US3719040A

US3719040A - Gas generator and tubular solid charge construction therefore

Info

Publication number: US3719040A
Application number: US00097029A
Authority: US
Inventors: H Hofmann
Original assignee: Messerschmitt Bolkow Blohm AG
Current assignee: Airbus Defence and Space GmbH
Priority date: 1968-04-30
Filing date: 1970-12-10
Publication date: 1973-03-06
Anticipated expiration: 1990-03-06
Also published as: DE1751268B1

Abstract

A device for generating gases such as a combustion chamber of a rocket engine includes a solid fuel propellant in the form of a substantially cylindrical hollow main charge having a central bore therethrough and located within the combustion chamber at a spaced location from the interior walls thereof and between a detonating charge and a priming charge. The end of the main charge facing the detonating charge is located adjacent an outlet for the combustion chamber and it is formed in a manner to ensure that some of the ignition and combustion gases which are formed by ignition of the center of the main charge will flow downwardly through the opposite end of the main charge past the priming charge and around the outside of the main charge to cause and maintain ignition of this outside portion of the main charge. The remaining portion of the gases will flow directly from the end of the charge adjacent the detonator to the outlet. To accomplish the desired flow conditions, the main charge itself is formed with a total or partial constriction or nozzle portion at its end which is adjacent the combustion chamber outlet.

Description

United States Patent Hofmann GAS GENERATOR AND TUBULAR SOLID CHARGE CONSTRUCTION THEREFORE Heinrich Hofmann, Germany Messerschmitt-Bolkow GmbH, Ottobrunn near Munich, Germany Filed: Dec. 10, 1970 Appl. No.: 97,029

Inventor: Grobenzell,

Assignee:

Related US. Application Data Division of Ser. No. 806,898, March 13, 1969.

Foreign Application Priority Data April 30, 1968 Germany ..P 17 51 268.4

References Cited UNITED STATES PATENTS FOREIGN PATENTS OR APPLICATIONS 951,237 3/1964 Great Britain ..60/27l 1,167,120 4/1964 Germany ..60/253 Primary Examiner-Manuel A. Antonakas Attorney-McGlew & Toren [57] ABSTRACT A device for generating gases such as a combustion chamber of a rocket engine includes a solid fuel propellant in the form of a substantially cylindrical hollow main charge having a central bore therethrough and located within the combustion chamber at a spaced location from the interior walls thereof and between a detonating charge and a priming charge. The end of the main charge facing the detonating charge is located adjacent an outlet for the combustion chamber and it is formed in a manner to ensure that some of the ignition and combustion gases which are formed by ignition of the center of the main charge will flow downwardly through the opposite end of the main charge past the priming charge and around the outside of the main charge to cause and maintain ignition of this outside portion of the main charge. The remaining portion of the gases will flow directly from the end of the charge adjacent the detonator to the outlet. To accomplish the desired flow conditions, the main charge itself is formed with a total or partial constriction or nozzle portion at its end which is adjacent the combustion chamber outlet.

7 Claims, 3 Drawing Figures '1 7. I Za Ga PATENTED W73 3,719,040

Za Ga 1.1 Fig. 3

INVEN TOR Heinrich Hofmann y CGM m ATTORNEYS GAS GENERATOR AND TUBULAR SOLID CHARGE CONSTRUCTION THEREFORE This is a division of application, Ser. No. 806,898, filed Mar. 13, 1969.

SUMMARY OF THE INVENTION This invention relates, in general, to the construction of a solid charge for a gas generator and to an arrangement of a gas generator and, in particular, to a new and useful rocket engine combustion chamber with a solid charge therein arranged between a detonating charge and a priming charge and having a hollow bore with one end adjacent the combustion chamber outlet which is formed in a manner to ensure the flow of gases generated by ignition of the interior of the hollow charge around the exterior of the hollow charge to cause and maintain ignition thereof.

