US2459163A

US2459163A - Thermal igniter

Info

Publication number: US2459163A
Application number: US538314A
Authority: US
Inventors: Clarence N Hickman
Original assignee: Individual
Current assignee: Individual
Priority date: 1944-06-01
Filing date: 1944-06-01
Publication date: 1949-01-18
Anticipated expiration: 1966-01-18

Description

c. N. HICKMAN THERMAL IGNITER Jan. 18 1949.

Filed June 1, 1944 Clarence N -Hitkman Patented Jan. 18, 1949 $459,163 THERMAL IGNITER Clarence N. HickmanpJackson Heights, N. 35., assignor to the Unitedv States of America as represented by the Secretary of War Application June 1, 1944, Serial No. 538,314

(Granted under the act of March 3, 1883, as amended April 30, 1928; 370 0. G. 757) 5 Claims.

The invention described herein may be manufactured and used by or for the Government for governmental purposes without the payment to me of any royalty thereon.

This invention relates to an ignition device for explosive charges, more particularly to a thermal igniter for rocket projectiles.

In rockets containing parachute flares, it is frequently desirable to ignite a delay train leading to an explosive charge for the purpose of ejecting such flare from the rocket at a predetermined time after the departure of the rocket from the projector.

The acceleration forces of rockets are comparatively small so that arming and ignition devices actuated by such forces to effect discharge of the flare would ordinarily have to be made extremely sensitive, which would render such devices unsafe. However, considerable heat is developed within the motor housing of such rockets and it is the purpose of this invention to utilize such heat to ignite a powder train thru a thermal delay element for the purpose of igniting a secondary powder charge for the ejection of the flare from the rocket projectile.

Accordingly, it is an object of this invention to provide an igniter for a rocket carried charge actuated by the heat of the propulsion gases of the rocket projectile.

A further object of this invention is to provide a thermal igniter for the pay load of a rocket projectile which is energized by the heat of the propulsion gases of the rocket projectile after a fixed time delay. It is a particular object of this invention to provide such an igniter structure of simple and readily manufactured design.

The specific nature of the invention as well as other objects and advantages thereof will clearly appear from a description of a preferred embodiment as shown in the accompanying drawing in which:

Fig. l is a longitudinal sectional view of a portion of a rocket projectile embodying this invention.

Fig. 2 is a cross sectional view of the rocket projectile taken along the planeZ-Z of Fig. 1.

Fig. 3 is a cross sectional view of the rocket projectile taken along the plane 3-3 of Fig. 1.

The igniter to be presently described in detail comprises mainly a hollow externally threaded thermal element filled with a readily ignitable material such as black powder. The thermal element'is mounted on an adapter within the rocket projectile in such fashion that one end of the thermal element projects into the propulsion charge chamber. The propulsion charge primer preferably surrounds the end of the thermal element which projects into the propulsion charge chamber. The other end of the thermal element projects into the pay load chamber and is arranged to transmit an igniting flame to the pay load which, for example, may comprise a flare and a flare ejection charge. Thus when the rocket projectile is discharged, the heat generated by the propulsion gases is conducted thru the thermal element to ignite a powder charge within such element after a fixed time delay, which in turn ignites the flare ejection charge thus effecting ejection of the flare from the rocket body and its ignition. The delay introduced is of course designed to produce ignition of the pay load, such as the flare, at a desired point in the flight of the projectile. If more delay is desirable, a powder train delay element may be arranged between the thermal element and the pay load.

It is desired to point out that the thermal igni tion to be described herein, has numerous other obvious applications. The particular application of the thermal igniter to a rocket propelled para chute flare is merely one specific example of its application.

There is shown in Fig. 1 in assembled relation a portion of a rocket projectile embodying this invention. A rocket motor housing I is provided to house the propulsion elements of the rocket. An adapter 2 is inserted in one end of housing 1 and is secured thereto as by threads 3. The for ward end 4 of adapter 2 is of reduced diameter to be insertable in one end of a parachute flare housing 5 and is secured therein by screws 6. The forward end of flare housing 5 connects with an ogival head portion (not shown). A suitable cowling l surrounds housing 5 adjacent the shoulder formed by the reduced end 4 of adapter 2.

An axial cylindrical recess 8 is provided in the rear end of adapter 2. Within recess 8 a plurality of rods 9 having threaded ends it are screwed into concentrically spaced threaded holes cient. It is of course necessary that thermal element [3 has sufficient strength and rigidity to withstand the relatively high pressure produced by the propulsion gases inside the motor housing i. A thermal element having smooth exterior side walls thin enough to permit sufficient conduction of heat to the hollow interior i i to ignite a black powder charge therein has been found to collapse under the pressure developed inside of motor housing i. smooth side walls of sufficient thickness to withstand the internal pressure, the thickness of such walls would prevent the transmissionof suificient heat during the relatively short burning time of the propulsion charge to ignite a black powder charge in the hole I4. By the use of a threaded thermal element i3 sufficient strength and rigidity is provided while at the. same time such thread, which in effect is a helical fin, will rapidly conduct heat to the interior it. A primer l6 comprising a small sack of black powder and an electric squib i8 is located in the bottom of recess 8 of adapter 2 and surrounds thermal element i 3. Squib i6 thus can ignite propulsion charge H andis actuated by any conventional electric firing mechanism of a rocket projector (not shown).

Within flare housing a ring spacer ii is inserted and located adjacent the forward face of adapter 2. A plurality of vent holes it thru botl'i housing 5 and ring I T are provided about the periphery of flare housing 5. A disc is abuts the end of ring spacer H. An axial hole 26 is provided in disc [9 in which is inserted a small rearwardly projecting tube 2!. Tube 2i is provided with an intergral rim 22 at its forward end which abuts the face of disc i9.

