RU2511961C1 - Cooling system of liquid-propellant engine chamber - Google Patents

Abstract

FIELD: engines and pumps.

SUBSTANCE: in a cooling system of a liquid-propellant engine chamber, which includes a cylindrical combustion chamber and a nozzle containing in its turn convergent and divergent parts and a critical section between them, which are made in the form of an external cover, an internal cover with the main fins of constant thickness, which form a cooling circuit, according to the invention, on the inner surface of the external cover in the area of convergent and divergent parts of the combustion chamber there are additional longitudinal fins; with that, height and thickness of additional longitudinal fins does not exceed height and thickness of the main fins.

EFFECT: improved cooling of a critical section of the nozzle and increased specific thrust of an engine.

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Description

The invention relates to the field of rocket propulsion and can be used to create chambers of liquid rocket engines (LRE), in particular for generatorless liquid propellant rocket engines operating on cryogenic components, such as oxygen and hydrogen.

A known cooling path of a liquid-propellant rocket chamber containing an inner profiled shell with ribs of constant thickness made therein, forming cooling duct channels having a trapezoidal profile of variable width with radially rounded corners adjacent to the inner surface of the shell, a profiled outer shell mounted on the inner and bonded to it (MV Dobrovolsky et al. "Liquid rocket engines. Fundamentals of design", Moscow, "Higher school", 1968, Fig. 4.26., with r.166-167).

In the indicated cooling path, the cooler is supplied between the fins made on the back side of the shell. The working surface facing the heat source, in this case, interacting with the combustion products, is made smooth cylindrical. Combustion products, in contact with the working surface, give it heat. Due to the thermal conductivity of the metal, heat from the shell is transferred to the ribs of the cooling path, which are washed by the cooler. The cooler, passing through the cooling channels, is in contact with the surfaces of the ribs and the back side of the shell, and while heating itself, it cools the ribs and the inner working surface of the inner shell.

With this design of the cooling duct, it is necessary to select the optimal thickness of the working fire wall, fins and heat exchange surface area. On the one hand, with the thinning of the fire working wall, the conditions of heat exchange improve, on the other hand, the wall thickness is limited by the strength and manufacturing conditions. An increase in the number of ribs leads to an improvement in heat transfer conditions, but at the same time leads to clutter of the duct, which increases the hydraulic resistance of the cooling duct and leads to an increase in the pump power for supplying the cooler to the cooling duct.

When using this cooling path in engines operating according to a generatorless circuit, the required heat removal from the surface of the combustion chamber is not provided. The cooler, which is then used to drive the TNA turbine, does not heat up to a predetermined temperature, which leads to a decrease in the efficiency of the turbine and the entire turbopump assembly as a whole.

A known cooling system of a chamber of a liquid propellant rocket engine according to the patent of the Russian Federation for invention No. 2158841, IPC F02K 9/62, publ. November 10, 2000 This chamber contains a housing, ignition means and a mixing head. The mixing head consists of an inner fire bottom, a middle bottom, an outer bottom, two-component nozzles are fixed in the inner fire bottom and the middle bottom. Some of the two-component nozzles are installed protruding beyond the internal fire bottom, and the other part is recessed in the fire bottom. The ignition means are made of jet nozzles installed in the power housing behind the internal fire bottom. The axis of the flow openings of the jet nozzles are located at an acute angle to the exit of the power housing and are deflected in a circle in the transverse plane from the longitudinal axis of the power housing in the same direction. The camera body includes a combustion chamber and a nozzle made of a power shell, a fire wall. The regenerative cooling duct is located between the power shell and the fire wall. The annular slit of the curtain belt is made in the inner fire wall before the critical section of the nozzle. The regenerative cooling section of the chamber is made with a branched entrance. One of its branches is connected with the cavity of the cooling path between the critical section of the nozzle and its cut, the second branch is with the cavity of the cooling path before the critical section of the nozzle, and the third is with the cavity of the cooling path in front of the annular slit of the curtain belt. This embodiment of the camera and the housing will improve the technical and operational characteristics of the engine and its resource when turned on repeatedly.

The disadvantage is a decrease in engine specific thrust due to the use of curtain cooling.

Known cooling system for the rocket engine chamber according to the patent of the Russian Federation for invention No. 2403424, IPC F02K 9/64, publ. 06/27/2010, the prototype.

This chamber contains a cylindrical combustion chamber and a nozzle having a tapering and expanding part and a critical section between them, the nozzle having inner and outer walls of the chamber, with ribs connected by brazing to the outer wall on the inner wall.