It is generally known to use tubular rocket charge compositions for the operation of solid fuel rocket engines and to set them afire by means or one or more priming charges which are arranged in the combustion chamber head. Such priming charges are usually ignited by detonator caps. The ignition of such solid fuel rocket compositions, particularly for small engines and engines which must work at very low ambient temperatures, is difficult. Special problems arise with the use of tubular charges because the outer jacket or outer surface of the charge is not sufficiently admitted by the ignition gases of the priming charge and it is not maintained at a high enough temperature for a long enough period of time to ensure its continued ignition. The reason is that the flow resistance in the annular passage defined between the exterior wall of the tubular charge and the combustion chamber is substantially higher than that from the end of the hollow charge which is adjacent the combustion chamber outlet. This means that the ignition gases produced by the ignition of the interior of the hollow charge and by directing the gases of the priming charge through the interior will follow the path of least resistance and flow primarily through the interior of the rocket composition to the combustion chamber outlet. In addition, the smaller amount of ignition gases which flow through the narrow outlet provided by the annular passage around the exterior of the hollow charge will cool off very rapidly due to the great heat transfer to the cold combustion chamber wall which forms a comparatively large surface surrounding the tubular charge. This causes a great ignition delay and poor combustion efficiency and reduced specific impulse. In many cases, the ignition temperature which is necessary for kindling the combustion of the charge is not reached adjacent the outer wall of the tubular charge. This means that the pressure necessary for the operation of the combustion chamber will drop below a critical value and may even result in the complete extinction of the rocket charge composition. It is for these reasons that the use of the tubular burners which have certain advantages from other aspects in terms of combustion and less insulation requirements are counter balanced and this is particularly true in respect to small rocket engines and engines which must work above the normal operating conditions. However, until the present time, the installations of this nature required front end pure inner burners.

In accordance with the present invention, the com bustion chamber is constructed so that a tubular burner may be employed without the disadvantages of the prior art and to achieve such an arrangement the tubular burner is advantageously made of a construction so that gas flow through one end thereof is inhibited to the extent that it ensures flow of the generated gases produced by combustion around the opposite end of the charge and the exterior wall of the charge at least initially. The present invention is directed specifically to the particular construction of the tubular charge and the combustion chamber arrangement. The tubular charge has an end which is adapted to be located adjacent the combustion chamber outlet which is provided with a closing or throttling portion to reduce or temporarily stop the direct flow of generated gases and ignition gases from this end to the outlet and to promote a flow around the opposite end and around the exterior wall of the charge to the outlet of the combustion chamber. In this manner, the gases produced by ignition of the interior of the hollow charge and by the primer charge and detonator will flow around the exterior wall of the tubular charge at least for a period and to an extent such that the exterior will burn freely and continue burning. In accordance with one embodiment of the invention, the end of the charge is constructed in a nozzle form or otherwise narrowed to define a constriction or closure at the end closest to the combustion chamber outlet. This constriction or closure exists at least during the ignition phase or initial phase of the combustion end and is progressively reduced by combustion.

Since solid rocket compositions suitable for charges and thrust gas generators can be either cast or formed by many other production methods into a desired shape without great expenditure, it is easily possible to form the charge in a desired manner with a nozzle formation or a closed end. By constructing the charge itself of the desired configuration, it is assured that the optimum operating conditions would prevail independent of the environmental conditions.

The end of the tubular main charge may, for example, be formed as a cylindrical jacket having a starshaped inner end burner adjacent one end which ex tends inwardly beyond the uniform diameter bore of the remaining portion and provides a nozzle restriction at this end. By constructing the end in a star-shaped configuration, a faster bum-off is assured than would be the case if the surface were smooth. The remaining portion of the bore of the hollow charge may be made smooth, however, so that the nozzle constriction will be burned down at a rate which is slightly faster than the burning of the interior of the hollow charge so that a non-restricted end formation will result after operation progresses to the extent that the exterior surface of the charge has achieved a desired burning characteristic. In an alternate arrangement, a fuel insert is arranged at the end of the hollow charge which is adjacent the combustion chamber outlet. The fuel insert provides a nozzle throat section which is narrower than the remainingbore of the hollow charge and it advantageously is made of a chemical substance having a faster combustion rate than the main charge. In those instances in which the detonator cap and the priming charge are arranged at the same end of the charge, that is, the end which is remote from the combustion chamber outlet, then I the end which is adjacent the combustion chamber outlet may be closed by the charge material itself, the thickness of the charge material being relatively slight so that the burn-off will be effected after the exterior of the charge has fully ignited. After the burnoff, the gases will flow through each end of the charge through the outlet of the combustion chamber.

Accordingly, it is an object of the invention to provide, an improved tubular charge for use in a gas generator such as a rocket engine wherein one end of the charge is formed with a closure, or a restriction to either throttle or temporarily reduce the flow of combustion gases out through this end after ignition of the interior of the charge takes place, the closure or restriction being preferably formed of a material which will burn-off to open this end and to permit gas flow therethrough.

