Method and apparatus for achieving sole means navigation from global navigation satelite systems

• US 6,317,688 B1
• Filed: 01/31/2000
• Issued: 11/13/2001
• Est. Priority Date: 01/31/2000
• Status: Expired due to Term

• Current Assignee: Rockwell Collins, Inc. (Raytheon Technologies Corporation d/b/a RTX)
• Original Assignee: Rockwell Collins, Inc. (Raytheon Technologies Corporation d/b/a RTX)
• Inventors: Hwang, Patrick Y., Bruckner, Juergen M.
• Primary Examiner(s): Cuchlinski, Jr., William A.
• Assistant Examiner(s): MANCHO, RONNIE M
• Application Number:
  US09/494,498
• Time in Patent Office: 652 Days
• Field of Search: 701/213, 701/214, 701/215, 701/220, 701/221, 340/990, 340/995, 340/994, 342/357.08, 342/357.14, 342/357.06, 342/357.05, 342/353, 342/352, 375/259, 375/357, 244/3.2, 244/3.1
• US Class Current: 701/213
• CPC Class Codes: G01C 21/005 with correlation of navigation data from several sources, e.g. map or contour matching G01C21/30 takes precedence
• CPC Class Codes: G01C 21/165 combined with non-inertial navigation instruments
• CPC Class Codes: G01S 19/08 providing integrity information, e.g. health of satellites or quality of ephemeris data
• CPC Class Codes: G01S 19/20 Integrity monitoring, fault detection or fault isolation of space segment
• CPC Class Codes: G01S 19/49 whereby the further system is an inertial position system, e.g. loosely-coupled
• CPC Class Codes: G01S 19/53 Determining attitude

Abstract

A sole means global navigation apparatus adapted for use on an aircraft includes a GPS receiver to provide GPS measurement data and an inertial sensor system adapted to provide inertial translational and rotational data which during time periods is independent of the GPS position data. A navigation system coupled to both of the GPS receiver(s) and the inertial sensor system determines a navigation solution as a function of both the condition of the GPS satellite data and the uncertainty in the Inertial data. An augmentation system coupled to the navigation solution determining system is used to increase the accuracy and/or integrity of the GPS/inertial sensor, thereby achieving style means navigation requirements.

Claims

1. A global positioning system (GPS) navigation apparatus adapted for use on an aircraft, the apparatus comprising:

a GPS receiver adapted to provide GPS measurement data;

an inertial sensor adapted to provide inertial measurement data which is independent of the GPS measurement data;

navigation solution determining means coupled to each of the GPS receiver and the inertial sensor for determining a navigation solution as a function of both the GPS measurement data and the inertial measurement data; and

an augmentation system coupled to the navigation solution determining means wherein the augmentation system is adapted to reduce errors in the inertial measurement data used by the navigation solution determining means in a manner which is independent of the GPS measurement data provided by the GPS receiver.

2. The GPS navigation apparatus of claim 1, wherein the navigation solution determining means further comprises:

a first Kalman filter coupled to the GPS receiver and adapted to filter the GPS measurement data to provide first filtered GPS data;

a second Kalman filter coupled to the inertial sensor and adapted to filter the inertial measurement data to provide first filtered inertial data; and

a processor adapted to implement receiver autonomous integrity monitoring algorithms using the first filtered GPS data and the first filtered inertial data to provide the navigation solution.

3. The GPS navigation apparatus of claim 2, and further comprising an integrity monitoring receiver adapted to receive integrity monitoring signals from a ground based satellite integrity monitoring station, the integrity monitoring signals being indicative of whether a satellite failure has been detected by the satellite integrity monitoring station.

4. The GPS navigation apparatus of claim 3, and further comprising:

a third Kalman filter coupled to the inertial sensor and adapted to filter the inertial measurement data to provide second filtered inertial data; and

wherein the navigation solution determining means is adapted to alternate between utilizing the first filtered inertial data and utilizing the second filtered inertial data in determining the navigation solution.

5. The GPS navigation apparatus of claim 2, wherein the augmentation system includes another GPS antenna system allowing GPS attitude data to be derived.

6. A global positioning system (GPS) navigation apparatus providing a sole source of navigation for an aircraft, the apparatus comprising:

GPS receiver means for providing GPS measurement data;

inertial sensor means for providing inertial measurement data which during time periods is independent of the GPS measurement data;

navigation solution determining means for generating navigation data as a function of both the GPS measurement data and the inertial measurement data; and

augmentation means for optimizing an accuracy of the inertial measurement data.

7. A method of generating navigation data using a global positioning system (GPS) navigation apparatus adapted for use on an aircraft, the method comprising:

generating GPS measurement data using a GPS receiver;

providing augmentation data independently of the GPS measurement data;

generating inertial measurement data, which during time periods is independent of the GPS measurement data using an inertial sensor; and

generating the navigation data as a function of both the GPS measurement data, the inertial measurement data, and the augmentation data.

