US3308624A

US3308624A - Rocket engine with consumable casing

Info

Publication number: US3308624A
Application number: US352486A
Authority: US
Inventors: Ciancitto Antonino
Original assignee: A GIAMBROCONO ING
Current assignee: A GIAMBROCONO ING; ING A GIAMBROCONO
Priority date: 1963-05-03
Filing date: 1964-03-17
Publication date: 1967-03-14
Anticipated expiration: 1984-03-14

Description

March 14, 1967 A. CIANCITTO ROCKET ENGINE WITH CONSUMABLE CASING 4 Sheets-Sheet 1 Filed March 17, 1964 ATTORNEY March 14, 1967 Filed March 17, 1964 A. CIANCITTO 3,308,624

ROCKET ENGINE WITH CONSUMABLE CASING 4 Sheets-Sheet 2 ATTORNEY March 14, 1967 A. CIANCITTO V ROCKET ENGINE WITH CONSUMABLE CASING Filed March 17, 1964 4 Sheets-Sheet 5 j INVENTOR [L's-e :1 r r 3 mm 3. K Ii v 1/ O0 Wk mm H 1 l 1|, in N mu u n u a 1].. n u L 4 3 a ww a n f Q 1 u k u H i: l mv W I qwh. 4 MR om M M i I #0 To mu 3 9 wk a a 8mm Q mv .vw k. i k Q I 1 M Q w k IQ I 1i a will ww l hwl i i m. Wlumw. n H lfl i l l l| ATTORNEY March 14, 1967 A. CiANClTTO 3,308,624

ROCKET ENGINE WITH CONSUMABLE CASING Filed March 17, 1964 4 Sheets-Sheet 4 &

ATTORNEY United States Patent 3,308,624 ROCKET ENGINE WITH CONSUMABLE CASING Antonino Ciaucitto, Ing. A. Giambrocono, Paterno, Catania, Italy Filed Mar. 17, 1964, Ser. No. 352,486 Claims priority, application Italy, May 3, 1963, 695,265/ 63 8 Claims. (Cl. 60-224) This invention relates to a rocket engine with consumble casing, and more precisely to a new propulsion system for the proper missile.

The main object of the invention is the provision of a rocket engine :wherein the consumption of propellant carried by it, is associated with a progressive destruction of the container wherein the propellant is stored, in order to obtain a material increase in the so called mass ratio, which is defined as the ratio of initial weight, to final useful weight of the rocket engine.

A further object of this invention is the provision of a rocket engine wherein the progressive destruction of the container wherein the propellant is contained, results in an added propulsion effect, whereby the container may be considered as a propellant, or more precisely as a combustible material.

The above objects are attained by a rocket engine, which is characterized essentially in that it comprises means designed for the drive of compressors, and tools by which the propellant container is progressively destroyed; means wherein fragments of container material are discharged, along with the gases designed for driving turbines, which in turn drive the compressors and the means for the destruction of the container, and finally a row of nozzles by which the fragments of container material are utilized as propellant, along with the liquid fuel and oxidizer stored in same container.

Any one of the magnesium base alloys (with magnesium as major component) as usually utilized in the air craft industry, may be selected as a material suitable for the container. For illustrative purposes only, the variety of alloys that are known under the commercial denomination of Electron might be utilized; however, even other alloys having again magnesium as main component-may be utilized for the purposes of the invention.

According to a particular embodiment form of the invention, the missile comprises a fore section, wherein the cockpit for the controls and possible passengers is contained, and that must show a streamlined shape in order to reduce the drag. In the case in question the ogive form is the most suitable. Said missile fore section is followed by a double wall shell, made of an alloy having a high magnesium content, in which the oxidizer is stored inside of one of the chambers thereby defined, while the liquid fuel is contained in an outer annular chamber. The lower ends of said chambers are closed by a supporting plate, and a number of motorcompressor sets are alternately fitted all around its periphery. Each one of such motor-compressor sets is fed independently from all others, and comprises a turbine by which both compressors designed for the liquid propellant are driven. A portion of such propellant is utilized for driving the proper turbine, which in turn drives the means for progressively destroying the shells, which fragments fall into a chamber located downstream of the turbine. The latter chamber is annularly shaped, and the propelling nozzles, fitted between two adjacent motorcompressor sets, are connected therewith.

