JPS62150060A - Cooling device of rocket engine - Google Patents

Abstract

PURPOSE:To make it possible to improve cooling while enhancing combustion efficiency of a low output rocket engine such as an apogee engine by arranging an annular fin integral with a burner near the periphery of a throat section. CONSTITUTION:In a burner 1, fuel from an injector 2a and propelling agent from an oxidizing agent injector 2b are burned in a combustion chamber 3, accelerated and jetted out of a nozzle 4. Near the periphery of a throat section 5 formed between the combustion chamber 3 and the nozzle 4 is integrally fixed on annular fin 6 and a film-like cooling layer 7 is formed with liquid fuel along the inner circumferences of the combustion chamber 3 and the throat section. A heat shield 9 is also attached around them by means of a support body 8. Since radiation cooling in the throat section is increased in this manner, cooling with the film-like cooling layer can be reduced, making it possible to improve cooling while increasing the combustion efficiency.

Description

DETAILED DESCRIPTION OF THE INVENTION [Field of Industrial Application] The present invention relates to a cooling device for a rocket engine, particularly a low thrust rocket engine such as an apogee engine.

[Conventional technology]

Generally, in a rocket engine, it is necessary to cool the combustion chamber in order to withstand long-term operation.

Since the thermal load reaches its maximum at the throat, cooling the throat is a critical issue in engine design, and various cooling means have been used heretofore.

In other words, in conventional large rocket engines,
A thin tube is placed outside the combustion chamber to form a coolant passage through which fuel or oxidizer is passed for cooling (regenerative cooling), or a material that melts and vaporizes when exposed to high heat, such as silica glass or phenol. The inner wall of the combustion chamber is formed using a composite material made of resin, etc., and the heat load from the combustion gas is absorbed by the heat of evaporation of the inner wall material, and the inner wall is made of a thin plate of heat-resistant material such as niobium. Radiation cooling methods are used in which the combustion chamber and nozzle are configured and cooled by heat radiation from the outer surface.

In small engine engines used in apogee engines, cooling is generally performed using a radiation cooling method, but cooling by radiant ion alone has a limit to the amount of heat radiated due to the temperature limit of the constituent materials of the combustor. In particular, the throat section, which has the largest thermal load, is placed in an extremely severe thermal condition. For this reason, usually a film cooling method is used in which liquid fuel is poured in a film along the inner surface of the engine combustor to cool the inner surface in the liquid or gas phase, and the temperature at the throat part is adjusted to the level of the material used. The temperature is below the allowable temperature.

[Problem that the invention seeks to solve]

However, if an attempt is made to increase cooling using a film cooling method to supplement the cooling of the engine combustor using radiant sol, there is a problem in that the combustion efficiency decreases due to incomplete combustion of the injected fuel.

The present invention has been made in order to solve the above-mentioned problems in conventional rocket engine cooling systems. The object of the present invention is to provide a cooling device for a rocket engine that increases cooling of the engine.

[Means and effects for solving the problems] In order to solve the above problems, the present invention provides an annular fin that is integrally arranged with the combustor near the outer periphery of the throat portion of the combustor of the rocket engine. It is something to do.

By integrally arranging the annular fin near the outer periphery of the throat part in this way, it is possible to increase the radiation area and increase the cooling effect of radiation, thus reducing the amount of cooling required by the film cooling method. This makes it possible to increase the cooling effect of the rocket engine without reducing combustion efficiency. In addition, by connecting this annular fin to the engine support structure of the combustor, it can be used as a support member for the combustor, thereby reducing the burden on the engine support strength at the base end of the combustor. It becomes possible.

Generally, in internal combustion engines, etc., fins are installed in the combustion chamber to increase the heat dissipation effect. It is provided to increase the effect. Cooling by such a heat transfer effect is not applicable to the cooling of rocket engines used in outer space in a high vacuum state, and therefore, it has not been possible to provide fins as a cooling means for rocket engines in the past. This is the first proposed method in the present invention, and good results were obtained through simulation.