8. The method of claim 7, and further comprising;

applying Kalman filter functions to the GPS measurement data to provide first filtered GPS data;

applying Kalman filter functions to the inertial measurement data to provide first filtered inertial data; and

implementing receiver autonomous integrity monitoring (RAIM) algorithms using the first filtered GPS data and the first filtered inertial data to provide the navigation solution data.

9. The method of claim 8, and further comprising:

applying Kalman filter functions to the inertial measurement data to provide second filtered inertial data; and

receiving integrity monitoring signals from a ground based satellite integrity monitoring station, the integrity monitoring signals being indicative of whether a satellite failure has been detected by the satellite integrity monitoring station; and

alternating between implementing the RAIM algorithms using the first filtered inertial data and the second filtered inertial data to determine the navigation solution.

10. The method of claim 7, wherein providing augmentation data further comprises receiving GPS attitude data from a multi-antenna GPS attitude system.

11. The method of claim 10, wherein generating the inertial data further includes generating the inertial data using the inertial sensor and as a function of the GPS attitude data during the time periods.

12. A global positioning system (GPS) navigation apparatus adapted for use on an aircraft, the apparatus comprising:

a GPS receiver adapted to provide GPS measurement data;

an inertial sensor adapted to provide inertial measurement data which is independent of the GPS measurement data;

navigation solution determining means coupled to each of the GPS receiver and the inertial sensor for determining a navigation solution as a function of both the GPS measurement data and the inertial measurement data and wherein the navigation solution determining means further comprises;

a first Kalman filter coupled to the GPS receiver and adapted to filter the GPS measurement data to provide first filtered GPS data;

a second Kalman filter coupled to the inertial sensor and adapted to filter the inertial measurement data to provide first filtered inertial data;

a third Kalman filter coupled to the inertial sensor and adapted to filter the inertial measurement data to provide second filtered inertial data; and

a processor adapted to implement receiver autonomous integrity monitoring algorithms using the first filtered GPS data and the first filtered inertial data to provide the navigation solution and

wherein the navigation solution determining means is adapted to alternate between utilizing the first filtered inertial data and utilizing the second filtered inertial data in determining the navigation solution and

wherein the navigation solution determining means is adapted to alternate between utilizing the first filtered inertial data and the second filtered inertial data during consecutive windows of time, each of the consecutive windows of time starting at a time of receipt of the integrity monitoring signals by the integrity monitoring receiver;

an augmentation system coupled to the navigation solution determining means wherein the augmentation system is adapted to reduce errors in the inertial measurement data used by the navigation solution determining means in a manner which is independent of the GPS measurement data provided by the GPS receiver; and

an integrity monitoring receiver adapted to receive integrity monitoring signals from a ground based satellite integrity monitoring station, the integrity monitoring signals being indicative of whether a satellite failure has been detected by the satellite integrity monitoring station.

13. The GPS navigation apparatus of claim 12, wherein the navigation solution determining means is adapted to determine the navigation solution as a function of a selected one of the first and second filtered inertial data, and wherein the navigation solution determining means is adapted to utilize a non-selected one of the first and second Kalman filters to begin a new pure inertial solution at a time corresponding to a next scheduled transmission of the integrity monitoring signals from the integrity monitoring station.

14. The GPS navigation apparatus of claim 13, wherein the first, second and third Kalman filters are implemented in the processor.

15. A method of generating navigation data using a global positioning system (GPS) navigation apparatus adapted for use on an aircraft, the method comprising:

generating GPS measurement data using a GPS receiver;

providing augmentation data independently of the GPS measurement data;

generating inertial measurement data, which during time periods is independent of the GPS measurement data using an inertial sensor;

applying Kalman filter functions to the GPS measurement data to provide first filtered GPS data;

applying Kalman filter functions to the inertial measurement data to provide first filtered inertial data;

applying Kalman filter functions to the inertial measurement data to provide second filtered inertial data;

implementing receiver autonomous integrity monitoring (RAIM) algorithms using the first filtered GPS data and the first filtered inertial data to substantiate the GPS measurement data, the augmentation data and the inertial measurement;

receiving integrity monitoring signals from a ground based satellite integrity monitoring station, the integrity monitoring signals being indicative of whether a satellite failure has been detected by the satellite integrity monitoring station;

alternating between implementing the RAIM algorithms using the first filtered inertial data and the second filtered inertial data to determine the navigation solution wherein alternating further comprises alternating between utilizing the first filtered inertial data and the second filtered inertial data during consecutive windows of time, with each of the consecutive windows of time starting at a time corresponding to a time of receipt of the integrity monitoring signals; and

generating navigation data as a function of both the GPS measurement data, the inertial measurement data, and the augmentation data.