GB1586109A - Solid propellant rocket propulsion means for accelerating a projectile along a launching tube - Google Patents

Description

(54) A SOLID PROPELLANT ROCKET PROPULSION MEANS FOR ACCELERATING A PROJECTILE ALONG A LAUNCHING TUBE (71) We,MESSERSCHMITT-BÖLKOW- BLOHM Gesellschaft mit beschränkter Haftung, of 8000 Munchen, German Federal Republic, a Company organised and existing under the laws of the German Federal Republic, do hereby declare the invention for which we pray that a Patent may be granted to us, and the method by which it is to be performed, to be particularly described in and by the following statement: This invention relates to a solid propellant rocket propulsion means for the acceleration of a payload along a launching tube using the expanding gases produced on ignition of the propellant.

According to this invention there is provided a solid propellant rocket propulsion means for accelerating a projectile or the like payload along a launching tube, comprising a casing housing the propellant and forming a combustion chamber with closed end walls, a plurality of orifices through the casing side wall spread radially over said wallsurface and through which gas may pass from the propellant during burning, an end wall of the casing being arranged to define in conjunction with an integral or separate outer tube an annular thrust nozzle concentric with the casing axis, the outer surface of the casing with said outer tube defining a channel in communication with said nozzle, the other end wall of the casing being integral with or sealingly engaging said outer tube, the propellant comprising a plurality of discs concentric with the casing axis and carried on a support extending along the axis.

The outer tube referred to may be the launching tube itself into which the means is inserted for use, or a tube integral with the casing may be provided.

Thus, the invention provides a solid propellant rocket propulsion means with radially distributed outlets for use with known short-time burning propellants, using a very high thrust cross-sectional area.

The advantages offered by the invention are as follows.

Proper design of the combustion chamber and a careful selection of flow parameters will give an opportunity for the hot gases produced by the propellant to flow radially from the zones of burning over the entire length of the propellant, thereby avoiding choking. The discharge from the individual orifices change direction and expand in the channel to form a gas flow towards the convergent/divergent thrust nozzle.

With selection of design parameters the losses can be smaller than in a comparable axial flow of gases discharged from a combustion chamber directly into a thrust nozzle. In this way it is possible to obtain higher thermal and dynamic loads and, consequently, a considerable shortening of the propellant-charge length. The radia] outlets lead generally to a more constant and higher flow velocity.

A number of spaced discs of solid propellant, in the form of circular plates of constant thickness, with the gaps between them leading to the orifices in the casing side wall may be provided mounted on a central support tube. This tube may carry an igniter charge and has radially distributed outlets for ignition located at the gaps between the propellant discs. This arrangement holds the discs parallel and thereby maintains the gaps and also provides a controlled discharge of the combustion products through the orifices.

The casing may taper conically to the nozzle to create an expanding flow channel with the outer tube being of constant diameter or cross section. This radial discharge through the orifices provides a significantly higher flow velocity than in the case of an axial discharge from combustion chamber directly into a thrust nozzle.

An embodiment according to the invention is described in conjunction with the accompanying drawing.

The drawing shows the rear end, in crosssection, of a launching or projector tube 1 including a short burning time rocket propulsion device 2 with a solid propellant in a combustion chamber 3 defined by a closed front end wall 4, closed rear end wall 5 and a conically tapered casing 6 between the walls 4 and 5. The wall 4 locates in tube 1 with a small clearance and the outer circumference of same has an annular groove 7 for a gasket 8. The casing 6 has a plurality of gas outlet orifices 9 over the whole surface. A profiled insert body 10 extends integrally from the rear wall 5 and forms the central part of an annular shaped convergent/divergent thrust nozzle 11. A channel 12 leads from the casing 6 to the nozzle 11.The outer boundary of the thrust nozzle 11 and channel 12 are both defined by either, (a) as shown in the upper half of the figure, part la of tube 1, or (b) a tube 13 as shown in the lower half of the figure and integral with end wall 4. Within space 14 of the combustion chamber casing 3 is an assembly of spaced juxtaposed discs 18 of a solid propellant charge with the radially extending spaces 16 aligned with orifices 17. The discs are mounted on central support tube 19 and additional support is provided for the propellant assembly by rods 20 with spacing washers 21 between adjacent propellant discs.

Support tube 19, which serves as an anchor between end walls 4 and 5, and which gives same added rigidity and strength, serves also as a chamber for the igniter charge 22. Through the inside of the tube 19 runs the electrical ignition leads 24 for the igniter charge 22. The tube 19 has apertures 25 through which igniter gases propagate through gaps 17 between the propellant charge discs 18.

This feature with the short radial distance through which the gases flow from the combustion space 14 to the channel 12 makes possible a rapid build-up of the full thrust and more even burning than in the case of axial discharge from a combustion chamber.

Because of the reduced (20 to 30%) combustion chamber pressures, pressure differences between combustion chamber 14 the flow channel 12 are such that casing 6 of the combustion chamber structure 3 can be made correspondingly thin.

WHAT WE CLAIM IS: 1. A solid propellant rocket propulsion means for accelerating a projectile or the like payload along a launching tube, comprising a casing housing the propellant and forming a combustion chamber with closed end walls, a plurality of orifices through the casing side wall spread radially over said wall surface and through which gas may pass from the propellant during burning, an end wall of the casing being arranged to define in conjunction with an integral or separate outer tube an annular thrust nozzle concentric with the casing axis, the outer surface of the casing with said outer tube defining a channel in communication with said nozzle, the other end wall of the casing being integral with or sealingly engaging said outer tube, the propellant comprising a plurality of discs concentric with the casing axis and carried on a support extending along the axis.

