阅读说明：本技术 一种燃烧室的点火、稳燃结构 (Ignition and stable combustion structure of combustion chamber ) 是由 李光熙 张玫 刘昊 张蒙正 刘晓伟 豆飞龙 路媛媛 于 2019-11-07 设计创作，主要内容包括：本发明涉及高超声速飞行器动力系统燃烧室技术领域,公开了一种燃烧室的点火、稳燃结构,包括设置在支板上的凹腔,在凹腔的壁面设置有至少一个自燃推进剂喷注单元,自燃推进剂喷注单元包括在凹腔的壁面上设置的氧化剂喷注孔和燃料喷注孔,氧化剂喷注孔与燃料喷注孔的喷射角度呈夹角β,通过氧化剂喷注孔向凹腔内喷射氧化剂,通过燃料喷注孔向凹腔内喷射点火用推进剂,通过自燃推进剂撞击后燃烧产生的高温燃气,不仅可以实现RBCC燃烧室的点火,而且在飞行状态极端、恶劣工况下可以作为引导火焰保持燃烧室火焰稳定,同时,将凹腔和点火器一体化设计不用再单独设置点火器,简化了燃烧室结构,降低了热防护难度。(The invention relates to the technical field of combustion chambers of power systems of hypersonic aircrafts, and discloses an ignition and stable combustion structure of a combustion chamber, which comprises a concave cavity arranged on a support plate, wherein the wall surface of the concave cavity is provided with at least one spontaneous combustion propellant injection unit, the spontaneous combustion propellant injection unit comprises an oxidant injection hole and a fuel injection hole which are arranged on the wall surface of the concave cavity, the injection angle between the oxidant injection hole and the fuel injection hole is β, oxidant is injected into the concave cavity through the oxidant injection hole, propellant for ignition is injected into the concave cavity through the fuel injection hole, and high-temperature fuel gas generated by combustion after the spontaneous combustion propellant impacts is used, so that the ignition of an RBCC combustion chamber can be realized, the high-temperature fuel gas can be used as guide flame to keep the flame stability of the combustion chamber under extreme and severe working conditions of a flight state, and an igniter is not required to be arranged independently due to the integrated design of the concave cavity, the structure of the combustion chamber is.)

1. The ignition and stable combustion structure of the combustion chamber is characterized by comprising a concave cavity (1) arranged on the combustion chamber, wherein the wall surface of the concave cavity (1) is provided with at least one spontaneous combustion propellant injection unit, the spontaneous combustion propellant injection unit comprises an oxidant injection hole (2) and a fuel injection hole (3) which are arranged on the wall surface of the concave cavity (1), the injection angle between the oxidant injection hole (2) and the fuel injection hole (3) is an included angle β, oxidant is injected into the concave cavity (1) through the oxidant injection hole (2), and ignition propellant is injected into the concave cavity (1) through the fuel injection hole (3).

2. The ignition and stable combustion structure of the combustion chamber as claimed in claim 1, wherein the included angle β is 50-80 °.

3. The ignition and stable combustion structure of the combustion chamber as claimed in claim 2, wherein: the propellant for ignition is spontaneous combustion hydrazine, and the oxidant is a nitro oxidizer.

4. The ignition and stable combustion structure of the combustion chamber as claimed in claim 3, wherein: the ignition energy Q of the combustion chamber is selected according to the total heat capacity of the combustion chamber.

5. The ignition and stable combustion structure of the combustion chamber as claimed in claim 4, wherein: the flow rate of the ignition propellant is determined according to the ignition energy Q requirement of the combustion chamber.

6. The ignition and stable combustion structure of the combustion chamber as claimed in claim 5, wherein: the structural form and the quantity of the self-ignition propellant injection units are determined according to the flow rate of the propellant for ignition.

7. The ignition and stable combustion structure of the combustion chamber as claimed in claim 5, wherein: the diameters of the oxidant injection hole (2) and the fuel injection hole (3) are determined according to the flow rate of the propellant for ignition.

