RU2448268C1 - Chamber of low-thrust rocket engine running on two-component anergolic gas fuel - Google Patents

Abstract

FIELD: engines and pumps.

SUBSTANCE: low-thrust rocket engine chamber comprises combustion chamber with nozzle and mixing head, prechamber with igniter, and fuel components feed pipelines. Mixing head comprises fuel and oxidiser feed pipelines communicated with components feed pipelines. Oxidiser feed pipeline comprises axial channel making a nozzle with its one end communicated with oxidiser feed pipeline. Axial channel wall has one or several radial bores to communicated with prechamber cavity and tails in auger to crease oxidiser swirled flow into combustion chamber. Fuel feed line comprises radial channel changing into annular manifold communicated via tangential channels with annular intermediate chamber arranged outside of axial channel communicates, on one side, with combustion chamber and, on opposite side, with, at least, one lengthwise channel provided with baffle. Axial channel making a nozzle communicating intermediate chamber with combustion chamber may be restricted by annular ledge made on outer wall of said axial channel.

EFFECT: reliable repeated start of engine at minimised engine weight.

Description

The invention relates to rocket and space technology and can be used to develop designs of small thrust rocket engines (RDMT), using two-component non-combustible gaseous fuels. Fields of application of such RDMTs are control systems for inter-orbital transportation means, orbital manned space stations and spacecraft. The invention can also be used in aircraft and in industrial power units.

A known design of the RDMT operating on a two-component gaseous non-combustible fuel: oxygen-hydrogen [1], including a combustion chamber with a nozzle and a mixing head, a pre-chamber with an igniter, supply pipelines. The device provides a constant, during the entire time the engine is running, supply of the entire flow of oxidizer (oxygen) and a small portion of the flow of fuel (hydrogen) to a pre-chamber located on the axis of the engine, in which the fuel is ignited, and most of the flow of fuel is fed into the combustion chamber. An electric spark plug is used to ignite the fuel.

In this design, radial and axial channels are used to supply fuel components to the pre-chamber and the combustion chamber. In the combustion chamber, an axial flow of combustion products flowing out of the pre-chamber is mixed with hydrogen flowing out of the annular opening from the outside with respect to the flow of combustion products.

The disadvantage of this device is the implementation of the picture of the components entering the combustion chamber with a predominance of a longitudinal layered flow and, as a result, limited possibilities for intensive mixing of the fuel components, as a result of which an increased length of the combustion chamber may be required to achieve high completeness of fuel combustion. This circumstance may be insignificant for large liquid propellant rocket engines when a large number of mixing elements are used and the interaction of flows from adjacent nozzles contributes to the mixing of fuel components. Due to the small consumption of fuel components and the size of the chamber, the RDMT can use a limited number of nozzles, up to a single amount (as is the case in the analog design, where one mixing element is used to mix the flow of fuel combustion products in the prechamber with the main fuel stream) , and this drawback entails significant losses in the completeness of fuel combustion.

A known design of the RDMT operating on a non-combustible two-component gas fuel [2], which contain a combustion chamber with a nozzle and a mixing head, a pre-chamber with an igniter, supply pipelines for supplying fuel components. The mixing head includes fuel and oxidizer feed lines connected to respective component supply pipelines.

Taken as a prototype, the design of the chamber of the thrust rocket engine (second embodiment of the invention) is characterized by the fact that the oxidizer enters the combustion chamber through the axial jet nozzle, and into the pre-chamber through a radial hole in the wall of the axial channel. The fuel supply line includes a radial channel passing into the annular manifold, communicating with the nozzle channels of the fuel supply to the combustion chamber and the channels with an annular intermediate cavity located on the outside of the oxidizer nozzle channel, which, on the one hand, communicates with the annular nozzle channel with a combustion chamber and, on the other hand, communicates with one or more longitudinal channels with a precamera.

The technical problem to which the invention is directed is the creation of an RDMT design with high completeness of combustion of a non-combustible bicomponent gaseous fuel, with the provision of reliable multiple start-up of the RDMT with a minimum mass of the engine structure.

In the proposed design, as in the prototype, pre-chamber ignition of the fuel is used with a constant supply of a portion of the oxidizer flow rate to the pre-chamber and a starting portion of fuel from the intermediate cavity to the pre-chamber when the engine is started. An increase in the completeness of fuel combustion in the combustion chamber is achieved by intensifying the mixing of the fuel components by swirling the entire gas flow.

To solve this problem, the following version of the structural solution of the RDMT camera is proposed.

The chamber of a small thrust rocket engine (RDMT) operating on a two-component non-combustible gaseous fuel contains a combustion chamber with a nozzle and a mixing head, a pre-chamber with an igniter, and supply pipelines for supplying fuel components. The design of the mixing head includes a fuel and oxidizer supply lines connected to respective component supply pipelines.

The oxidizer feed line includes an axial nozzle channel connected at one end to the oxidizer feed pipe and connected to the combustion chamber at the other end. The wall of the axial channel of the nozzle has one or more radial openings for communication with the cavity of the prechamber and ends with a screw forming a swirling flow of oxidizer into the combustion chamber.