2. A means in accordance with Claim 1, wherein the propellant discs are in spaced juxtaposed relationship, the orifices being provided between adjacent discs.

3. A means in accordance with Claim 2, wherein the support includes an igniter charge with ignition orifices therethrough and positioned between adjacent discs.

4. A means in accordance with any preceding Claim, wherein the periphery of the propellant is spaced from the inner surface of the casing.

5. A means in accordance with any preceding Claim, wherein the channel tapers conically outwardly towards the nozzle.

6. A means in accordance with any preceding Claim, wherein the outer tube is separate and is provided by a launching tube in which the means is inserted, one end wall of the combustion chamber having a sealing means engaging the inner surface of said tube.

7. A means in accordance with any preceding Claim 1 to 5, wherein the outer tube is integral with one end of the casing and extends along the casing concentric therewith.

8. A means in accordance with Claims 6 or 7, wherein the outer tube is cylindrical and of constant cross-section along its length.

9. A solid propellant propulsion means in accordance with any preceding Claims substantially as herein described with reference to and as shown in the accompanying drawing.

10. The means of Claim 10 included within a projectile launching tube.

**WARNING** end of DESC field may overlap start of CLMS **.

Claims (10)

**WARNING** start of CLMS field may overlap end of DESC **. The drawing shows the rear end, in crosssection, of a launching or projector tube 1 including a short burning time rocket propulsion device 2 with a solid propellant in a combustion chamber 3 defined by a closed front end wall 4, closed rear end wall 5 and a conically tapered casing 6 between the walls 4 and 5. The wall 4 locates in tube 1 with a small clearance and the outer circumference of same has an annular groove 7 for a gasket 8. The casing 6 has a plurality of gas outlet orifices 9 over the whole surface. A profiled insert body 10 extends integrally from the rear wall 5 and forms the central part of an annular shaped convergent/divergent thrust nozzle 11. A channel 12 leads from the casing 6 to the nozzle 11.The outer boundary of the thrust nozzle 11 and channel 12 are both defined by either, (a) as shown in the upper half of the figure, part la of tube 1, or (b) a tube 13 as shown in the lower half of the figure and integral with end wall 4. Within space 14 of the combustion chamber casing 3 is an assembly of spaced juxtaposed discs 18 of a solid propellant charge with the radially extending spaces 16 aligned with orifices 17. The discs are mounted on central support tube 19 and additional support is provided for the propellant assembly by rods 20 with spacing washers 21 between adjacent propellant discs. Support tube 19, which serves as an anchor between end walls 4 and 5, and which gives same added rigidity and strength, serves also as a chamber for the igniter charge 22. Through the inside of the tube 19 runs the electrical ignition leads 24 for the igniter charge 22. The tube 19 has apertures 25 through which igniter gases propagate through gaps 17 between the propellant charge discs 18. This feature with the short radial distance through which the gases flow from the combustion space 14 to the channel 12 makes possible a rapid build-up of the full thrust and more even burning than in the case of axial discharge from a combustion chamber. Because of the reduced (20 to 30%) combustion chamber pressures, pressure differences between combustion chamber 14 the flow channel 12 are such that casing 6 of the combustion chamber structure 3 can be made correspondingly thin. WHAT WE CLAIM IS:

1. A solid propellant rocket propulsion means for accelerating a projectile or the like payload along a launching tube, comprising a casing housing the propellant and forming a combustion chamber with closed end walls, a plurality of orifices through the casing side wall spread radially over said wall surface and through which gas may pass from the propellant during burning, an end wall of the casing being arranged to define in conjunction with an integral or separate outer tube an annular thrust nozzle concentric with the casing axis, the outer surface of the casing with said outer tube defining a channel in communication with said nozzle, the other end wall of the casing being integral with or sealingly engaging said outer tube, the propellant comprising a plurality of discs concentric with the casing axis and carried on a support extending along the axis.

2. A means in accordance with Claim 1, wherein the propellant discs are in spaced juxtaposed relationship, the orifices being provided between adjacent discs.

3. A means in accordance with Claim 2, wherein the support includes an igniter charge with ignition orifices therethrough and positioned between adjacent discs.

4. A means in accordance with any preceding Claim, wherein the periphery of the propellant is spaced from the inner surface of the casing.

5. A means in accordance with any preceding Claim, wherein the channel tapers conically outwardly towards the nozzle.

6. A means in accordance with any preceding Claim, wherein the outer tube is separate and is provided by a launching tube in which the means is inserted, one end wall of the combustion chamber having a sealing means engaging the inner surface of said tube.

7. A means in accordance with any preceding Claim 1 to 5, wherein the outer tube is integral with one end of the casing and extends along the casing concentric therewith.

8. A means in accordance with Claims 6 or 7, wherein the outer tube is cylindrical and of constant cross-section along its length.

9. A solid propellant propulsion means in accordance with any preceding Claims substantially as herein described with reference to and as shown in the accompanying drawing.

10. The means of Claim 10 included within a projectile launching tube.