# US 6,298,288 B1 · Autonomous gyro scale factor and misalignment calibration

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- Number: US 6,298,288 B1
- Document ID: US6298288B1
- Status: Granted
- Kind: B1
- Application number: 09/212,454
- Filed: 1998-12-16
- Granted: 2001-10-02
- Earliest claimed: 1998-12-16
- Assignees: Hughes Electronics Corporation
- Inventors: Rongsheng Li; Yeong-Wei A. Wu; Garry Didinsky
- Web page: https://patentgrep.com/patent/US6298288B1

## CPC

- B64G 1/365: {using horizon or Earth sensors}
- B64G 1/244: {Spacecraft control systems}
- B64G 1/36: using sensors, e.g. sun-sensors, horizon sensors
- B64G 1/361: {using star sensors}
- B64G 1/363: {using sun sensors}
- B64G 1/369: {using gyroscopes as attitude sensors}
- B64G 1/245: {Attitude control algorithms for spacecraft attitude control}
- B64G: COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR

## Abstract

A system and method of estimating the attitude of a spacecraft is disclosed. A three-axis inertial-based estimate of spacecraft attitude is produced and is compared to a stellar-based estimate of spacecraft attitude. A Kalman filter having states associated with gyro scale factor and/or misalignment errors compares the stellar-based attitude estimate to the inertial-based estimate of attitude and apportions the total error into three time varying matrices. A first time varying matrix is associated with gyro scale factor and misalignment errors, a second time varying matrix is associated with gyro bias errors, and a third time varying matrix is associated with attitude errors. The time varying matrices are applied as corrective feedbacks to the inertial-based estimate of spacecraft attitude and are adaptively adjusted to minimize the error therein.

## Claims

### Claim 1

A system for estimating the attitude of a spacecraft comprising:
first, second, and third inertial sensors;
a stellar position sensor;
a digital filter;
the filter adapted to receive a first estimated spacecraft attitude derived from the stellar position sensor and a second estimated spacecraft attitude derived from the first, second, and third inertial sensors;
the filter further adapted to compare the first and second estimated attitudes to produce corrective data signals that are representative of the misalignment between the inertial sensors and the scale errors associated with each of the three inertial sensors; and
a means for correcting the second estimated spacecraft attitude using the corrective data signals.

### Claim 2 (depends on 1)

The system of claim 1, wherein the first, second, and third inertial sensors are uniquely associated with substantially orthogonal axes.

### Claim 3 (depends on 1)

The system of claim 1, wherein the digital filter adjusts its filtering coefficients to minimize error in the inertially estimated attitude.

### Claim 4 (depends on 3)

The system of claim 3, wherein the digital filter is a Kalman filter.

### Claim 5 (depends on 3)

The system of claim 3, wherein the digital filter is a Kalman filter having fifteen states.

### Claim 6 (depends on 1)

The system of claim 1, wherein the digital filter produces corrective data signals representative of gyro bias errors.

### Claim 7 (depends on 1)

The system of claim 1, wherein the digital filter produces corrective data signals representative of attitude errors.

### Claim 8

A method of estimating the attitude of a spacecraft, the method comprising the steps of:
receiving inertial information from three inertial sensors;
correcting the inertial information using a first time varying matrix having corrective data associated with scale factor and misalignment errors in the inertial information computed by a digital filter;
propagating a three axis attitude using the corrected inertial information to produce a propagated attitude estimate; and
correcting the propagated attitude estimate for attitude errors to produce a final output attitude estimate, the correction being made using a second time varying matrix computed by the digital filter.

### Claim 9 (depends on 8)

The method of claim 8, wherein the steps of generating, correcting, propagating, and correcting are executed more than once.

### Claim 10 (depends on 8)

The method of claim 8, further including the step of correcting the inertial information for sensor bias errors.

### Claim 11 (depends on 8)

The method of claim 8, wherein the digital filter compares measured star positions to the final output attitude estimate and based on the comparison produces the first and second time varying matrices.

### Claim 12 (depends on 11)

The method of claim 11, wherein the first and second time varying matrices are produced using a Kalman filter.

### Claim 13

A system for estimating the attitude of a spacecraft comprising:
a first sensor that acquires three axis inertial data;
a second sensor that acquires star position data;
a first processor that processes the three axis inertial data to produce a first estimate of spacecraft attitude;
a second processor that processes the star position data to produce a second estimate of spacecraft attitude;
a digital filter that compares the first and second estimates of spacecraft attitude, and therefrom produces at least a first time varying matrix having corrective data associated with scale factor and misalignment errors in the three axis inertial data; and
the first processor further adapted to apply the corrective data to the first estimate of spacecraft attitude.

### Claim 14 (depends on 13)

The system of claim 13, wherein the digital filter produces a second time varying matrix comprising corrective data associated with bias errors in the three axis inertial data.

### Claim 15 (depends on 13)

The system of claim 13, wherein the digital filter produces a third time varying matrix comprising corrective data associated with attitude errors in the three axis inertial data.