The disadvantage is poor cooling of the critical section of the nozzle and the zone around it (in front of and behind the critical section) due to the high specific heat fluxes. Additional ribs increase the heat removal, but do not reduce the temperature of the inner shell. The presence of ribs in the gas path of the nozzle reduces the specific thrust of the engine due to additional gas-dynamic losses.

The objective of the invention is to remedy these disadvantages and create a cooling path, the design of which will improve the heat transfer conditions between the combustion products and the cooler in the critical section area and increase the specific engine thrust.

The solution of these problems was achieved in the cooling system of a liquid-propellant rocket engine chamber, containing a cylindrical combustion chamber and a nozzle, which, in turn, tapering and expanding parts and a critical section between them, made in the form of an outer shell, an inner shell with main ribs of constant thickness, forming cooling path, in accordance with the invention, on the inner surface of the outer shell in the region of the tapering and expanding parts of the combustion chamber, additional longitudinal ribs, the height and thickness of additional longitudinal edges does not exceed the height and thickness of the main ribs

The most optimal conditions for heat transfer and cooling are achieved when the height and thickness of the additional longitudinal ribs does not exceed the wall thickness of the bottom of the channel. A further increase in geometric dimensions above the specified limits leads to a deterioration in the conditions of heat removal, burning and fusion of additional ribs.

The invention is illustrated by the drawings of figures 1 ... 4, where:

figure 1 shows a longitudinal axial section of the combustion chamber,

figure 2 is a cross section of the cooling path.

figure 3 and 4 shows a section of the combustion chamber with the placement of the main and additional ribs.

The design of the chamber is presented in figure 1 ... 4 and contains a combustion chamber 1 and a nozzle 2.

The combustion chamber 1 is cylindrical. The nozzle 2 has a tapering part 3, an expanding part 4 and a critical section 5 between them.

The combustion chamber 1 and the nozzle 2 have an inner shell 6 and an outer shell 7. On the inner shell 6 are made of the main ribs 8 of constant thickness, forming between them the channels 9 of the cooling path. Channels 9 have a trapezoidal profile of variable width

On the inner surface of the outer shell 7 in the region of the critical section 5 (before and after the exhaust gas flow), additional ribs 10 are made. Additional ribs 10 are placed between the main ribs 8 and have a smaller thickness and height than the main ribs 8.

The placement zone of the additional ribs 10 (figure 3 and 4) is determined from the condition:

- to critical section 5:

11 = (0.4 ... 0.6) L1, where L1 is the length of the subsonic part of the nozzle.

- after critical section 5:

12 = (0.4 ... 0.6) L2,

where L2 is the length of the expanding part (supersonic) of the nozzle from the critical section to the section having a diameter D1 equal to the inner diameter of the combustion chamber.

Every second main rib 10 is made shortened 11. In this case, additional ribs 12 are placed in the vacant gap of the cooling path 9 (Fig. 4).

The proposed device operates as follows.

During operation of the LRE chamber, the combustion products of the fuel components move along the channels 9 along the wall of the inner shell 6 and transfer heat and it to the main ribs 8. Due to the thermal conductivity, the entire inner shell 6, including the main ribs 8, is warmed up. A cooler enters the channels 9 of the cooling path, which washes the inner shell 6, the main ribs 8 and the bottom of the channel 9. The cooler, having a temperature below the temperature of the ribs and the bottom of the channel 0, takes away them warm and heats itself.

To improve the conditions of heat transfer from the combustion products to the cooler, additional ribs 10 are made on the inner surface of the outer shell 7. In this case, part of the heat will be taken from the combustion products using these additional ribs 10. In addition, the presence of additional ribs 10 clutters the channels 9 and increases they speed of the cooler is approximately 2 times and thereby increases the coefficient of heat transfer to the cooler by 80% ... 90%. This will significantly improve the cooling of the critical section 5 and the zone around it.

Claims (1)

The cooling system of a liquid-propellant rocket engine chamber, comprising a cylindrical combustion chamber and a nozzle, which, in turn, comprises a tapering and expanding part and a critical section between them, made in the form of an outer shell, an inner shell with main fins of constant thickness forming a cooling path, characterized in that on the inner surface of the outer shell in the region of the tapering and expanding parts of the combustion chamber, additional longitudinal ribs are made, while the height and thickness of the additional s longitudinal edges does not exceed the height and thickness of the main ribs.

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