US3653288A

US3653288A - Tubular-shaped launcher for projectiles, in particular for missiles

Info

Publication number: US3653288A
Application number: US435785A
Authority: US
Inventors: Emile Stauff; Jean Guillet; Pierre Allard; Johannes Schubert; Heinz Topfer; Erich Prier
Original assignee: NORD AVIAT SOC NATIONALE DE CO
Current assignee: NORD AVIAT SOC NATIONALE DE CO; NORD-AVIATION SOC NATIONALE DE CONSTRUCTIONS AERONAUTIQUES
Priority date: 1964-02-26
Filing date: 1965-02-25
Publication date: 1972-04-04
Anticipated expiration: 1989-04-04
Also published as: DE1428637C1; FR1604112A; GB1222501A

Abstract

The present invention relates to a tubular-shaped launching device for projectiles, in particular for missiles.

Description

United States Patent 1151 3,653,288

Staufi et a1. [4 Apr. 4, 1972 [54] TUBULAR-SHAPED LAUNCHER FOR PROJECTILES, IN PARTICULAR FOR 1 References t d MISSILES UNITED STATES PATENTS [72] Invent: Emile Versailles; Jean Guam" 1 380 35s 6/1921 Cooke ..s9/1.7 A

Chatenay-Malabry; Pierre Allard, Fontenay-Sous-Bois, all of France; Johannes Schubert, Putzbrunn, near Munich; Heinz 2,421,522 6/1947 Pope 2,598,256 5/1952 Hickman 2 834 255 5/1958 Musser... Tf,Mh-P1'h;EhP, fi f a" 5 i m 2,967,460 1/1961 Mussel... ..s9/1.7 A 2,986,973 6/1961 Waxman. ..89/1.703 Assigneer Noni-Aviation Societe Nationale De 2,987,965 6/1961 Musser ..s9/1.7 A

a or struction Aeronautiques, Paris, France a 9 a [22] Filed: Feb. 25, 1965 FOREIGN PATENTS 0R APPLICATIONS [21] pp 435785 460,353 10/1949 Canada ..89/1.7A

Primary Examiner-Samuel W. Engle [30] Foreign Application Priority Data Att0rneyl(ar1W.F10CkS Feb. 26, 1964 France .965 258 [5 ABSTRACT [52] [1.8. CI ..89/1.703, 89/1.704 Th ent invention relates to a tubular-shaped launching [51 Int. Cl .1 41 de i e for roje tiles in particular for mi55i1e5 [58] Field ofSearch ..89/1, 1.7 A, 1.7 B, 1.703,

89/ 1 .704 9 Claims, 5 Drawing Figures lullllllllllllgmlllll TUBULAR-SHAPED LAUNCHER FOR PROJECTILES, IN PARTICULAR FOR MISSILES The present invention relates to a tubular-shaped launching device for projectiles, in particular for missiles.

Devices of this kind are already known which operate in accordance with the principle of the single chamber or of the double chamber. According to the single chamber conception, the propellant charge is disposed between the projectile and a rear closing wall which, as the propellant charge is consumed, is pushed rearwards, thus ensuring practically recoilless firing.

The single chamber structure also has the advantage of simplicity of construction and of operation.

However, a drawback has manifested itself: the field of use of this single chamber design is considerably limited by the minimum pressure which is necessary for a stable combustion of the propellant powder. As a result of the definite limits in the ignition conditions of the powder, strong accelerations are produced of such a nature that the operation of a firing device embodying the principle of the single chamber, for example for the launching of bodies provided with measuring instruments or equipments which are sensitive to accelerations, is no longer possible.

An improvement on the single chamber design is achieved by the use of a launcher constructed under the principle of the double chamber and comprising a high-pressure combustion chamber in which is disposed the propellant charge, as well as a low-pressure chamber connected to the first chamber and comprising an expansion nozzle. This design has the advantage that all the parameters, such as acceleration of the projectile, acceleration path, discharge speed, etc. can be altered within wide limits irrespective of the properties of the propellant powder. The shape and size of the projectile or of the missile launched, as well as the possible existence of control equipment or directional equipment which are sensitive to accelerations do not limit the field of use of this double chamber concept. However, it has been found that launchers operating under the double chamber principle have an unfavorable volume for easy handling of the weapon, either because the diameter of the tube is too large or its length is too long. These two drawbacks arise from the manner in which the propellant charge is disposed and from the shape of the expansion nozzle which is usually designed along the lines of a de Laval nozzle.

