Guided missile subsystem

• US 4,589,610 A
• Filed: 11/08/1983
• Issued: 05/20/1986
• Est. Priority Date: 11/08/1983
• Status: Expired due to Fees

• Current Assignee: Northrop Grumman Corporation
• Original Assignee: Westinghouse Electric Corporation (Paramount Global (f/k/a ViacomCBS Inc.))
• Inventors: Schmidt, Donald J.
• Primary Examiner(s): Jordan, Charles T.
• Application Number:
  US06/549,861
• Time in Patent Office: 924 Days
• Field of Search: 244/3.19, 244/3.15, 343/5 CM
• US Class Current: 244/3.19
• CPC Class Codes: F41G 7/22 Homing guidance systems
• CPC Class Codes: F41G 7/2246 Active homing systems, i.e. comprising both a transmitter and a receiver
• CPC Class Codes: F41G 7/2286 using radio waves

Abstract

A guided missile subsystem including a Kalmanized radar track loop driven by acceleration signals of the missile generated by an inertial measuring unit (IMU), and a missile control loop driven by estimates of the relative kinematics of the missile and target computed by the radar track loop is disclosed. The IMU driven Kalmanized radar track loop accommodates the use of a high performance radar, like a synthetic aperture radar, for example, which operates to measure radar data at a low rate on the order of 1 Hz, to generate estimates of relative target and missile kinematics to drive the control loop at rates compatible with high performance missile kinematics. The Kalmanized track loop effects an exchange of IMU errors for "dynamic lag" errors of conventional track loops which cannot be modeled very well, and can change very rapidly. In contrast, the IMU errors can be modeled well, and in addition change very slowly which is what permits the Kalmanization function to work well in the track loop at reduced rates. Because of the dynamic exactness of the track loop, very good estimates of the relative kinematics of the missile may be supplied to the control loop to effect more accurate computations of maneuver commands which drive the controls of the missile. Moreover, the Kalmanized track loop does not let large amounts of angle glint noise into the control loop prior to missile impact. An effective bandwidth decrease as glint noise increases is provided without incurring a dynamic lag error penalty.

Claims

1. A missile guidance subsystem disposed onboard a missile and operative during the flight of said missile to cooperate in guiding said missile to the location of a target, said missile guidance subsystem comprising:

a radar including an antenna system, a front end, and a signal processor, said antenna system governed by beam steering commands to maintain the beam pattern of said radar antenna on said target location, said front end for receiving radar echo signals within said beam pattern and conditioning said radar echo signals for processing by said signal processor, said signal processor for deriving true radar measurements of said missile kinematics in relation to said target kinematics from said conditioned radar echo signals;

control means governed by a set of maneuver commands to control said missile kinematics;

an inertial measuring unit for generating signals corresponding to the acceleration of said guided missile in accordance with predetermined spatial coordinates;

means for integrating said acceleration signals of said inertial measuring unit to generate estimates of the relative kinematics of said missile and target in accordance with said spatial coordinates;

means for converting said estimates of the relative kinematics into a priori estimates of radar measurements of said missile and target relative kinematics, and into beam steering commands for said radar antenna system, said signal processor operative to compute signals representative of the differences between corresponding estimated and true radar measurements;

filtering means for deriving error signals based on an estimating function of said computed radar measurement difference signals, said integrating means operative to generate intermediate relative kinematics estimates according to said spatial coordinates in the integration process thereof, said error signals derived in accordance with said spatial coordinates for correcting corresponding intermediate relative kinematics estimates of said integrating means to render said relative kinemaic estimates; and

means for generating said set of maneuver commands based on a control law function of said relative kinematics estimates.

2. A missile guidance subsystem in accordance with claim 1 wherein the integrating means, converting means, radar and filtering means constitute, in combination, a radar tracking loop governed by the acceleration signals generated by the inertial measuring unit;

and wherein said filtering means includes a Kalman filter portion for generating estimates of tracking loop errors based on optimal filter techniques using a priori information of error processes of said tracking loop.

3. A missile guidance subsystem in accordance with claim 2 wherein the Kalman filter portion includes:

a set of Kalman gain vectors for operating on the radar measurement difference signals for generating the tracking loop error estimates; and

means for extracting from the radar measurement difference signals an index of filter performance for use in adjusting said Kalman gain vectors used in said error estimation process to compensate for errors in modeling target accelerations.

4. A missile guidance subsystem in accordance with claim 1 wherein the integrating means, converting means, radar and filtering means constitute, in combination, a radar tracking loop governed by the acceleration signals generated by the inertial measuring unit at a first rate;

wherein the radar signal processor is operative at a second rate, substantially slower than said first rate, to measure the differences between true and estimated radar measurements;

wherein the filtering means is operative to derive the error signals at a rate commensurate with said second rate; and

wherein the integrating means includes means for accommodating the acceleration signals at said first rate and the error signals at said slower rate to generate the relative kinematics estimates at a rate commensurate with said first rate.

