Multi-pulse rocket motor

Patent metadata

Field	Value
Publication number	US-2021381468-A1
Application number	US-202016896271-A
Country	US
Kind code	A1
Filing date	Jun 9, 2020
Priority date	Jun 9, 2020
Publication date	Dec 9, 2021
Grant date	—

Abstract

A flight test system uses a flight termination destruct charge that is configured to overpressurize a pressure vessel in a rocket motor to terminate thrust. The flight termination destruct charge is an electroexplosive detonator arranged on a final burn surface of a propellant contained in the pressure chamber. In a multi-pulse rocket motor, one of the pulses is ignited by the activation of the detonator. The activated detonator is configured to ignite the propellant grain without venting of the gas resulting from the burning of the propellant. Due to the burning of the propellant, the surface area in the pressure vessel is increased which causes increased pressure in the pressure vessel until a critical pressure is reached. When the critical pressure is reached, the rocket motor casing structural capabilities are exceeded. The overpressurized rocket motor casing then ruptures and thrust of the rocket motor is terminated.

First claim

1 . A flight test system for a rocket motor, the flight test system comprising:
flight termination system electronics arranged in a casing of the rocket motor; and an electroexplosive detonator coupled to a propellant inside a pressure vessel defined by the casing, the electroexplosive detonator being configured for activation by the flight termination system electronics to ignite the propellant whereby the pressure vessel is overpressurized to rupture the casing.

2 . The flight test system according to claim 1 , wherein the electroexplosive detonator is arranged on a final burn surface of the propellant.

3 . The flight test system according to claim 1 , wherein the electroexplosive detonator is arranged on a forward-facing surface of the propellant.

4 . The flight test system according to claim 1 , wherein the casing has a forward-facing dome portion and the electroexplosive detonator is mounted to the dome portion.

5 . The flight test system according to claim 4 , wherein the dome portion defines a pluggable port for receiving the electroexplosive detonator.

6 . The flight test system according to claim 1 , wherein the electroexplosive detonator is configured to ignite the propellant without a vent path.

7 . The flight test system according to claim 1 , wherein the propellant comprises a solid propellant grain.

8 . A flight test system comprising:
a multi-pulse rocket motor module including a first pulse chamber containing a first burnable propellant that is burned during a first stage of the multi-pulse rocket motor and a second pulse chamber containing a second burnable propellant that is burned during a second stage of the multi-pulse rocket; flight termination system electronics arranged in the multi-pulse rocket motor module; and an electroexplosive detonator coupled to the second burnable propellant inside the second pulse chamber, the electroexplosive detonator being activated by the flight termination system electronics to ignite the second burnable propellant whereby the pressure vessel is overpressurized.

9 . The flight test system according to claim 8 , wherein the electroexplosive detonator is arranged on a final burn surface of the second burnable propellant.

10 . The flight test system according to claim 8 , wherein the electroexplosive detonator is arranged on a forward surface of the second burnable propellant.

11 . The flight test system according to claim 8 , wherein the second pulse chamber has a forward-facing dome portion, wherein the electroexplosive detonator is mounted to the dome portion.

12 . The flight test system according to claim 11 , wherein the second pulse chamber defines a pluggable port for receiving the electroexplosive detonator.

13 . The flight test system according to claim 8 , wherein the electroexplosive detonator is configured to ignite the second burnable propellant without a vent path.

14 . The flight test system according to claim 8 , wherein the first burnable propellant and the second burnable propellant comprise a solid propellant grain.

15 . A method of flight testing a multi-pulse rocket motor, the method comprising:
using an additional pulse that occurs after a first pulse in the rocket motor as a thrust termination charge; overpressurizing a pulse chamber containing a burnable propellant during the additional pulse; and rupturing the burnable propellant to fail a casing of the rocket motor and terminate thrust of the rocket motor.

16 . The method according to claim 15 further comprising:
selecting an electroexplosive detonator from a plurality of detonators based on at least one of a size of the pulse chamber, size of the burnable propellant, burn rate of the burnable propellant, and density of the burnable propellant;
inserting the detonator into the pulse chamber; and

activating the electroexplosive detonator as the thrust termination charge.

17 . The method according to claim 16 further comprising:
coupling the electroexplosive detonator to a final burn surface of the burnable propellant; and
igniting the burnable propellant without a vent path using the electroexplosive detonator.

18 . The method according to claim 17 further comprising:
inserting flight termination system electronics in a casing of the rocket motor; and
activating the electroexplosive detonator using the flight termination system electronics.

19 . The method according to claim 16 further comprising integrating the electroexplosive detonator into the pulse chamber of a predefined casing of the rocket motor.

20 . The method according to claim 19 further comprising:
plugging a port formed in the casing of the pulse chamber prior to flight testing;
unplugging the port; and
inserting the electroexplosive detonator into the port for flight testing.

Assignees

• Raytheon Co

Inventors

• Lefcourt Adam I
• Meschberger John A

Classifications

• F02K9/28

  having two or more propellant charges with the propulsion gases exhausting through a common nozzle · CPC title

• F02K9/38

  Safety devices, e.g. to prevent accidental ignition · CPC title

• F02K9/92

  incorporating means for reversing or terminating thrust · CPC title

• F02K9/96 Primary

  characterised by specially adapted arrangements for testing or measuring · CPC title

• F02K9/95

  characterised by starting or ignition means or arrangements (safety devices F02K9/38) · CPC title