The object of the present invention is a tubular-shaped launcher of the type hereinbefore described and exhibiting good possibilities of adaptation to the highest discharge speeds while having the lowest volume and weight compatible with the requirement of absence of recoil.

The object of the invention is achieved by means of a highpressure combustion chamber fixedly secured within the launching tube and provided with means for receiving solid propergol and at its front portion with apertures for the passage of combustion gases flowing into a low-pressure combustion chamber, the wall of the high-pressure combustion chamber forming with the launching tube a supersonic annular nozzle through which the combustion gases flow.

In contradistinction to known launching devices, the device constructed in accordance with the above principle is characterized not only by a substantial absence of recoil due to the annular nozzle design but has the advantage of a reduced length together with a constant tube diameter over the whole of its length. The reduced length is due to the fact that in launches of the known type the total length results from the length of the nozzle of the combustion chamber and from the length of the space reserved for the propergol, and to the fact that an annular nozzle is much shorter in length than a de Laval nozzle.

To ensure an absence of recoil of the launcher, it is not necessary that the forces be in balance at each moment of the combustion. It suffices that the sum of the impulses as a whole be zero during the very short period of combustion. In case the thrust through the supersonic nozzle which is necessary to balance the quantities of motion, is no longer sufficient per se,

according to a unique feature of the invention, the high-pressure chamber has, at its rear portion, openings in the form of reaction propelling nozzles, which provide an additional thrust. This arrangement has, furthermore, the advantage of maintaining a small caliber for a given thrust and pressure within the combustion chamber.

Should the launcher be provided with spin-imparting devices such as guides for example, or any other rotation-imparting means, with a view to stabilizing the projectile or missile, the angular acceleration of the projectile produces a moment of rotation during the firing phase, which acts on the firing installations. On recoil-less firing devices it is however desirable to avoid applying to the tube, during the firing phase, not only forces but also couples. To this end, according to the invention, there is provided in the annular slot of the supersonic noule a device which produces an antagonistic moment.

In the firing devices according to the invention, making use of the double chamber principle involving the production of a thrust in the high-pressure chamber and constant sections of flow, the impulses are practically negligible during the whole of the combustion. However, it is not impossible that under certain conditions, with constant sections of flow, the recoil forces are larger than the counter-thrust of the nozzles at the start of combustion, whereas, at the end of combustion, the counter-thrust of the nozzles is preponderant. In this case, the equipment is subjected to axial forces of variable magnitude and direction. To avoid the production of these forces, ac-

cording to the present invention, the openings provided in the" front portion of the high-pressure combustion chamber comprise fusible or combustible gaskets whereby the flow section from the high-pressure combustion chamber can be increased during the combustion period.

Another solution for avoiding residual impulses consists in providing in the front portion of the high-pressure combustion chamber openings produced by the retarded combustion of certain substances.

In this manner, with either of these solutions, it is possible to adapt the gas flow to the increase of the tube volume resulting from the motion of the projectile or of the missile within the tube, and to reduce to an appreciable extent the variations of pressure in the low-pressure chamber, as well as the nonbalanced residual forces.

A form of embodiment of the launching device according to the invention is shown in the figures of the accompanying drawing. All portions of the device which have no direct connection with the invention have not been shown in detail, for the sake of clarity.

In said figures:

FIG. 1 is a diagrammatical longitudinal section of the extreme end of a launcher;

FIG. 2 is an end view of the launching tube as seen in the direction of the arrow A of FIG. 1;

FIG. 3 is a partial view in longitudinal section of the rear part of the high-pressure combustion chamber with additional outlet orifices;

FIG. 4 is a view of the forward portion of the combustion chamber provided with gaskets which produce a variation in the section of the orifices for the passage of the gases;

FIG. 5 is a developed view of the additional gaskets closing the orifices for the passage of the gases in the forward portion of the combustion chamber.