5. A missile guidance subsystem in accordance with claim 1 including a navigational update filtering means for deriving error signals based on an estimating function of the radar-measured kinematics difference signals to compensate for errors in the acceleration signals generated by the inertial measuring unit.

6. A missile guidance subsystem in accordance with claim 1 wherein the inertial measuring unit includes means for generating the acceleration signals corresponding to the acceleration of the missile along each coordinate of a predetermined orthogonal 3-axis coordinate system;

wherein the integrating means includes an integration function to generate intermediate relative velocity and position estimates from said acceleration signals corresponding to each axis of said orthogonal coordinate system;

wherein the filtering means derives an error signal for each intermediate relative position and velocity estimate from the differences between corresponding estimated and true radar measurements; and

wherein the integrating means includes a correction function to correct each intermediate relative position and velocity estimate with its corresponding error signal to render a relative position and velocity estimate for each axis of said orthogonal coordinate system.

7. A missile guidance subsystem in accordance with claim 6 wherein the converting means includes means for converting the relative position and velocity estimates into estimated radar measurements including range, range rate, and antenna beam pointing angle errors in azimuth and elevation;

wherein the signal processor includes means for deriving true radar measurements including true range, true range rate, and true antenna beam pointing angle errors in azimuth and elevation from the conditioned radar echo signals and said corresponding estimated radar measurements, said signal processor further including means for measuring the differences between said corresponding estimated and true radar measurements and for generating signals representative thereof; and

wherein the filtering means includes means for deriving the position and velocity error signals for each axis of the orthogonal coordinate system based on an estimating function of said measurement difference signals in range, range rate, antenna azimuth beam pointing angle error and antenna elevation beam pointing angle error.

8. A missile guidance subsystem in accordance with claim 7 wherein the integrating means, converting means, radar and filtering means constitute, in combination, a radar tracking loop governed by the acceleration signals generated by the inertial measuring unit;

and wherein the filtering means includes a Kalman filter portion for generating estimates of position and velocity tracking loop errors for the 3-axis orthogonal coordinate system based on optimal filter techniques using a priori information of error processes of said tracking loop.

9. A missile guidance subsystem in accordance with claim 8 wherein the Kalman filter portion includes:

a set of Kalman gain vectors for operating on the radar measurement difference signals in range, range rate, and antenna beam pointing angle errors in azimuth and elevation for generating estimates of position, velocity and acceleration tracking loop errors; and

means for extracting from said radar measurement difference signals an index of filter performance for use in adjusting said Kalman gain vectors used in said error estimation process to compensate for errors in modeling target acceleration.

10. A missile guidance subsystem in accordance with claim 6 wherein the generating means includes means for generating a set or maneuver commands according to an up-down and left-right coordinate axis system referenced to the velocity vector of the missile, said generation based on a control law function of the relative position and velocity estimates of the predetermined 3-axis orthogonal system generated by the integrating means.

11. A missile guidance subsystem in accordance with claim 10 wherein the control law function of said generating means includes a proportional navigation control law function.

12. A missile guidance subsystem in accordance with claim 6 wherein the integrating means, the generating means, the control means and the inertial measuring unit, in combination, constitute a control loop of the guidance subsystem.

13. A missile guidance subsystem in accordance with claim 6 including a navigational update loop comprising a navigation update filtering means governed by signals including the radar measurement difference signals and predetermined check point coordinates to derive error correction signals for the inertial measuring unit.

14. A missile guidance subsystem in accordance with claim 6 wherein the radar signal processor includes a synthetic aperture radar processor governed by the relative position and velocity estimates to compensate for the motion of the missile in deriving a radar image of a ground target location.

15. A missile guidance subsystem in accordance with claim 6 wherein the converting means includes means for converting the relative position estimates of the predetermined orthogonal 3-axis coordinate system into beam steering commands including estimated antenna beam pointing angle errors along the azimuth and elevation axis coordinates of the radar antenna;

and wherein the radar antenna includes means governed by said estimated antenna azimuth and elevation beam pointing angle errors to steer the antenna beam generated thereby.

16. A missile guidance subsystem in accordance with claim 15 wherein the converting means includes means for converting the relative position estimates of the predetermined orthogonal 3-axis coordinate system into beam steering commands including estimated antenna beam pointing angle errors along azimuth and elevation axis in accordance with unit direction vectors coordinatized in antenna system reference axes for positioning an antenna beam electronically.