Methods for attitude control of a satellite in survival mode without a priori knowledge of the local time of the satellite's orbit

Patent metadata

Field	Value
Publication number	US-11597539-B2
Application number	US-202017440559-A
Country	US
Kind code	B2
Filing date	Mar 16, 2020
Priority date	Mar 20, 2019
Publication date	Mar 7, 2023
Grant date	Mar 7, 2023

Abstract

A method for attitude control of a satellite in inclined low orbit in survival mode is disclosed, the satellite including at least one solar generator, at least one solar sensor, magnetic torquers capable of forming internal magnetic moments in a satellite reference frame having three orthogonal axes X, Y, and Z, and inertial actuators capable of forming internal angular momentums in the satellite reference frame. The at least one solar sensor has a field of view at least 180° wide within the XZ plane around the Z axis, the method including a step of attitude control using a first control law, a step of searching for the sun by means of the at least one solar sensor, when a first phase of visibility of the sun is detected, and a step of attitude control using a second control law.

First claim

The invention claimed is:

1. A method for attitude control of a satellite in inclined low orbit in survival mode, the satellite comprising at least one solar generator, at least one solar sensor, magnetic torquers capable of forming internal magnetic moments in a satellite reference frame comprising three orthogonal axes X, Y, and Z, and inertial actuators capable of forming internal angular momentums in said satellite reference frame, wherein the at least one solar sensor has a field of view at least 180° wide in a XZ plane around the Z axis, the at least one solar generator is stationary in the satellite reference frame during survival mode and directed so as to generate electrical energy when the sun (S) is located along the Z axis within a field of view of the at least one solar sensor, wherein the method comprises:
a step of attitude control using a first control law according to which the magnetic torquers are controlled to form torques along the X, Y, and Z axes in order to limit the variations of the Earth's magnetic field in the satellite reference frame, and the inertial actuators are controlled to form an internal angular momentum along the X axis,
a step of searching for the sun by means of the at least one solar sensor configured to detect whether the satellite is in a phase of visibility of the sun,
responsive to a detection of a first phase of visibility of the sun: a step of attitude control using a second control law according to which the magnetic torquers are controlled to limit variations of the Earth's magnetic field in the satellite reference frame by forming attitude control torques along the Z axis, and the inertial actuators are controlled to form torques along the X and Y axes to place and maintain the satellite in an attitude in an inertial reference frame in which the at least one solar generator is directed towards the sun.

2. The method according to claim 1 , comprising, during a first phase of eclipse of the sun detected after the first phase of visibility of the sun, a step of attitude control using the first control law.

3. The method according to claim 1 , comprising:
during a phase of visibility of the sun detected after the first phase of visibility of the sun: a step of attitude control using a third control law which corresponds to the second control law and which further comprises a controlling of the magnetic torquers to form a torque in a direction of the sun, and a controlling of the inertial actuators to form an internal angular momentum in the direction of the sun forming a torque that opposes the torque formed by the magnetic torquers in the direction of the sun,
during a subsequent phase of eclipse of the sun: a step of no attitude control, during which the magnetic torquers and inertial actuators are not controlled.

4. The method according to claim 3 , wherein, during the use of the third control law, the magnetic torquers and inertial actuators are controlled until an internal angular momentum of predetermined norm is reached in the direction of the sun.

5. The method according to claim 1 , wherein the magnetic torquers are controlled according to a biased b-dot law in the first control law and/or in the second control law.

6. The method according to claim 1 , wherein the second control law further comprises a controlling of the magnetic torquers to desaturate the inertial actuators along the X and Y axes.

7. The method according to claim 1 , wherein the step of searching for the sun starts to be executed, during the step of attitude control using the first control law, when a norm of a rotational speed of the satellite on itself in the inertial reference frame becomes less than or equal to a predetermined positive threshold value.

8. The method according to claim 7 , wherein the first predetermined positive threshold value is equal to K1·|w 0 |, wherein w 0 is the orbital angular frequency of the satellite in the inertial reference frame, and wherein K1 is a positive parameter such that 3≤K1≤5.

9. The method according to claim 1 , wherein, during the use of the first control law, the magnetic torquers are controlled to limit the variations of the Earth's magnetic field in the satellite reference frame so as to obtain a rotational speed of the satellite on itself in the inertial reference frame of a norm strictly greater than 2·|w 0 | and less than or equal to 4·|w 0 |, wherein w 0 is the orbital angular frequency of the satellite in the inertial reference frame.

10. Computer program product characterized in that it comprises a set of program code instructions which, when executed by a processor, configure said processor to implement the method for attitude control according to claim 1 .

11. A satellite intended to be placed in a inclined low orbit, comprising at least one solar generator, at least one solar sensor, magnetic torquers capable of forming internal magnetic moments in a satellite reference frame comprising three orthogonal axes X, Y, and Z, and inertial actuators capable of forming internal angular momentums in said satellite reference frame, wherein the at least one solar sensor has a field of view at least 180° wide in the XZ plane around the Z axis, the at least one solar generator is configured to be stationary in the satellite reference frame during a survival mode and to be directed so as to generate electrical energy when the sun (S) is located along the Z axis within the field of view of the at least one solar sensor, said satellite further comprising a control module for controlling the magnetic torquers and inertial actuators, wherein said control module is configured to implement a method for attitude control in survival mode according claim 1 .

12. The satellite according to claim 11 , wherein the at least one solar generator is arranged along the Y axis.

13. The satellite according to claim 11 , comprising two solar sensors each having a field of view at least 100° wide within the XZ plane, and arranged so as to jointly present a field of view at least 180° wide within said XZ plane around the Z axis.

14. The satellite according to claim 11 , wherein the inertial actuators are reaction wheels.

Assignees

• Airbus Defence & Space Sas
• Airbus Oneweb Satellites Sas

Inventors

• Neubrand Yannick
• Jullien Alexandre
• Sachot Marie

Classifications

• B64G1/363

  using sun sensors · CPC title

• B64G1/285 Primary

  using momentum wheels · CPC title

• B64G1/283

  using reaction wheels · CPC title

• B64G1/32

  using earth's magnetic field · CPC title

• B64G1/244 Primary

  Spacecraft control systems · CPC title