FIG. 1 illustrates a tube 1 into which a partially shown missile 2 has been introduced. This missile is provided with foldable surfaces 2a for stabilization purposes as is well known in the art. With a view to stabilizing the projectile or part of the projectile by a spin effect, the inner wall of the launching tube comprises guides 1a one of which only is shown in the drawing. A piston 3 is disposed at the rear portion of the missile 2, said piston having on its outer periphery rings 3a and 3b to ensure a tight seal of the low-pressure chamber in relation to the space into which the missile is introduced.

Corresponding to the number of guides, the piston is provided on its periphery with recesses 30 which is engaged by a block 4 sliding in one of said guides. In view of the mechanical connection of the piston 3 with the guides 1a referred to above, on the one hand, and of the dynamical connection of said piston with the missile, on the other hand, for example through the medium of the shoulder, 3d, as the volume of gas increases the missile is ejected from the tube with a spin.

In accordance with the invention, the gas (and therefore the thrust) is produced in a high-pressure chamber 6 fixedly secured to the tube 1 by a ring provided with arms 5a; the propellant charge, composed of several disks, is mounted within the high-pressure combustion chamber on a tube 8 and is secured against axial movements. Rearwardly, the highpressure chamber 6 is hermetically sealed by a disk 9 on which is also secured the tube 8. At the front of the combustion chamber gas discharge openings 6a issue into the intermediary or low-pressure chamber 10 arranged rearwardly of the piston 3.

According to the invention, the high-pressure combustion chamber 6 is so designed and arranged that its wall 6b, in its central portion, forms with the launching tube 1 a supersonic annular nozzle 11 through which the combustion gases flow, whereby the advantages described at the beginning of the present specification can be obtained.

The firing of the propellant charge 7 is carried out in the illustrated form of embodiment by electrical means (ignition wires 12) which cause an igniting device (not described in detail) to detonate. This igniting device is disposed within the tube 8 and ignites the propellant charge through apertures (not shown) formed in the tube 8.

In case the caliber must be maintained as low as possible, for a predetermined thrust and chamber pressure, or in case the balancing of the quantities of motion obtained by the thrust produced by the supersonic annular nozzle is not sufficient per se, reactive thrust nozzles 13 can be provided at the rear end of the chamber 6, on the end disk 9, in addition to the forward orifices 6a for the flow of gases. These reactive thrust nozzles 13 can be designed in the form of de Laval nozzles or of annular nozzles. An arrangement of this type is shown in FIG. 3.

In order to compensate the moment of rotation resulting from the torsional couple produced at the start of the missile, and to maintain, as far as possible, the whole launching installation in equilibrium as to moments, the radial fixing arms 5a of the high-pressure combustion chamber are disposed obliquely in relation to the direction of incident flow (FIG. 2). In this manner, a tangential component is obtained which is added to the thrust produced by the ejection of the gases discharged, said component giving rise to a couple which is antagonistic to the moment of rotation of the launching tube.

According to FIG. 4, with a view to avoiding axial stresses of variable magnitudes and directions which are liable to be produced on the launching tube in certain cases at the beginning and at the end of the combustion, the front orifices 6a of the high-pressure combustion chamber are provided with fusible or combustible gaskets 14 which enable the section of the orifices 6a, which are small at the onset of combustion, to be increased during the combustion phase, proportionally to the increase of the volume in the launching tube during the launching operation.

With a view to avoiding these residual forces another form of embodiment is shown in FIG. 5 wherein the front portion of the wall of the combustion chamber 6b is provided with apertures 6a shown in developed form. As shown in this figure, a plurality of apertures 6a are completely closed by gaskets 15 at the start of combustion. During the launching operation, a number of additional apertures are formed due to the combustion or the fusion of said gaskets.

The invention is not limited to the embodiments according to FIGS. 4 or 5; on the contrary it is possible to combine both been described and illustrated in an exglanatory manner without any intention to limit the same, an that alterations 0 detail can be made thereto within the spirit of the invention, without falling outside its scope.