〔Example〕

Examples will be described below. FIG. 1 is a schematic diagram partially showing an embodiment of the present invention in cross section. In the figure, 1 is a combustor of a rocket engine, and the combustor 1 includes a fuel injector 2a for injecting propellant, an oxidizer injector 2b, and a combustor for burning the propellant injected from each injector 2a, 2b. It is composed of a combustion chamber 3 and a nozzle 4 that accelerates and injects the combustion gas burned within the combustion chamber 3, and a throat portion 5 is formed between the combustion chamber 3 and the nozzle 4.

Reference numeral 6 denotes an annular fin disposed near the outer periphery of the throat portion 5, which is formed of an annular plate made of, for example, niobium, which is the same material as the constituent material of the combustor 1, and is attached to the outer periphery of the throat portion by electron beam welding or the like. It is integrally fixed near the part. Note that although it is possible to form the fins 6 integrally with the combustor 1, it is difficult to cut them, so it is preferable to form them separately as described above and fix them together by welding or the like. . In addition, it is possible to use copper in addition to niobium as the constituent material of this fin, but the combustor 1
Since there is a difference in expansion coefficient from the niobium that constitutes the combustion chamber, its use may be difficult, so it is advantageous in terms of the expansion coefficient to use the same material as that of the combustion chamber as described above. Reference numeral 7 denotes a film-like cooling layer made of liquid fuel and formed along the inner surfaces of the combustion chamber 3 and the throat portion 5.

Reference numeral 8 denotes an engine support body, which is normally configured to support the base of the combustor 1 in which injectors and the like are arranged, and to hold the engine. 9 is a heat shield attached to the engine support 8, and the nozzle 4 of the combustor 1
The engine combustor 1 is arranged around the engine combustor 1 and provides heat shielding to equipment arranged around the engine combustor 1.

In the engine combustor configured in this way, the combustor main body is cooled by the heat of vaporization of the film-like cooling layer formed along the inner surface and by radiation from the outer surface. Cooling of the portion 5 is promoted by increasing radiation due to the annular fins 5 provided on the outer surface thereof.

When adopting the above-mentioned cooling means for a small engine with a combustion chamber diameter of approximately 7 to 3 cm, a diameter of the lower end of the nozzle of approximately 65 cm, and a thrust of approximately 200 kg, the diameter of the combustion chamber is approximately 2 cm.
When an annular fin of about 0 cm is integrally arranged on the outer periphery of the throat part, cooling by the film-like cooling layer is improved by 2 times compared to conventional methods.
It has been confirmed through simulation that even if the combustion chamber temperature is reduced by about 5%, the inner wall temperature of the throat can be suppressed to about 1000°C when the combustion chamber temperature is about 2500°C.

Further, if the annular fin 6 integrally disposed on the outer circumferential surface of the throat portion 5 is configured so that its outer circumferential portion is fixed to the engine support body 8, it can also function as an engine support portion. The attachment to the support body 8 at the base of the engine combustion chamber can reduce the strength.

〔Effect of the invention〕

As described above based on the embodiments, according to the present invention, radiation cooling of the throat portion is increased, so cooling by the film cooling method can be reduced, and therefore, cooling is achieved while increasing combustion efficiency. can be improved.

[Brief explanation of drawings]

FIG. 1 is a schematic cross-sectional view of a part of an embodiment of a cooling device for a rocket engine according to the present invention. 1... Engine combustor, 2a... Liquid fuel injector, 2b... Oxidizer injector, 3... Combustion chamber, 4... Nozzle, 5... Throat part, 5... Annular fin, 7... Film cooling layer, 8...
-Engine support, 9...heat shield.

Claims (1)

[Claims] 1. A cooling device for a rocket engine comprising a combustor having a throat portion, wherein an annular fin is disposed integrally with the combustor near the outer periphery of the throat portion.