Launch vehicle testing system

• US 9,665,453 B2
• Filed: 09/12/2012
• Issued: 05/30/2017
• Est. Priority Date: 09/12/2012
• Status: Active Grant

• Current Assignee: Honeywell International Inc.
• Original Assignee: Honeywell International Inc.
• Inventors: Martin, Kenneth Lee, Sjoquist, Gregory Allan
• Primary Examiner(s): McCullough, Michael
• Assistant Examiner(s): Butler, Michael E
• Application Number:
  US13/611,296
• Publication Number:
  US 20140074347A1
• Time in Patent Office: 1,721 Days
• Field of Search: 701 315
• CPC Class Codes: G06F 11/2005 using redundant communication controllers
• CPC Class Codes: G06F 11/2294 by remote test

Abstract

A system for testing a launch vehicle comprises a flight control section including flight electronics, and a non-flight test section coupled to the flight control section. The non-flight test section comprises a first control module in communication with the flight electronics, and a second control module that provides redundancy in communication with the flight electronics. A first network serial interface is coupled to the first control module and configured to provide communication to ground support equipment, and a second network serial interface is coupled to the second control module and configured to provide communication to the ground support equipment. Only one of the first and second control modules is selected at a given point in time to send data to the flight control section. When a non-flight hardware error occurs in the selected control module, the other control module is automatically selected to send data to the flight control section.

Claims

1. A system for testing a launch vehicle, the system comprising:

a flight control section including radiation hardened flight electronics in the launch vehicle, the flight electronics including a radiation hardened flight processor;

a non-flight test section operatively coupled to the flight control section, the non-flight test section comprising;

a first electronic control module in operative communication with the flight electronics; and

a second electronic control module that provides redundancy with respect to the first electronic control module, the second electronic control module in operative communication with the flight electronics;

wherein the first and second electronic control modules include non-radiation hardened electronics;

a first flight processor test interface coupled between the first electronic control module and the radiation hardened flight processor;

a second flight processor test interface coupled between the second electronic control module and the radiation hardened flight processor;

a flight processor control interface configured to assure that the test interfaces are controlled by the flight electronics;

an isolation interface comprising a set of radiation hardened buffers configured to assure that the test interfaces are isolated from the flight processor control interface;

a set of flight control section interfaces configured to assure that enables and direction control to the radiation hardened buffers are managed by the radiation hardened flight processor;

a power interface configured to assure that power switching is controlled by the radiation hardened flight processor and radiation hardened switching electronics, the power interface further configured to assure that test logic is powered-off during flight such that the non-radiation hardened electronics in the non-flight test section are electrically isolated from the flight electronics; and

an embedded software control mechanism configured to;

select one of the first and second electronic control modules to send data to the flight control section during a test;

determine whether a non-flight hardware error has occurred in the selected electronic control module;

automatically select the other electronic control module to send data to the flight control section after the non-flight hardware error is determined to have occurred; and

isolate the electronic control module where the non-flight hardware error occurred;

wherein the first and second flight processor test interfaces are configured to provide for redundant flight system test and debug, such that any flight system faults are isolated from any non-flight system faults.

2. The system of claim 1, wherein the flight electronics are space qualified electronics.

3. The system of claim 1, wherein the radiation hardened buffers include:

a first voltage translation buffer operatively coupled between the first electronic control module and the flight control section; and

a second voltage translation buffer operatively coupled between the second electronic control module and the flight control section;

wherein the first and second voltage translation buffers are configured to isolate the flight electronics from the first and second electronic control modules during flight of the vehicle.

4. The system of claim 1, wherein the first and second electronic control modules each comprise:

a processor;

a first memory unit operatively coupled to the processor and including a non-volatile memory storage;

a second memory unit operatively coupled to the processor and including a dual port memory; and

a third memory unit operatively coupled to the processor and including an operating system memory.

5. The system of claim 4, wherein the processor in each of the first and second electronic control modules is in operative communication with a respective serial interface connector.

6. The system of claim 4, wherein the processor in each of the first and second electronic control modules is operatively coupled to a respective one of the first and second flight processor test interfaces.

7. The system of claim 1, further comprising:

a first network serial interface communicatively coupled to the first electronic control module and configured to provide signal communication between the first electronic control module and ground support equipment; and

a second network serial interface communicatively coupled to the second electronic control module and configured to provide signal communication between the second electronic control module and the ground support equipment;

wherein the first and second network serial interfaces are Ethernet connectors.

8. The system of claim 1, wherein the first and second electronic control modules are covered with a protective material that provides protection from vibration and shock.

9. The system of claim 1, wherein the flight control section is configured to control power application to the non-flight test section when needed prior to launch of the vehicle.

10. The system of claim 1, wherein the first and second electronic control modules are configured to receive discrete signals from outside of the system.

11. The system of claim 1, wherein the non-flight test section is configured to be powered down during flight of the launch vehicle.

12. The system of claim 1, wherein the first and second electronic control modules are configured to receive select and control signals from the flight control section interfaces.

13. The system of claim 1, wherein the first and second electronic control modules are configured to send interrupt signals to the flight control section.

14. The system of claim 1, wherein the first and second electronic control modules are configured to receive or send a plurality of signals comprising shared data, control, address, simulation data, flight processor test interface data, and reset.

15. A method for testing a launch vehicle, the method comprising:

performing a test operation using a testing system for the launch vehicle, the testing system comprising;

a flight control section including radiation hardened flight electronics in the launch vehicle, the flight electronics including a radiation hardened flight processor;

a non-flight test section operatively coupled to the flight control section, the non-flight test section comprising;

a first electronic control module in operative communication with the flight electronics; and

a second electronic control module that provides redundancy with respect to the first electronic control module, the second electronic control module in operative communication with the flight electronics;

wherein the first and second electronic control modules include non-radiation hardened electronics;

wherein the test operation is performed by a method comprising;

selecting one of the first and second electronic control modules to send data to the flight control section;

determining whether a non-flight hardware error has occurred in the selected electronic control module;

automatically selecting the other electronic control module to send data to the flight control section after the non-flight hardware error is determined to have occurred; and

isolating the electronic control module where the non-flight hardware error occurred.

16. The method of claim 15, wherein the test system further comprises:

a first network serial interface communicatively coupled to the first electronic control module; and

a second network serial interface communicatively coupled to the second electronic control module;

wherein the first and second network serial interfaces are connected to high-speed commercial media.

17. The method of claim 16 wherein the first and second network serial interfaces are connected to an Ethernet network.

18. The method of claim 15, wherein the flight control section controls power application to the non-flight test section when needed prior to launch of the vehicle.

19. The method of claim 15, wherein the first and second electronic control modules receive discrete signals from ground support equipment.

20. The method of claim 15, further comprising powering down the non-flight test section during flight of the launch vehicle.