Tube 2i comprises adelay element and is filled with a slow burning, sulfurless powder 23. As the sulfurless powder 23 is difi icult to ignite a cap 24 of readily ignitable black powder is mounted on the'end of tube 2i which is adjacent thermal element 13. The black powder cap is then readily ignited by the charge iii in thermal element 19, as will be presently described. A 132W: ejection charge 25 is placed adjacent disc is and is retained there by a disc 21 which is separated from disc ill by a spacer ring A plurality of gas escapeholes 28 are provided in the disc 2? to permit the gases from charge '25 to escape thru such holes when charge 25 is ignited to impinge on the base of a conventional parachute flare 29 mounted forward of disc 2i within flare housing 5.

When primer I6 is discharged by squib 16' the propulsion charge II is in turn ignited and considerable heat and pressure is immediately generated by the resulting combustion of such charge. The gases formed launch the rocket projectile from its projector in conventional manner. In the meantime heat from the igniter and the combustion of the propulsion charge is rapidly conducted thru the thermal element 13 to the hollow interior l4 thereof. When the temperature within the hollow portion I4 reaches approximately 550, which is the combustion temperature of the preferred black powder igniting charge, the charge I5 is ignited. It is obvious that for reasons of safety, that charge it should not completelyburn until the projectile is Well out of the projector. The length of time it takes to ignite and burn charge 15 is of course governed by the length and thickness of the thermal element l3. By proper selection of these dimensions the time delay may thereby be increasedor de- Conversely, if thermal element l3 had I i ing powder 23.

. in the order of eight seconds.

creased. In this particular application the length of time of heat transfer to the hollow portion i4 is on the order of one tenth of a second. The rocket will thus be well beyond the muzzle of the projector when such ignition occurs.

The black powder cap 24 mounted on the end of tube 2! is ignited shortly after the ignition of charge l5. Cap 2 in turn ignites the slow burn- The time for complete combustion of the powder in tube 2i may of course be readily regulated by proportioning of the amount of such powder within tube 2 i. In this particular application the length of time required to com pletely burn the powder train within tube 2i is This permits the rocket projectile to reach the optimum height for the release of the parachute flare 29 at which time the flare ejection charge 25 will be ignited.

Upon ignition of such charge the gas blast from the resulting explosion escapes thru vent holes 28 and impinges on the base of flare 29. The flare 29 is thus forcibly propelled out of the end of flare housing 5, blowing on" whatever ogival head member (not shown) which may be provided on the forward end of housing 5.

From the foregoing description it is readily apparent that a safe, easily and simply constructed thermal igniter for the delayed ignition of the pay load of a rocket projectile is hereby provided. The structure represents an optimum combination of simplicity, ruggedness and efficient, reliable performance.

I claim:

1. In combination, a first chamber, a second chamber, common wall between said chambers,

said wall having an opening therethrough, a hollow tub-uiar exteriorly threaded member threaded into said opening in the wall thereby sealing said opening, said tubular member having one end projecting into said first chamber and being threaded for a substantial portion of said projecting end, the hollow portion of said tubular member opening into said second chamber and arranged to secure an ignitable charge therein whereby a high teir erature condition in said first el' ambor-wili orot delayed ignition 0i: said charge by h t. ansmission thru the threaded portion of said tubular member.

2. In a rocket projectile, a motor chamber arranged to contain a propellant charge, a load chamber arranged to receive combustible load, a wall separating said chambers, an igniting charge for saidcombustible load, and thermal delay means mounted. in said wall arranged to conduct heat from said motor chamber to said igniting charge to ignite said igniting charge said delay means comprising a hollow tubular member having a helical reinforcing and heat transmitting fin on the surface thereof.

3. In a rocket projectile, a motor chamber arranged to contain a propellant charge, a load chamber arranged to receive combustible load, a wall separating said chambers, said wall having a threaded hole therethrough, a hollow tubular member threaded into said. hole, thereby sealing the hole, said tubular member having one solid end projecting into said motor chamber and a helical reinforcing and heat-conducting fin formed on the exterior thereof, the hollow portion of said tubular member connecting withv said load chamber, and an igniting charge in said hollow portion ignitable by heat conducted from said motor chamber thru the walls of said tubular member.

4. In a rocket projectile, a motor chamber, ar-

ranged to contain a propellant charge, a load chamber arranged to receive combustible load, a wall separating said chambers, said wall having a threaded holetherethrough, a hollow tubular exteriorly threaded member threaded into said hole thereby sealing the hole, said tubular mem ber having one solid end projecting into said motor chamber, an igniter for the propellant charge mounted on said projecting end of said tubular member, the hollow portion of said tubular member connecting with said load chamber, and an igniting charge in said hollow portion ignitable by heat conducted from said motor chamber thru the walls of said tubular member.

5. In a rocket projectile, a motor chamber arranged to contain a propellant charge, a load chamber arranged to receive combustible load, a wall separating said chambers, said Wall having a threaded hole therethrough, a hollow tubular exteriorly threaded member threaded into said hole and thereby sealing the hole, said tubular member having one solid end projecting into said motor chamber, the hollow portion of said tubular member connecting with said load chamber, an igniting charge in said hollow portion ignitable by heat conducted from said motor chamber thru the walls of said tubular member, and a powder train delay element mounted in said load chamber, one end of said element arranged adjacent the hollow portion of said tubular element, the other end arranged adjacent the combustible load.

CLARENCE N. HICKMAN.

REFERENCES CITED The following references are of record in the file of this patent:

UNITED STATES PATENTS Number Name Date 1,326,494 Gowdy Dec. 30, 1919 2,093,353 Geitmann Sept. 14, 193'! FOREIGN PATENTS Number Country Date 669,007 Germany Dec. 14, 1938