The invention will be better appreciated from a consideration of the following detailed description of a preferred embodiment form thereof, as shown in the ac companying drawings, whereby the description and drawings are given only as a not restrictive example. In the drawings:

FIG. 1 diagrammatically shows a rocket engine according to the invention.

FIG. 2 is a diagrammatic section taken on the line IIII of FIG. 1.

FIG. 3 is an enlarged and more detailedeven if yet diagrammatic, from certain viewpointssection taken on the line TIL-III of FIG. 2, and:

FIG. 4 is a section taken on the line IVIV of FIG. 2.

FIG. 5 is an enlarged cross-section taken inside of circular zones as shown in the FIG. 4.

Referring now to above figures, 1 is the cone-shaped fore section of the missile, that can be realized in any already known manner, and may comprise a cockpit for the possible crew. Such detail shall be no more discussed in the following description, since the invention does not cover it.

Said fore section 1 is connected, by any known means, with a

double shell

2, 3, consisting of tubular bodies, made of an aircraft type magnesium alloys, e.g. electron, and that define an annular outer chamber 4, wherein the liquid fuel, e.g. kerosene, is stored, and an inner cylindric chamber 5, wherein the oxidizer-Le. liquefied oxygen-4s stored. The bottom of both

chambers

4 and 5 is closed by a base plate 6, and double truncated cone shaped propelling jets 7, of known type, are fitted along the outer edge of said base plate 6, and more precisely within the annular ring as defined by the chamber 4. Motor-compressor sets, as indicated in the whole with 8, are also fitted, alternately with said jets, in the same annular area. In more detail: an annular body 10, having downwardly diverging walls, is connected with said base plate, and a row of stationary blades 12, made a metal carbide-e.g. WIDLA or other equivalent material-are carried by its upper flange 11. Said stationary blades 12 are arranged all along two concentric rings, and are located directed below the

shells

2 and 3, the blades 12 are designed to cooperate with two rings 13, having a toothing formed on their sides 14 directed toward the inside of chamber 2. The profile of said toothings is oblique, to make them able to cooperate with the bevel gears 15 (see FIGS. 3 and 4), as explained in more detail later. The two rings 13 (driven by the bevel gears) are formed with flanges 1 6, having rows of cutting teeth made of a metal carbide, e.g. W'IDIA, and designed to remove chips from the

shells

2 and 3.

Such chips are then discharged into an annular chamber 17, defined by the ring 10. The combustion chambers 7, and the motor-compressor sets 8, as shown in detail in the FIG. 4, are supported by said ring.

The double walled, double truncated cone shaped propelling nozzles 7 are of the conventional type. The compressed fuel flows through the intervening space 18, and is injected by jets arranged in a circle 19, into the combustion chamber 7. The annular chamber 17 is .put in communication with the combustion chamber 7 through a port 20, as defined by said jets. Said port can be closed by a shutter 21, which is fitted on the end of a rod 22, fast with a piston 23, that is slidingly fitted, against the action of a spring 25, in a cylindric chamber 24. Said piston 23 can be acted upon by a pressure fluid, fed through a duct 25a from a suitable pressure fluid source (not shown), and controlled by suitable control devices, e.g. pressure regulators, fitted within the cone 1, and hand operated by the crew, or indirectly controlled, e.g. by a programming device. By the closing of port 20, that occurs simultaneously to closing of

propellant feeding valves

82 and 84, the related combustion chamber can be put out of operation. It is thus possible to adjust the trajectory of the missile, and also to materially change the thrust imparted thereto. For such a purpose, the propellant valves are to be conveniently closed, and the gear 74 is to be thrown out of mesh, thus releasing all other mechanisms. A disengageable clutch 79 is therefore provided.

The plate 9 carries a side wall 29 (FIG. 3), that is connected therewith, and is fiitted with an annular packing 30, designed to ensure a seal against the inside face of shell 3. In order that the closing assembly 6 be kept connected with the

shells

2 and 3, even when no thrust is imparted, the side wall 29 is fitted with braking shoes 31, that can be forced, against the action of return springs 33, into engagement with the shell 3, by a pressure fluid that is fed through the duct 32. Said braking shoes may be controlled by hand, or automatically.

