Attitude Determination and Control System
Altitude determination and control is responsible for continuously tracking the orbit, the spacecraft’s orientation and if required modify either to make sure other subsystems and the payload points in the required direction.
| Sensor | How it works | Adventages | Disadventages | Rough Accuracy (arcmin) |
|---|---|---|---|---|
| Magnetometer | Measures the local magnetic field and compares with on- board magnetic field model to determine spacecraft attitude. |
-Available as IC: –Low mass –Small size –Low power |
- Accuracy dependent on that of gravitational field models held on-board if calculations are performed on-board -Requires knowledge of spacecraft position at all times |
30 |
| Sun Sensors | Detect incident light levels on the spacecraft. Allows satellite-sun vector to be determined. |
-Can be very simple -Accuracy-size trade- off can be adjusted as required -Solar panels can be used as basic sun sensors |
-Doesn’t work during eclipse -Far all time operation requires sensors on each side (360 degrees FOV) |
1 |
| Earth Horizon Sensors | Detect boundary between Earth’s IR and space. Uses this to calculate the nadir vector direction (spacecraft- to-geocenter). |
-Constant observation for Earth pointing spacecraft |
-Fuzzy boundary -Heat detectors are required -Cannot detect errors around the local vertical -Limited by horizon definition and degrades at lower altitudes |
5 |
| Star Sensors | Determinates attitude of spacecraft compared to known star patterns in field of view. |
-One sensor is sufficent to provide reference vector |
-High mass, power and processing requrement -May be confused be other light sources such as satellites |
1/60 |
| Inertial Measurement Unit |
Consists of accelerometers and gyroscopes. Plots the rotation rates and calculates attitude over time. |
-Available as IC: –Low mass, power and small size -Provides attitude information in between readings from other sensors |
-not a reference sensor, inertia only -requires other sensors to acquire and update attitude |
N/A |
table 1 – Sensors
| Sensor | How it Works | Adventages | Disadventages | |
|---|---|---|---|---|
| Magnetorquer | Create a magnetic field on satellite which attempts to align itself with the local magnetic field of the Earth |
-Torque about any axes -Virtually unlimited lifetime |
-Possible magnetic interference between systems -Accuracy dependant on accuracy of on- board magnetic field model or magnetometer |
|
| Gravity Gradient | Uses Earth’s gravity field to stabelise the satellite. |
-Passive system | -Significant mass -Low accuracy -Requires deployable structure -Contol only along z-axes -Boom must always be perpendicular to Earth’s surface therefore limited manoeuvrability |
|
| Spin Stabelisation | Spun spacecraft are stable about the axis with the largest momentum of inertia. Typically 10 to 60 rpm. |
-Passive system | -Must be used in parallel with other ACS to start spin -Spinning may effect payload and sensors |
|
| Passive Magnets | Magnets align themselves with the Earth’s magnetic field such that the spacecraft rotates twice each orbit |
-Passive system | -Heavy -Only aligned along one axes One axes attitude is predetermined therefore limited manoeuvrability |
|
| Thrusters | Direct application of external torques. |
-Torque about any axes -Suitable for any torque size |
-Heavy -Limited amount of fuel |
|
| Adjustable Spacecraft Geometry |
Spacecraft geometry adjusted to utilise naturally occuring external tourques (i.e.: aerodynamic drag or solar radiation pressure) |
-Potentially could move the spacecraft in any direction -Only deployment requires power |
-Complicated deployable structures | |
| Mass Movement | Moving mass inside the satellite such that spacecraft attitude must change to conserve angular momentum |
-Little additional mass | -Complicated mechanism -Large space requirement |
|
| Reaction Wheels | Wheels used to rotate satellite by introducing momentum |
-Relatively rapid and accurate response | -Moving parts may cause complications -Requires additional system for momentum damping |
|
| Momentum Wheel | Wheels spin constantly to control flow of momentum in spacecraft by storing or releasing momentum |
-Relatively rapid and accurate response -Momentum bias increases stability of satellite |
-Requires constant power -Moving parts may cause complications -Requires additional system for momentum damping |
|
| Control Momentum Gyros | Run at constant speed and may have one or two gimbals, absorbing angular momentum by gimbal rotation. |
-Large torque capability -Relatively low weight, pover and size compared to reaction wheel |
-relatively complex -Complicated control algorithm |
|
| Dual-Spin | Similar principle to pure spin satellite but allows part of the spacecraft to be de-spun |
-No duel required | -No existing structure design which would allow it -Complex |
table 2 – Stabelisation
Any variation these sensors can be used however at least two reference sensons are required to determine the attitude of the spacecraft. Additionally one external torque actuator is required in each axes to stabelise the spacecraft. This can be as simple as a graviti gradient boom or as complex as a four wheel redundant reaction wheel system.
| CASE STUDY 2010 University of Southampton. SUSat’s mission objective was to look at the Moon several times each orbit. For this the ADCS have to be more advanced than usual. SUSat uses magnetorquers and reaction wheels for actuators, magnetorquer, gyroscopes and sun-sensors for sensors. To find out more visit the SUSat site. |
OPEN DESIGNSSUSat ADCS, 2010 University of Southampton downloads:
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