The fuel supply line includes a radial channel connected to the fuel supply pipe, passing into an annular manifold communicating with tangential channels with an annular intermediate cavity located on the outside of the oxidizer nozzle channel of the oxidizer, which, on the one hand, is connected by an annular nozzle channel to the combustion chamber and, on the other hand, one or more longitudinal channels with a precamera.

The portion of the nozzle channel connecting the intermediate cavity to the combustion chamber may have a narrowing with an annular protrusion made on the outer wall of the nozzle channel.

The essence of the invention is illustrated in the drawing.

The RDMT chamber includes a combustion chamber 1 with a nozzle 2, a pre-chamber 3 with an igniter 4, a supply pipe 5 for supplying an oxidizer chamber and a supply pipe 6 for supplying a fuel chamber, a mixing head 7. The oxidizer enters the combustion chamber through an axial channel-nozzle 8 with a screw 9 at the end and in the chamber, through a radial hole 10 in the wall of the axial channel 8. The fuel supply line from the pipeline 6 to the combustion chamber 1 sequentially includes a radial channel 11, an annular collector 12, tangential channels 13, an intermediate cavity 14, an annular channel-injector 15 for introducing fuel into the combustion chamber 1 and the channel 16 for supplying fuel to the pre-chamber when starting the engine

Consider the operation of this device.

Before starting, the pressure is the same in all cavities of the engine chamber and is equal to the ambient pressure, as a rule, vacuum.

When the engine is started, the fuel components are supplied simultaneously: the oxidizing agent through the pipeline 5 and the fuel through the pipeline 6. The oxidizing agent through the axial channel-nozzle 8 enters through at least one radial channel 10 into the chamber 3 and into the screw 9, from which the swirling oxidant stream flows into the combustion chamber 1.

Fuel from the pipeline 6 through the radial channel 11 enters the annular manifold 12, from which through the tangential channels 13 it enters the intermediate cavity 14, from where it flows into the combustion chamber 1 by swirling flow through the annular channel-nozzle 15, and when the engine is started, into the pre-chamber 3 through axial channel 16.

As fuel and oxidizer enter the pre-chamber, a mixture is formed in the ratio of fuel components necessary for ignition, which ignites from the energy of the igniter 4.

High-temperature products of fuel combustion from the pre-chamber expire through the axial channel 16, the intermediate cavity 14 and the nozzle channel 15 into the combustion chamber 1, as a result of which the combustion process propagates into the combustion chamber and the engine starts.

As the engine is running, the oxidizing agent entering through the channel 10 into the pre-chamber 3 creates a pressure in it greater than the pressure in the intermediate chamber 14, and the fuel ceases to enter the pre-chamber, as a result, the combustion process in the pre-chamber stops.

Analysis of the prior art for compliance of the claimed solution with the condition of patentability of the invention “inventive step” showed the following.

In the inventive device, the intensification of mixing of the fuel components is achieved by spinning the entire flow rate of the fuel components, while a significant peripheral component of the speed will contribute to achieving a sufficient mixing length for mixing the fuel components with a small design length of the combustion chamber. Depending on the requirements for the design, the twist of the fuel components can be either one-sided or oncoming. The increase in the completeness of fuel combustion due to the swirling of gas flows at the inlet to the combustion chamber along with the implementation of reliable multiple start-up and the possibility of minimizing the mass of the chamber structure is confirmed by the results of experimental work at the stands of the Keldysh Center.

Information sources

1. Appel M.A., Schoeman L., Berkman D.K. Oxygen / Hydrogen Thrusters for the Space Station Auxilary Propulsion Systems. JANNAF Propulsion Conference, 1984, p.29-37.

2. "The chamber of a small thrust rocket engine running on a two-component non-combustible gaseous fuel." Application for invention No. 2008117806 dated 05/07/2008, patent No. 2369766.

Claims (2)

1. The chamber of the small thrust rocket engine (RDMT), operating on a two-component non-combustible gaseous fuel, comprising a combustion chamber with a nozzle, a pre-chamber with an igniter, pipelines for supplying fuel components and a mixing head including an oxidizer feed line, which represents an axial channel-nozzle, connected at one end to the oxidizer feed pipe and at the other end connected to the combustion chamber and through one or more radial holes in the wall of the nozzle channel, with with a pre-chamber cavity, a fuel supply line including a radial channel connected to the fuel supply pipe, passing into an annular manifold communicating with channels with an annular intermediate cavity located on the outer side of the oxidizer nozzle channel of the oxidizer, which is connected on one side by an annular channel nozzle with a chamber combustion and, on the other hand, communicates with one or more longitudinal channels with a prechamber, characterized in that a screw is installed at the exit from the axial channel of the oxidizer nozzle, and the channels of the fuel supply line communicating the annular manifold with the annular intermediate cavity are made tangential. 2. The chamber of the thrust rocket engine according to claim 1, characterized in that the portion of the nozzle channel connecting the intermediate cavity to the combustion chamber has a narrowing with an annular protrusion made on the outer wall of the nozzle channel.

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