What is claimed is:

l. A tubular launching installation for projectiles comprising a launching tube having a portion of substantially constant inner diameter,

a tubular means forming a high-pressure combustion chamber disposed within said tube and spaced therefrom and containing a propellant charge,

said tubular means having forward and rearward closure means and peripheral gas outlet openings adjacent sai forward closure means, a

a supersonic annular nozzle formed by a portion of the outer surface of said tubular means in cooperation with a portion of the inner surface of said launching tube at a region of smallest cross-sectional area between said portions of said surfaces located at the rearward end of said portion of substantially constant inner diameter of said launching tube,

means for supporting said tubular means in said launching tube located in the region of said supersonic annular nozzle,

a piston disposed in said launching tube behind the projectile,

and a low-pressure chamber located behind said piston and in front of said region of smallest cross-sectional area of said supersonic nozzle and concentric to said tubular means.

2. The tubular launching installation of claim 1 further characterized by said rearward closure means having additional openings in the form of reaction thrust nozzles.

3. The tubular launching installation of claim 2, further characterized by guide means on the inner surface of said launching tube for imparting a spin to the projectile,

and said means for supporting said tubular means including gas deflecting means for producing an antagonistic moment to said spin.

4. The tubular launching installation of claim 1, further characterized by gaskets in said peripheral gas outlet openings adapted to vanish due to the heat of combustion during the combustion of the propellant charge.

5. The tubular launching installation of claim 4, further characterized by said gaskets being fusible.

6. The tubular launching installation of claim 4, further characterized by said gaskets being combustible.

7. The tubular launching installation of claim 1, further characterized by supplemental peripheral gas outlet openings adjacent said forward closure means having gaskets adapted to vanish due to the heat of combustion during the combustion of the propellant charge.

8. The tubular launching installation of claim 7, further characterized by said gaskets being fusible.

9. The tubular launching installation of claim 7, further characterized by said gaskets being combustible.

PatentNo. 3,653,288 Dated April 4, 1972 n (s) EMILE STAUFF, JEAN GUILLOT,PIERRE ALLARD,

JOHANNES SCHUBERT, HEINZ TOPFER, AND ERICH PRIER It is certified that error appears in the above-identified patent and that said Letters Patent are hereby corrected as shown below:

In column 1, lines 10 and ll, the names of the assignees should read: Nerd-Aviation Societe Nationale de Constructions Aeronautiques Paris, France and Bolkow Gesellschaft mit beschrankter Haftung Munich, West Germany Signed and sealed this 3rd day of October 1972,

(SEAL) Attest:

EDWARD MQFLETCHER J Attesting Officer ROBERT GOTTSCHALK Commissionerof Patents FORM FO-105O (1069) USCOMM'DC 60376-F59 9 U.S, GOVERNMENT PRINTING OFFICE: 1969 0-366-334 UNITED STATES PATENT OFFICE CERTIFICATE 0F CORRECTION Patent N v Dated April 4,

Inventor(s) EMILE STAUFF, JEAN GUILLOT,PIERRE ALLARD,

In column 1, lines 10 and I 11, the names of the assignees should read: Nord-Aviation Societe Nationale de Constrnctions Aeronautiques Paris, France and Bolkow Gesellschaft mit beschrankter Haftung Munich, West Germany Signed and sealed this 3rd day of October 1972.

(SEAL) Attest:

EDWARD M.FLETCHER,JR.

testing Officer ROBERT GOTTSCHALK Commissioner of Patent:

, FORM PO-IOSO (10- USCOMM-DC 60376-P69 9 U.S. GOVERNMENI' PRINTING OFFICE: \969 0-356-334

Claims

1. A tubular launching installation for projectiles comprising a launching tube having a portion of substantially constant inner diameter, a tubular means forming a high-pressure combustion chamber disposed within said tube and spaced therefrom and containing a propellant charge, said tubular means having forward and rearward closure means and peripheral gas outlet openings adjacent said forward closure means, a supersonic annular nozzle formed by a portion of the outer surface of said tubular means in cooperation with a portion of the inner surface of said launching tube at a region of smallest cross-sectional area between said portions of said surfaces located at the rearward end of said portion of substantially constant inner diameter of said launching tube, means for supporting said tubular means in said launching tube located in the region of said supersonic annular nozzle, a piston disposed in said launching tube behind the projectile, and a low-pressure chamber located behind said piston and in front of said region of smallest cross-sectional area of said supersonic nozzle and concentric to said tubular means.

3. The tubular launching installation of claim 2, further characterized by guide means on the inner surface of said launching tube for imparting a spin to the projectile, and said means for supporting said tubular means including gas deflecting means for producing an antagonistic moment to said spin.