Referring now again to the drawing, and more specifically to FIGS. 4 and 5, a description will be given of the means that are designed to feed the propellant into the combustion chamber 7, and to bring about the progressive destruction of

shells

2 and 3. Such means substantially consists of the assemblies 8.

With the aid of a flowrate adjusting valve 84, that can be manually or program controlled, the possibility is given to adjust the fiow of an oxidizer, that is fed to compressor 43 through the duct 42. Said compressor is of the conventional, centrifugal type, and its delivery side is connected with an outlet duct, comprising two

branches

44 and 45, having different diameters. By the first duct 44 the pressurized oxidizer is fed to the different combustion chambers (see FIG. 4) of the propelling nozzles, while the second duct 45 extends annularly all around a body 47. The oxidizer is injected into a precombustion chamber 59, by a row of nozzles 48 arranged in a circle.

The valve 84 is fitted inside of shell 3, and is securd to bottom plate 9, through which passes the duct 42, which is controlled by said valve.

The

ducts

44 and 45 are designed in such a manner that the major portion of oxidizer is fed to the chambers of nozzles 7, while the minor portion thereof is fed to precombustion chamber 50, for driving the rotor 51 of a turbine. The fuel that is fed into the precombustion chamber 50, consists of kerosene, taken from its storage shell, through a flowrate adjusting valve 82, that is fitted in the ring 26, and fed through the duct 60 to a centrifugal type compressor 61. The major portion of thus delivered kerosene is fed to duct 62, which end section 63 extends all around the precombustion chamber 50, and the communication between duct and chamber is established through a crown of nozzles 64. The turbine 51 is then driven by the gases produced by the partial combustion of kerosene, that passes through the rotor blades of the turbine, and is discharged into the annular chamber 17, wherein the chips cut from the

shells

2 and 3 are also discharged.

Such chips are melted by injecting an oxidizer through melting nozzles 81, that are fitted on ends of ducts 81a extending along the wall by which the annular chamber 17 (downstream of turbine) is defined. Such ducts are connected with the main duct 45. Said molten chips are then conveyed, along with the remainder of the propellant, through parts 20 to propelling nozzles 7, wherefrom they are discharged to generate the thrust as required for the motion of missile.

The turbine rotor 51 drives, through the ring gear 70, two gears 71, respectively keyed on the

shafts

72 and 73, whereon the compressor rotors are also keyed. A gear 74, that may be torsionally connected with a shaft 75 by the pressure fluid clutch 79, is driven by the gears 73a, also fitted on said

shafts

72 and 73. The shaft 75 extends through the rotor 51, and is supported by the

bearings

75a, 76 and 77, being thereby independent from said rotor 51; shaft 75 is designed to impart the required rotary cutting motion to crown gears 13, that are in mesh with the pinions 15. The shaft 75 is axially bored as in 77a, to allow a portion of kerosene, that is fed by the compressor 61, to flow therethrough, entering from the channel 78, for cooling purposes. Same kerosene flows then from the end of shaft 75, into the chamber 17a, thus cooling also the counter-revolving rings 13, that are sealed against the annular piston 26 by a labirinth seal 80.

The bearing 77 is fitted in the ring 26, while the bearing 76 is interposed between the rotor and its shaft, and the thrust bearing 75a is fitted in a housing of the body 47, wherein the different gears, the shafts, the compressor rotors and the turbine are also fitted.

The operation of the abovedescribed device is as follows:

The liquid oxidizer, to be fed to motor-compressor sets 8, and to related nozzles 7a, is taken from the chamber 5, through the valve 84 and the ducts 42, and is sent to compressors, wherefrom a major portion thereof is fed to the nozzles 7 through the ducts 44, while the remaining minor portion is fed to precombustion chambers 50, through the ducts 45 and the nozzles 48. The liquid fuel is taken through a duct 60 and the valves 82, and is fed by the compressors 61 through the precombustion chamber 50, through the ducts 62 and the crown of nozzles 64.