A further object of the invention is to provide a combustion chamber construction in which a tubular charge is arranged between a detonator and a primer charge and which includes a bore having one end adjacent the detonator which is also located adjacent the outlet of the combustion chamber and an opposite end adjacent the primer charge, the end adjacent the detonator being closed at least partially or restricted in order to ensure that at least some flow of the combustion gases will be through the end adjacent the primer and around the exterior of the charge before passing to the outlet.

A further object of the invention is to provide a tubular charge construction and a combustion chamber construction which are simple in design, rugged in construction and economical to manufacture.

The various features of novelty which characterize the invention are pointed out with particularity in the claims annexed to and forming a part of this specification. For a better understanding of the invention, its operating advantages and specific objects attained by its uses, reference should be had to the accompanying drawings and descriptive matter in which there are illustrated preferred embodiments of the invention.

BRIEF DESCRIPTION OF THE DRAWINGS In the drawings:

FIG. 1 is a partial transverse sectional view of a rocket engine constructed in accordance with the invention;

FIG. 2 is an axial sectional view of another embodiment of the tubular charge of the type indicated in FIG. 1; and

FIG. 3 is a view similar to FIG. 2 of still another embodiment of the invention.

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS Referring to the drawings, in particular, the invention embodied therein in FIG. 1 comprises a tubular rocket charge or rocket composition 2 arranged within a combustion chamber 1 and spaced from the interior walls 13 of the combustion chamber by means of a centering and holding device 3. The combustion chamber 1 includes a closed wall 1a adjacent which is mounted a priming charge 4. The charge 4 is located at one end of a uniform diameter bore portion of the tubular charge 2 in a position to permit gas flow therethrough and over or through the primer charge 4 and passages of the centering and holding device 3 and into an annular passage 14 formed between an exterior wall 16 of the hollow charge and the interior wall 13 of the combustion chamber.

In accordance with the invention, a detonating cap 5 is carried in an elongated tubular housing 6 which is mounted on an opposite closed wall of the combustion v chamber (not shown) and which includes an inner tubular end 6a having a flange or extension 10 defining a screen with a plurality of bores 10a. The housing 6, together with the flange l0, define annular space 11. Ignition and combustion gases may pass from the interior J of the charge 2 through openings 10a of the screen 10 into the annular space 11 and .out through a discharge or outlet 12 of the combustion chamber.

In accordance with the invention, the outflow end of the tubular charge 2, that is, the end adjacent the outlet 12 is provided with a constriction 21 which, in accordance with this feature of. the invention, is an integral portion of the solid charge material. The constriction 1 includes a nozzle projection forming a starshaped burner 22 which extends inwardly further into the smooth bore portion 15 of the tubular charge 2.

In the embodiment according to FIG. 2, a tubular charge 2 includes a smooth bore portion 15 and an outlet end having a nozzle forming constriction 31. The constriction 31 comprises a rocket composition material which burns faster than the material of the tubular charge 2.

In the embodiment indicated in FIG. 3, there is provided a tubular charge 2" in which the constriction is reduced to zero in the form of a closure 41 at the outlet end of the tubular charge 2". The closure 41 is made very concave on its inner side so that the thinnest central region of the constriction burns-off first and rapidly during the initial phase of the combustion and before the full pressure build-up on the interior of the tubular charge 2" can take place. With this arrangement both the primer charge 4 and the detonator must be located adjacent the open end of the charge 2".

The method of operation of the device is as follows:

