Attitude determination using infrared earth horizon sensors

Patent metadata

Field	Value
Publication number	US-9809328-B2
Application number	US-201514644425-A
Country	US
Kind code	B2
Filing date	Mar 11, 2015
Priority date	Apr 22, 2014
Publication date	Nov 7, 2017
Grant date	Nov 7, 2017

Abstract

Described herein are systems and methods for attitude determination using infrared Earth horizon sensors (EHSs) with Gaussian response characteristics. Attitude information is acquired by detecting Earth's infrared electromagnetic radiation and, subsequently, determining the region obscured by Earth in the sensors' fields of view to compute a nadir vector estimation in the spacecraft's body frame. The method can be applied when two sensors, each with known and distinct pointing directions, detect the horizon, which is defined as having their fields of view partially obscured by Earth. The method can be implemented compactly to provide high-accuracy attitude within small spacecraft, such as CubeSat-based satellites.

First claim

What is claimed is:

1. A system for spacecraft attitude determination, the system comprising:
a plurality of horizon sensors, each of the plurality of horizon sensors having a respective field-of-view (FOV), each of the plurality of horizon sensors having a sensitivity to the portion of the respective FOV obscured by Earth;
an attitude determination module coupled to the plurality of horizon sensors, the attitude determination module configured to:
obtain horizon sensor readings from the horizon sensors;
determine an Earth disk radius in a spacecraft's body frame;
convert the horizon sensor readings to nadir angles using a Gaussian approximation including dividing the horizon sensor FOV into a plurality of constant sensitivity regions each having a Gaussian approximation coefficient; and
estimate a nadir vector using the nadir angles and the Earth disk radius.

2. The system of claim 1 further comprising a reference sensor coupled to provide reference sensor readings to the attitude determination module, wherein the attitude determination module is configured to:
determine a plurality of possible nadir vector solutions using the nadir angles and the Earth disk radius;
obtain a reference sensor reading; and
choose one of the plurality of possible nadir vector solutions based upon the reference sensor reading.

3. The system of claim 2 , wherein the plurality of horizon sensors comprise a first horizon sensor and a second horizon sensor mounted along orthogonal axes.

4. The system of claim 1 wherein at least one of the plurality of horizon sensors comprises a thermopile.

5. The system of claim 1 wherein at least one of the plurality of horizon sensors comprises an infrared radiation (IR) sensor.

6. The system of claim 3 wherein the reference sensor comprises a magnetometer or a Sun sensor.

7. The system of claim 1 further comprising an attitude control module coupled to receive estimated nadir vector information from the attitude determination module, the attitude control module configured to adjust a spacecraft's attitude based upon received estimated nadir vector information.

8. The system of claim 1 wherein the attitude determination module is configured to:
receive spacecraft position information; and
determine an Earth disk radius in a spacecraft's body frame based upon the spacecraft position information.

9. A method for spacecraft attitude determination, the method comprising:
obtaining horizon sensor readings from a plurality of horizon sensors, each of the plurality of horizon sensors having a respective field-of-view (FOV), each of the plurality of horizon sensors having a sensitivity to the portion of the respective FOV obscured by Earth;
determining an Earth disk radius in a spacecraft's body frame;
converting the horizon sensor readings to nadir angles using a Gaussian approximation including dividing the horizon sensor FOV into a plurality of constant sensitivity regions each having a Gaussian approximation coefficient; and
estimating a nadir vector using the nadir angles and the Earth disk radius.

10. The method of claim 9 further comprising:
determining a plurality of possible nadir vector solutions using the nadir angles and the Earth disk radius;
obtaining a reference sensor reading from a reference sensor; and
choosing one of the plurality of possible nadir vector solutions based upon the reference sensor reading.

11. The method of claim 10 , wherein obtaining horizon sensor readings from horizon sensors comprises obtaining a horizon sensor readings from a first horizon sensor and from a second horizon sensor mounted along orthogonal axes.

12. The method of claim 9 wherein obtaining horizon sensor readings from horizon sensors comprises obtaining readings from thermopiles.

13. The method of claim 9 wherein obtaining horizon sensor readings from horizon sensors comprises obtaining readings from infrared radiation (IR) sensors.

14. The method of claim 11 wherein obtaining a reference sensor reading comprises obtaining a reading from a magnetometer or a Sun sensor.

15. The method of claim 9 further comprising adjusting a spacecraft's attitude based upon an estimated nadir vector.

16. The method of claim 9 further comprising receiving spacecraft position information, wherein determining an Earth disk radius in a spacecraft's body frame comprises determining an Earth disk radius based upon the spacecraft position information.

17. The system of claim 1 wherein the attitude determination module is configured to determine the Earth disk radius in the spacecraft's body frame using a geocentric latitude of the spacecraft, an equatorial radius of the Earth, and a polar radius of the Earth.

18. The system of claim 1 wherein the plurality of horizon sensors comprise a first horizon sensor having a first boresight vector in the spacecraft's body frame and a second horizon sensor having a second boresight vector in the spacecraft's body frame, wherein the nadir angles include a first nadir angle and a second nadir angle, and wherein the attitude determination module is configured to estimate the nadir vector by finding the intersection of a first cone defined by the first boresight vector and the first nadir angle and a second cone defined by the second boresight vector and the second nadir angle.

19. A method for spacecraft attitude determination, the method comprising:
obtaining horizon sensor readings from a plurality of horizon sensors, each of the plurality of horizon sensors having a respective field-of-view (FOV), a sensitivity to the portion of the respective FOV obscured by Earth, and a boresight vector in a spacecraft body frame;
determining an Earth disk radius in a spacecraft's body frame using a geocentric latitude of the spacecraft, an equatorial radius of the Earth, and a polar radius of the Earth;
converting the horizon sensor readings to nadir angles; and
estimating a nadir vector using the Earth disk radius and by finding the intersection of a first cone defined by the boresight vector of a first one of the plurality of horizon sensors and a first one of the nadir angles and a second cone defined by the boresight vector of a second one of the plurality of horizon sensors and a second one of the nadir angles.

Assignees

• Massachusetts Inst Technology

Inventors

• Nguyen Tam Nguyen Thuc
• Cahoy Kerri L
• Quadrino Meghan K

Classifications

• B64G1/286

  using control momentum gyroscopes (CMGs) · CPC title

• B64G1/366

  using magnetometers · CPC title

• B64G1/26

  using jets · CPC title

• B64G1/365 Primary

  using horizon or Earth sensors · CPC title

• B64G2001/245

  Operations & Transport · mapped topic