Technical Field

The invention relates to the technical field of combustion chambers of power systems of hypersonic aircrafts, in particular to an ignition and stable combustion structure of a combustion chamber.

Background

A high-efficiency Ma 0-8 wide-range rocket-based combined cycle (RBCC) engine is an important development direction of power of a reusable space transport system, and a Ma 0-8 wide-range stamping combustion chamber is a core component of the RBCC engine. Reliable ignition, flame stabilization and efficient combustion organization technologies of the ramjet combustion chamber in the Ma 0-8 wide range are core key technologies of RBCC. Research shows that in the change range of 2-3 working Mach numbers in the range of Ma 4-7, the RBCC stamping combustion chamber adopts a flame stabilizing structure of a support plate and a concave cavity, and rocket gas pneumatic flame stabilization is assisted, so that better ignition and flame stabilization effects are realized. For example, CN105180211B discloses a combustor and a scramjet engine with a cavity flame stabilizer, in which a bleed air passage for bleeding air from the rear edge of the flame stabilizing cavity to the downstream is provided in the combustor, and the high-temperature and high-pressure fuel gas generated inside the cavity during combustion is led to the downstream of the cavity.

However, in the range of Ma 0-8, the incoming flow temperature of the inlet of the low-speed combustion chamber is low, the pressure is low, ignition and flame stabilization are difficult, and the wide-range reliable ignition and flame stabilization are difficult to realize by the support plate and concave cavity flame stabilization structure. If the rocket gas is adopted for pneumatic flame stabilization, a rocket engine with adjustable working condition depth needs to be designed for ensuring the specific impulse performance of the engine, and the design difficulty and technical risk of the engine are greatly increased.

Disclosure of Invention

The technical problem solved by the invention is as follows: the ignition and stable combustion structure of the combustion chamber can realize continuous supply of a propellant, realize continuous ignition and flame stabilization in a concave cavity, ensure high-efficiency and stable combustion of the wide-range combustion chamber and reliable work under extreme and severe working conditions, and avoid the design problem that the working condition of the rocket thrust chamber must be changed in a large range due to the adoption of hot gas pneumatic flame stabilization of the rocket thrust chamber.

The technical scheme includes that the ignition and stable combustion structure of the combustion chamber comprises a concave cavity arranged on the combustion chamber, at least one spontaneous combustion propellant injection unit is arranged on the wall surface of the concave cavity, the spontaneous combustion propellant injection unit comprises an oxidant injection hole and a fuel injection hole which are arranged on the wall surface of the concave cavity, the injection angle between the oxidant injection hole and the fuel injection hole is β, oxidant is injected into the concave cavity through the oxidant injection hole, and propellant for ignition is injected into the concave cavity through the fuel injection hole.

The principle of the invention is as follows: the liquid spontaneous combustion propellant injection unit is directly arranged on the wall surface of the concave cavity, the injection form of the spontaneous combustion propellant and the oxidant is designed into a direct current mutual impact type by utilizing the characteristic that the spontaneous combustion propellant burns after being impacted, and the high-temperature fuel gas generated by combustion after the spontaneous combustion propellant impacts can not only realize the ignition of the RBCC combustion chamber, but also be used as guide flame to keep the flame of the combustion chamber stable under extreme and severe working conditions of a flight state. Simultaneously, with the design of cavity and igniter integration no longer set up the point firearm alone, simplified the combustion chamber structure, reduced the heat protection degree of difficulty.

Further, the included angle β is 50-80 degrees, and the spontaneous combustion propellant sprayed from the fuel injection hole can be impacted and combusted by the oxidant sprayed from the oxidant injection hole.

Furthermore, the propellant for ignition is a spontaneous combustion hydrazine class, and the oxidant is a nitro oxidizer.

Further, the ignition energy Q of the combustion chamber is selected according to the total heat capacity of the combustion chamber.

Further, the flow rate of the ignition propellant is determined in accordance with the ignition energy Q requirement of the combustion chamber.

