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Attitude Control

The satellite is 3-axis stabilized by actuators, which consist of reaction wheels and magnetic torquers.

The four reaction wheels are aligned in a tetrahedron configuration to ensure redundancy. This makes it possible to compensate for the loss of one of the four reaction wheels. The magnetic torquers (torque rods) are mainly used to dump the momentum accumulated by the reaction wheels over time. They are also used during the LEOP phase (Launch and early operation phase) to stabilize the satellite after the separation from the launch vehicle, because all higher systems (e.g. reaction wheels) are turned off during that phase.

The attitude motion is monitored by five different types of sensors:

  • two 3-axis magnetometers
  • a 4π coarse sun sensors system
  • four Fiber Optic Gyros as rate sensors
  • a star tracker unit with two camera heads
  • three GPS receivers 

The Zarm AMR-magnetometer uses a magneto-resistive sensor and has a digital interface. For the measurement of the angular velocity, four fiber optical rate sensors will be used. A star tracker, the micro Advanced Stellar Compass (µASC), from the Technical University of Denmark will provide a relative pointing knowledge of up to 2 arcseconds. When the satellite is stabilized and rotates with a slew rate of less than 1.2 °/s the star tracker delivers regular attitude updates. To provide full accuracy about all axes and to decrease the probability of blinding during maneuvers, a second camera head unit is mounted on the satellite with its optical axis tilted away from the first one. To support accurate target-pointing of the spacecraft during imaging and ground station contacts, the satellite will be equipped with a GPS navigation system consisting of three Phoenix receivers from DLR/GSOC.

The attitude control system can command a total of five attitude control modes, each specifically designed for a certain operational mode or emergency condition.


  • The Detumbling Mode (Mode 0) is used after launcher separation or if the satellites body rates exceed 3 °/s and is initiated automatically. The Flying Laptop measures it's rotational velocity with the magnetometers and uses the magnetic torquers to slow the rotation down.
  • The Safe Mode (Mode 1) is initiated by the ground station or the automatic fault detection system in case of a mission critical error or failure. To stabilize the satellite and to ensure the solar panels are pointing towards the sun and the battery gets charged, the controller uses the sun sensors to orient the satellite's principle axis towards the sun and the magnetometers and magnetic torquers are used to initiate a 2 °/s spin about this axis. This ensures a stable attitude which is also robust against disturbances.
  • If the satellite is fully operational but is unused it is commanded to the Idle Mode (Mode 3). During the Idle Mode the solar panels are actively pointed to the sun unsing the reaction wheels, the fiber optical gyros, the sun sensors and the star trackers. The batteries will be charged most efficiently that way and the satellite is ready to carry out observations in the pointing modes immediately.
  • For image acquisition three different attitude control modes are defined and shown above: Inertial Pointing Mode (Mode 3), Nadir Pointing Mode (Mode 4) and Target Pointing Mode (Mode 5).
    In the target-pointing, also known as spotlight mode, the satellite points to a fixed spot on the surface of the earth during a fly-over. The slew rate for this maneuver is 1 °/s (max.) and follows a non-linear bell-shaped curve over time. This is the most demanding mode of the satellite in terms of control algorithms and accuracy.
    In the Nadir Pointing Mode the payload cameras are pointed "directly down" (towards nadir) and in the
    Inertial Pointing Mode the Cameras (or any side of the satellite) can be pointed to a celestial object like stars, the sun or the moon. I this mode the Flying Laptop is not rotating, it is inertially fixed in its attitude.
    The startrackers, the fiber optical gyros, the GPS and the reaction wheels are used in the pointing modes. Optionally the magnetometers and the torquers can be switched on to desaturate the reaction wheels.

The attitude control system needs to provide the selected earth observation instruments with a high pointing knowledge of 2.5 arcseconds and a pointing accuracy of 150 arcseconds as well as agile maneuvering capabilities which is a big challenge for a micro-satellite. This can only be achieved by a thorough control concept and high performance sensors/actuators.
New methods in implementing the attitude control algorithms are presently being pursued.