Thus, the fuel and the oxidizer are fed into the precombustion chambers, wherein they are partially burnt, the combustion being initially started by anyone suitable device for the ignition of mixture. The combustion gases are discharged from the chamber to the annular cham- 'ber 17, thus driving the turbine, which in turn drives, through the abovedescribed gears, both the

compressors

63, 61, and the shaft 68, which rotary motion results in a reverse rotation of ring cutters 13. The chips that are cut by latter rings from the material of

shells

2 and 3, are discharged into the annulus 17, wherein they are melted by the oxidizer injected through the melting nozzles 81. The only partially burnt fuel mixture is finally fed from the chamber 17, through the ports 20, into the combustion chambers of nozzles 7, wherein the thrust is produced as effect of the complete combustion of fuel components that are mixed therein.

Thus, a gradual destruction of

shells

2 and 3 is ob tained. The material of such shells can be utilized as propellant, due to its large magnesium contents, with the added advantage of a gradual decrease in the weight of the rocket engine.

What I claim is:

1. A rocket engine comprising, in combination, elongated casing means adapted to contain propellants for said engine and having a bottom wall and side wall means movably arranged with respect to said bottom wall; means forming an annular chamber attached to said bottom wall; a plurality of propelling nozzles mounted on said chamber forming means and communicating with the interior of said annular chamber formed thereby; prolellant pumping means mounted on said chamber forming means alternately with said nozzles for pumping propellant from said easing into said annular chamber; prime mover means for driving said pumping means; and cutter means driven by said prime mover means for gradually destroying said side wall means as the propellant is used up.

2. A rocket engine as set forth in claim 1, wherein said side wall means of said casing means comprise an inner and an outer shell arranged spaced from and substantially coaxial with each other and forming an annular outer chamber and an inner cylindrical chamber, said bottom wall closing said chambers at one end thereof.

3. A rocket engine as set forth in claim 2 including a plurality of cylinder means fixed to said bottom wall and fluid-operated piston means in said cylinder means and engaging the inner surface of said inner shell.

4. A rocket engine ase set forth in claim 1, wherein said prime mover means comprises a turbine having a delivery side communicating with said annular chamber connected to said bottom wall.

5. A rocket engine as set forth in claim 1, wherein said side Wall means are formed from combustible material.

6. A rocket engine as set forth in claim 5, wherein said side wall means are formed from a material having a high magnesium content.

7. A rocket engine as set forth in claim 5, wherein said cutter means and said annular chamber are arranged so that chips cut from said side wall means by said cutter means will drop into said annular chamber and so pass from there into said nozzles.

8. In a rocket engine, in combination, elongated easing means adapted to contain a propellant for said engine, said casing means having 'a bottom wall and side Wall means References Cited by the Examiner UNITED STATES PATENTS 3,073,113 1/1963 Faught 60-39.47 X 3,112,611 12/1963 Adamson 60--35.6 3,127,739 4/1964 Miller 6035.6

MARK NEWMAN, Primary Examiner.

D. HART, Assistant Examiner.

Claims

1. A ROCKET ENGINE COMPRISING, IN COMBINATION, ELONGATED CASING MEANS ADAPTED TO CONTAIN PROPELLANTS FOR SAID ENGINE AND HAVING A BOTTOM WALL AND SIDE WALL MEANS MOVABLY ARRANGED WITH RESPECT TO SAID BOTTOM WALL; MEANS FORMING AN ANNULAR CHAMBER ATTACHED TO SAID BOTTOM WALL; A PLURALITY OF PROPELLING NOZZLES MOUNTED ON SAID CHAMBER FORMING MEANS AND COMMUNICATING WITH THE INTERIOR OF SAID ANNULAR CHAMBER FORMED THEREBY; PROLELLANT PUMPING MEANS MOUNTED ON SAID CHAMBER FORMING MEANS ALTERNATELY WITH SAID NOZZLES FOR PUMPING PROPELLANT FROM SAID CASING INTO SAID ANNULAR CHAMBER; PRIME MOVER MEANS FOR DRIVING SAID PUMPING MEANS; AND CUTTER MEANS DRIVEN BY SAID PRIME MOVER MEANS FOR GRADUALLY DESTROYING SAID SIDE WALL MEANS AS THE PROPELLANT IS USED UP.