Further, the structural form and the number of the autoignition propellant injection units are determined according to the flow rate of the propellant for ignition.

Further, the diameters of the oxidizer injection hole and the fuel injection hole are determined according to the flow rate of the propellant for ignition.

Compared with the prior art, the invention has the advantages that:

(1) the liquid spontaneous combustion propellant injection unit is arranged in the concave cavity, and the design can be used for carrying out short-time ignition according to the working requirement of the combustion chamber or keeping flame stable as guide flame when the combustion chamber works for a long time under extreme and severe working conditions in a flying state, so that wide-range reliable work of the RBCC combustion chamber of Ma 0-8 is realized.

(2) Meanwhile, due to the integrated design of the concave cavity and the igniter, the design difficulty that the working condition of the rocket thrust chamber needs to be changed in a large range due to the adoption of the hot gas dynamic flame stabilization of the rocket thrust chamber is greatly reduced.

(3) With the design of cavity and ignition integration, need not special independent design RBCC combustion chamber's ignition, simplified the combustion chamber structure, reduced the heat protection degree of difficulty.

Drawings

FIG. 1 is a schematic illustration of the ignition and combustion stabilization configuration of the present invention;

fig. 2 is a partially enlarged view of a portion a of fig. 1.

The reference numerals 1-cavity, 2-oxidizer injector, 3-fuel injector, 4-oxidizer supply, 5-fuel supply, 6-groove, length of L-cavity, depth of D-cavity, rear edge angle of α -cavity, angle of impingement of β -autoignition propellant and oxidizer, do-oxidizer injector, df-fuel injector.

Detailed Description

The combined injection unit of the liquid spontaneous combustion propellant is directly arranged on the wall surface of the concave cavity, the injection form of the spontaneous combustion propellant and the oxidant is designed into a direct current mutual impact type by utilizing the characteristic that the spontaneous combustion propellant burns after being impacted, and the high-temperature fuel gas generated by the combustion of the spontaneous combustion propellant after being impacted can not only realize the ignition of the RBCC combustion chamber, but also be used as guide flame to keep the flame of the combustion chamber stable under extreme and severe working conditions of a flight state. Simultaneously, with the design of cavity and igniter integration no longer set up the point firearm alone, simplified the combustion chamber structure, reduced the heat protection degree of difficulty.

Specifically, as shown in fig. 1 and 2, the ignition and stable combustion structure of the combustion chamber comprises a cavity 1 arranged on a support plate, at least one spontaneous combustion propellant injection unit is arranged on the wall surface of the cavity 1, the spontaneous combustion propellant injection unit comprises an oxidant injection hole 2 and a fuel injection hole 3 which are arranged on the wall surface of the cavity 1, the injection angle between the oxidant injection hole 2 and the fuel injection hole 3 is an included angle β, the structural form and the number of the spontaneous combustion propellant injection units are determined according to the flow rate of ignition propellant, specifically, the oxidant injection hole 2 and the fuel injection hole 3 are respectively communicated with an oxidant supply channel 4 and a fuel supply channel 5, the oxidant supply channel 4 and the fuel supply channel 5 are arranged in parallel to facilitate machining, a groove 6 is arranged on the wall surface of the cavity 1 and used for machining and positioning of the spontaneous combustion hole and the fuel injection hole, the intersection of the axes of the oxidant injection hole 2 and the fuel injection hole 3 is coincided with the center of the groove 6, oxidant is injected into the cavity 1 through the fuel injection hole 3, the ignition propellant is specifically, the ignition energy of the oxidant is determined according to the total flow rate of the ignition propellant, and the total flow rate of the oxidant for ignition propellant.

Specifically, in order to ensure that the autoignition propellant ejected from the fuel ejection hole 3 can be impacted and combusted by the oxidizer ejected from the oxidizer ejection hole 2, the included angle β is 50-80 °, in fact β is the included angle between the axis of the oxidizer ejection hole 2 and the axis of the fuel ejection hole 3, and β is the impact angle between the autoignition propellant and the oxidizer.