Missile navigation method

• US 8,242,423 B2
• Filed: 06/02/2009
• Issued: 08/14/2012
• Est. Priority Date: 06/02/2009
• Status: Active Grant

• Current Assignee: Raytheon Company (Rtx Corporation)
• Original Assignee: Raytheon Company (Rtx Corporation)
• Inventors: Geswender, Chris E., Cribb, Matthew W., Hanson, Clyde R.
• Primary Examiner(s): Gregory, Bernarr
• Application Number:
  US12/476,486
• Publication Number:
  US 20120043410A1
• Time in Patent Office: 1,169 Days
• Field of Search: 244 31- 319, 244 32- 33, 244/158.1, 244/164, 89/11.1, 701/1, 701/3, 701/200, 701/207, 701/220, 701/221, 701/4, 701/13, 701/11, 701/12, 701/400, 701/408, 701468-491, 701500-512, 342/73, 342/89, 342/94, 342/95, 702/85, 702/104, 702/105, 702/127, 702/141, 702150-154, 702/182, 702/92, 702/142, 702/145, 702/189, 702/190, 702/191, 731/78.R, 731/81
• US Class Current: 244/3.2
• CPC Class Codes: F42B 15/01 Arrangements thereon for guidance or control steering arrangements F42B10/60; aircraft flight control B64C; guidance systems other than those installed aboard F41G7/00, F41G9/00; locating by use of radio or other waves G01S; flight control in general G05D1/00; computer aspects G06]

Abstract

A missile has a pair of systems to provide acceleration information during flight. The primary system is a microelectromechanical systems (MEMS) inertial measurement unit (IMU) that provides accurate rate sensor output, such as providing pitch and yaw rates, at low cost, over a wide range of conditions. However MEMS IMUs are susceptible to temporary incorrect responses when subjected to shocks, such as acoustic-range shocks, for instance in the range of 10-20 kHz. The missile includes a secondary system to temporarily provide acceleration data during the periods following shocks, when the MEMS IMU does not provide valid (reliable or usable) rate sensor output, for use in estimating pseudo pitch and yaw rates. The secondary system may be an accelerometer that does not provide navigation-quality acceleration data, but does provide a sufficiently accurate response in order to maintain stable flight during the post-shock period.

Claims

1. A method of navigating a missile, the method comprising:

using a microelectromechanical systems (MEMS) inertial measurement unit (IMU) as a primary data source, to provide angular rate data to a navigation system of the missile, wherein the using is temporarily suspended during periods following shock events to the missile;

using at least a secondary data source to provide pseudo rate data to the navigation system during the periods following the shock events; and

providing either the angular rate data or the pseudo rate data to an autopilot of the navigation system that directs a control section of the missile, to alter a course of the missile.

2. The method of claim 1, wherein the secondary data source includes an accelerometer.

3. The method of claim 2, wherein the accelerometer is a two-axis accelerometer configured to measure accelerations in pitch and yaw directions.

4. The method of claim 2,

wherein the MEMS IMU provides navigation quality accuracy;

and

wherein the accelerometer has an accuracy lower that is less than navigation quality accuracy.

5. The method of claim 1, wherein the secondary data source includes an magnetometer.

6. The method of claim 1, further comprising selecting use of either the angular rate data or the pseudo rate data by the autopilot using IMU validity logic, wherein the IMU validity logic bases its selection on data from the MEMS IMU.

7. The method of claim 6, wherein the selecting includes comparing acceleration measured by the MEMS IMU with a predetermined acceleration validity threshold.

8. The method of claim 1,

wherein the missile includes a rate estimator operatively coupled to the secondary data source;

and

further comprising producing the pseudo rate data in the rate estimator using acceleration data from the secondary data source.

9. The method of claim 8, wherein the producing includes using rates of change of angle of attack as approximations of the angular rates of change.

10. A missile navigation system comprising:

a microelectromechanical systems (MEMS) inertial measurement unit (IMU);

a secondary data source;

IMU validity logic;

a rate estimator; and

an autopilot;

wherein the secondary data source is operatively coupled to the rate estimator to provide data to the rate estimator to allow the rate estimator to produce pseudo angle rate changes in yaw and pitch directions;

wherein the MEMS IMU and the rate estimator are coupled to the validity logic, which is in turn operatively coupled to the autopilot; and

wherein the validity logic determines validity of data from the MEMS and IMU, and supplies to the autopilot either angle rate change data from the MEMS IMU or the pseudo angle rate changes produced by the rate estimator.

11. The missile of claim 10, wherein the secondary data source includes an accelerometer.

12. The missile of claim 11, wherein the accelerometer is a two-axis accelerometer configured to measure accelerations in pitch and yaw directions.

13. The missile of claim 11,

wherein the MEMS IMU provides navigation quality accuracy;

and

wherein the accelerometer has an accuracy lower that is less than navigation quality accuracy.

14. The missile of claim 10, wherein the secondary data source includes a magnetometer.

15. The missile of claim 10, wherein the rate estimator uses rates of change of angle of attack as approximations of the angular rates of change.

16. The missile of claim 10, wherein the IMU validity logic and the rate estimator are embodied as a circuit card assembly.