordingly. This closed-loop feedback system enables autonomous operation by allowing the spacecraft to self-correct without ground controller intervention
Data Source
AI summary
Methods and apparatus for a spacecraft (1) orbiting about a celestial body such as the Earth to reacquire operational three-axis orientation with respect to that body. A method embodiment of the invention comprises determining (201) a set of actual conditions of the spacecraft, comprising a position of the spacecraft (1) in inertial space as a function of time and a set of angular rotation rates of the spacecraft (1) with respect to a coordinate frame of the spacecraft (1), determining (202) an actual instantaneous direction of the sun with respect to the coordinate frame, and propagating (240) an estimated actual sun direction with respect to the coordinate frame as a function of time; determining (260) a desired sun direction with respect to the coordinate frame as a function of time; rotating (270) the spacecraft (1) and adjusting angular rotation rates of the spacecraft (1) so that an actual angle between the spacecraft (1) coordinate frame and the sun as a function of time substantially coincides with the desired sun direction with respect to the coordinate frame as a function of time; and, rotating (280) the spacecraft (1) around an axis defined by a line between the sun and the spacecraft until (1) the celestial body is observed by a celestial body sensor of the spacecraft.