Power plants – Reaction motor – Liquid oxidizer
Reexamination Certificate
1999-09-08
2001-01-09
Thorpe, Timothy S. (Department: 3746)
Power plants
Reaction motor
Liquid oxidizer
Reexamination Certificate
active
06170258
ABSTRACT:
FIELD OF THE INVENTION
The invention relates to rocket engineering and may be used in the production of rocket power plants based on the liquid propellant rocket engine.
BACKGROUND OF THE INVENTION
High-power liquid-propellant rocket engines are known in rocket engineering and they are widely used as parts of launchers designed for various purposes.
Known in the art under code name RD-253 is a liquid-propellant rocket engine (LRE) designed by the “Energomash” Scientific and Production Corporation (NPO Energomash) (see in the encyclopaedia: “Cosmonautics”, Editor-in-Chief V. P. Glushko, Moscow, 1985, pp. 330-331). This LRE comprises a combustion chamber, a gas generator, a turbopump assembly, automatic equipment, pipelines that interconnect hydraulically engine assemblies with each other, a gas line that connects the turbine of the turbopump assembly to the combustion chamber, units for fastening the LRE to a launcher so as to make it capable of turning (swinging) in a vertical plane.
A disadvantage of this prior art LRE design consists in that it can ensure turning (swinging) of the LRD, as required for changing its thrust vector direction, only in a single plane. It should be also pointed out that large size of the bellows balances used in this prior art design leads to an increased axial overall dimension of the engine.
These disadvantages of the prior art design also exist in the design of a dual-chamber LRE which also has a considerable axial overall dimension. The above disadvantages prevent the production of a launcher having smaller mass and overall dimensions.
Also known in the art is an F-1 LRE designed by Rocketdyne Co. of the U.S.A. (see in the book: TsIAM, “Foreign Aircraft and Rocket Engines”, 1971, pp. 436-439). This LRE comprises a combustion chamber, a gas generator, a turbopump assembly, automatic equipment, inner pipelines of the engine, and a gimbal assembly.
As applied to a dual-chamber engine design, this prior art design requires the development of a special frame. In a number of cases, it is also difficult to ensure smaller axial overall dimensions for this engine and, hence, to optimize the launcher as far as its overall dimensions and mass are concerned.
The closest to the liquid-propellant rocket engine of the present invention is an LRE under code name RD-219 designed by NPO Energomash (see in the encyclopaedia: “Cosmonautics”, Editor-in-Chief V. P. Glushko, Moscow, 1985, p. 330). This prior art liquid-propellant rocket engine comprises two combustion chambers fixed to a frame, a turbopump assembly fastened also to the frame and having a turbine, oxidizer and fuel pumps, and pipes for feeding the oxidizer and fuel to a gas generator and to the combustion chamber of the engine.
A limitation of this prior art design consists in that the engine chambers are rigidly fixed to the frame. They cannot turn so as to change the thrust vector direction and, in a number of cases, they require to provide special control spaces on board the launcher. Besides, this engine has no pipe between the turbine and the combustion chambers so that gas is exhausted from the gas generator overboard, instead of contributing to an increase in the specific thrust developed by the rocket engine as the case is for a system with gas from the gas generator being afterburned. If this engine is made with “swinging” combustion chambers using the system for afterburning the gas, then both the axial and diametric overall dimensions of the design will become larger.
SUMMARY OF THE INVENTION
The principal object of the present invention is to provide an LRE design which ensures a reduction in the overall axial dimension of the engine as well as in the overall axial dimension and mass of the launcher. Another object is an improvement in the use of the engine compartment space of the rocket by the LRE components and assemblies which are to be received therein. Still another object is to simplify the process of assembling the engine.
The necessity for arranging the components more compactly becomes particularly pressing in those instances when it is required to update obsolete launchers by equipping them with advanced modern rocket engines.
The essence of the invention consists in that in the prior art liquid-propellant rocket engine comprising two combustion chambers fixed to a frame, a turbopump assembly also fastened to the frame and having a turbine, oxidizer and fuel pumps, and pipes for feeding the oxidizer and fuel to a gas generator and to the combustion chambers of the engine, said frame is made demountable and comprises two sections disposed symmetrically relative to the engine axis, said frame is provided with supports for taking up forces from said rocket engine and with footpads for fastening said frame to the rocket body, said footpads and said supports being located in the planes perpendicular to the longitudinal axis of said rocket engine, the footpad location plane being located between the support location plane and the rocket body, wherein a branched curved pipe is incorporated therein additionally for feeding high-temperature oxidizing gas from the turbine outlet to the combustion chambers of the engine and having the unbranched end passage thereof connected to the turbine outlet and having the two branches thereof connected to two sections of the frame disposed symmetrically relative to the engine axis, and wherein said two branches of the branched curved pipe are connected to said combustion chambers through corresponding bellows balances, each being a swinging unit for its corresponding combustion chamber, and each of the pipes for feeding the fuel to said combustion chambers of the engine comprises two bellows balances.
In addition, said bellows balances of the pipes for feeding the fuel to said combustion chambers are made so as to be capable of compensating for the angular displacements of said pipes in two mutually perpendicular planes which are parallel to the axis of the liquid-propellant rocket engine.
Besides, the frame sections are made of rods welded together and they are fastened to each other by a plane rod joint.
Furthermore, said plane rod joint is shaped as a spider, and hollow rods are used as said rods.
Moreover, each of said bellows balances of the pipes for feeding the fuel to said combustion chambers of the engine is provided with a cardan mechanism.
Also, the bellows balances which are the swinging units for the combustion chambers are provided with appropriate cardan mechanisms.
Further, said branched curved pipe for feeding high-temperature oxidizing gas has the two branches thereof connected to the two sections of the frame disposed symmetrically relative to the engine axis by means of attachment units.
Besides, each of said attachment units comprises two trunnions inserted into bearing units of said supports for taking up forces from said rocket engine.
In addition, said support for taking up forces from said rocket engine is made on said frame in the form of a cross-arm.
As follows from the abovesaid, the LRE comprises two combustion chambers, a gas generator, a turbopump assembly, and engine pipelines communicating the corresponding engine assemblies with each other. The LRE also comprises a frame having footpads and supports located in different planes perpendicular to the engine axis and made up of at least two welded rod sections fastened to each other by a plane rod joint. Bellows balances or flexible hoses are mounted in the inner pipelines of the LRE which feed the fuel to the combustion chambers of the engine, said bellows balances or flexible hoses having their movable ends connected directly to the combustion chambers or to the pipes joined thereto. Here, those ends of the bellows balances or flexible hoses, which are moving together with the combustion chambers during their swinging motion relative to the rocket body, are considered to be the movable ends.
The frame has footpads and supports located in different planes perpendicular to the engine axis. The frame is made as a compensation frame due to the location of the footpad plan
Chelkis Felix Jurievich
Chvanov Vladimir Konstantinovich
Katorgin Boris Ivanovich
Murlykina Nina Ivanovna
Polushin Valentin Georgievich
Banner & Witcoff , Ltd.
Gartenberg Ehud
Otkrytoe Aktsionernoe Obschestvo “Nauchno” Proizvodstvennoe Obie
Thorpe Timothy S.
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