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Ignition and stable combustion structure of a combustion chamber

A combustion chamber and igniter technology, which is applied to combustion chambers, continuous combustion chambers, combustion methods, etc., can solve the problems of low temperature and low pressure at the inlet of the combustion chamber, and the flame stabilization structure is difficult to reliably ignite and stabilize the flame. The effect of simplifying the structure of the combustion chamber, reducing the difficulty of design, and reducing the difficulty of thermal protection

Active Publication Date: 2020-10-16
XIAN AEROSPACE PROPULSION INST
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AI Technical Summary

Problems solved by technology

[0003] However, in the range of Ma0-8, the inlet flow temperature and pressure of the combustion chamber in the low-velocity section are low, the ignition and flame stability are difficult, and the flame stability structure of the support plate + concave cavity is difficult to achieve reliable ignition and flame stability in a wide range
If rocket gas aerodynamic flame stabilization is used, in order to ensure the specific impulse performance of the engine, it is necessary to design a rocket engine with adjustable operating conditions, which greatly increases the difficulty of engine design and technical risk

Method used

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  • Ignition and stable combustion structure of a combustion chamber
  • Ignition and stable combustion structure of a combustion chamber

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Embodiment 1

[0024] Taking the air flow at the inlet of the firing chamber as 4.5kg / s, the total temperature as 1600K, the static pressure as 0.045MPa, and the speed as Ma3.2 as an example, the depth D of the concave cavity is 30mm, and the length-to-depth ratio L / D is 7.0. The design steps of the ignition stabilization structure with a trailing edge angle α of 40° are as follows:

[0025] (1) According to the air flow, temperature, pressure and speed at the inlet of the combustion chamber, the ignition energy Q of the combustion chamber is determined to be 0.3MW.

[0026] (2) The propellants in the rocket thrust chamber of the RBCC engine are dinitrogen tetroxide and anhydrous hydrazine, and the pyrophoric propellants for ignition are determined to be dinitrogen tetroxide and anhydrous hydrazine.

[0027] (3) According to the ignition energy Q of the combustion chamber, determine the flow rate of the spontaneous combustion propellant, the specific flow rate of dinitrogen tetroxide is 22.6...

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Abstract

The invention relates to the technical field of hypersonic flight vehicle power system combustion chambers, and discloses an ignition and stable combustion structure of a combustion chamber. The ignition and stable combustion structure of the combustion chamber comprises a concave cavity arranged on a support plate, wherein at least one hypergolic propellant injection unit is arranged on the wallsurface of the concave cavity and comprises an oxidizing agent injection hole and a fuel injection hole formed in the wall surface of the concave cavity; injection angles of the oxidizing agent injection hole and the fuel injection hole form an included angle beta; an oxidizing agent is injected into the concave cavity through the oxidizing agent injection hole; a propellant for ignition is injected into the concave cavity through the fuel injection hole; through a high-temperature fuel gas produced by combustion after a hypergolic propellant is impacted, the ignition of an RBCC combustion chamber can be realized, and the high-temperature fuel gas can be used as a guiding flame under the extreme and severe working conditions of flight states so as to keep the flame of the combustion chamber to be stable; and meanwhile, the concave cavity and an igniter are integrally designed, so that the igniter has no need to be arranged independently, the structure of the combustion chamber is simplified, and the difficulty in heat protection is reduced.

Description

technical field [0001] The invention relates to the technical field of a hypersonic aircraft power system combustion chamber, in particular to a combustion chamber ignition and stable combustion structure. Background technique [0002] The high-efficiency Ma0-8 wide-range rocket-based combined cycle (RBCC) engine is an important development direction for reusable space transportation system power, and the Ma0-8 wide-range ramjet combustor is the core component of the RBCC engine. Reliable ignition, flame stability and high-efficiency combustion organization technology of the ram combustion chamber in the wide range of Ma0-8 are the core key technologies of RBCC. Studies have shown that within the range of 2 to 3 working Mach numbers in the range of Ma4~7, the RBCC ram combustion chamber adopts the flame stabilization structure of the support plate + concave cavity, supplemented by rocket gas aerodynamic flame stabilization, and achieves better performance. Ignition and flam...

Claims

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Application Information

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Patent Type & Authority Patents(China)
IPC IPC(8): F02K9/44F02K9/56F02K9/62F02K9/95F23R3/28
CPCF02K9/44F02K9/56F02K9/62F02K9/95F23R3/28
Inventor 李光熙张玫刘昊张蒙正刘晓伟豆飞龙路媛媛
Owner XIAN AEROSPACE PROPULSION INST
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