The ignition jet of the detonator cap 5, as indicated in FIG. 1, will penetrate through the interior J of the rocket tubular charge 2 and impinge on the priming charge 4. This ignites the priming charge and its ignition gases will split into an inner partial current Zi and into an outer partial current Za. The ignition gas current Zi ignites the inner wall 15 of the rocket composition while the ignition gas current Za moves through the passages of the centering piece 3 and the annular space 14 to ignite the exterior walls 16 of the tubular charge 2. the constriction 21 ensures that a major part of the ignition gases produced which would otherwise take the direct way through the interior I and out through the combustion chamber 12, will be forced by the constriction at this end to move in an opposite direction and pass around the exterior wall 16 to ignite and to maintain ignition of this wall. In addition, a part U of the combustion gases which are produced by the burning of the interior wall 15 will be forced to move downwardly around the end remote from the outlet 12 and upwardly along the exterior wall to aid in the combustion of this wall by maintaining the temperature at this location very high. The partial gas current leaving the interior space J through the constriction 21 is designated with the legend Gi and it follows the path of the primer gas flow Zi. A paraboloidal design of the constriction 21 is provided to facilitate the threading of the ignition jet which is produced by the detonator cap The operation for the arrangement of FIG. 2 is substantially the same as that of FIG. 1. In respect to FIG. 2, the total amount of the ignition gases and the total amount of the propellant gases produced in the interior I is forced to flow during the ignition phase, and if necessary during the partial phase of the combustion, through the annular slot 14 along the outer wall 16 of the rocket composition.

The direct outflow of all propellant gasses produced by the combustion on the inner wall to the outlet 12 should be effected as soon as possible in order to avoid unnecessary flow losses inside the annular slot 14.

The screen 10 serves to retain any solid pieces which 10 jacent said first end initially at least restricting flow of may become detached from the tubular rocket charge 2. These pieces are retained until they are completely burned in order to avoid the carry-over and possible damage to an operating device receiving the gases such as a gas turbine. In addition, the flow screen 10 has the function, of reducing the very high temperatures of the propellant gases by the withdrawal of heat to an operating temperature which may be admissible for a turbine, for example. In addition, soot and other impurities are deposited on the screen by the cooling effect of the screen.

ln the embodiment of FIG. 3, a constriction extending inwardly into the bore and closing the outlet end thereof comprises the closed end 41 which burns away quickly to open the outlet and permits gas flow out this end after the initial gas flow causes ignition of the exterior wall 16".

; I What is claimed is:

1. A device for generating gases, such as thrust gases for a rocket engine, comprising wall means defining a cylindrical combustion chamber having an outlet, a hollow tubular combustible charge located within said combustion chamber and having a first end adjacent said combustion chamber outlet and an axially opposite gases from within said tubular charge through said first end thereof to said combustion chamber outlet to initially force at least a major portion of the gases to flow to said combustion chamber outlet around said second end and through said exterior flow passage in contact with the exterior wall of said tubular charge, and means for igniting said priming charge, said constriction means burning away following the initial combustion phase to provide substantially unrestricted flow of gases through said first end of said tubular charge to said combustion chamber outlet.

2. A device, according to claim 1, wherein said constriction means comprises a nozzle formation composed of the material of said tubular charge.

3. A device, according to claim 1, wherein said constriction means comprises a closure at said first end of said charge.

4. A device, according to claim 1, wherein said constriction means comprises an insert within said tubular charge formed of a combustible material having a combustion rate faster than that of said charge.

5. A device, according to claim 1, wherein said constriction means comprises a portion of said tubular charge material forming a nozzle of generally starshaped configuration.

6. A device, according to claim 1, wherein said interior of said tubular charge is forced into a paraboloidal form adjacent said first end and comprises said constriction means.

7. A device, according to claim 1, including said priming charge arranged adjacent said second end, and an igniter charge arranged adjacent said first end, said priming and said igniter charges being aligned with the bore of said tubular charge.

Claims

1. A device for generating gases, such as thrust gases for a rocket engine, comprising wall means defining a cylindrical combustion chamber having an outlet, a hollow tubular combustible charge located within said combustion chamber and having a first end adjacent said combustion chamber outlet and an axially opposite second end, the exterior lateral surface of said charge being spaced from said wall means to define therewith an exterior flow passage around the exterior of said charge communicating with said second end, a priming charge operable, when ignited, to produce ignition gases flowing into the interior of said tubular charge and into said exterior flow passage to ignite the internal and external surfaces of said tubular charge, combustible constriction means within said tubular charge adjacent said first end initially at least restricting flow of gases from within said tubular charge through said first end thereof to said combustion chamber outlet to initially force at least a major portion of the gases to flow to said combustion chamber outlet around said second end and through said exterior flow passage in contact with the exterior wall of said tubular charge, and means for igniting said priming charge, said constriction means burning away following the initial combustion phase to provide substantially unrestricted flow of gases through said first end of said tubular charge to said combustion chamber outlet.

5. A device, according to claim 1, wherein said constriction means comprises a portion of said tubular charge material forming a nozzle of generally star-shaped configuration.