A direct current ignition structure for a rocket engine combustion chamber

Abstract

The present invention discloses a direct current ignition structure for a rocket engine combustion chamber, which relates to the field of rocket engine technology and aims to solve the problem of insufficient ignition of fuel in traditional combustion chambers. The direct current ignition structure for the rocket engine combustion chamber includes an ignition agent nozzle and a liquid collecting ring. The combustion chamber is provided with one or more ignition nozzles evenly distributed along the circumference, and the wall surface of the combustion chamber and the outer wall of the ignition nozzle are sealed and fixedly connected; the liquid collecting ring is fixedly connected to the outer side of the wall surface of the combustion chamber, and an annular cavity is provided in the liquid collecting ring. Each ignition nozzle is connected to the annular cavity, and the annular cavity is used to connect to the ignition agent supply pipeline. The direct current ignition structure for the rocket engine combustion chamber provided by the present invention is used to achieve stable ignition of the liquid rocket engine, ensure sufficient ignition of the fuel, and make the engine thrust chamber structure simple and compact.

Description

Technical Field

[0001] The present invention relates to the technical field of rocket engines, and in particular to a direct current ignition structure of a rocket engine combustion chamber. Background Art

[0002] The combustion chamber ignition nozzle is a critical component in the thrust chamber of liquid rocket engines using non-hypergolic propellants, organizing combustion. The ignition nozzle's ignition performance is crucial for stable engine startup, and its performance directly impacts subsequent combustion stability.

[0003] Commonly used ignition structures include torch ignition structure, gunpowder ignition structure, electric ignition structure or plasma ignition structure, but the torch ignition system has problems such as complex structure and limited ignition energy in large-flow thrust chamber, which increases the redundancy of the system and structure for engines that are not started multiple times; and the ignition structures such as gunpowder ignition, electric ignition, and plasma ignition all have the problem of limited ignition energy. For large-thrust, large-flow liquid rocket engines, the combustion chamber diameter is large, and it is impossible to achieve full ignition of the fuel. Summary of the Invention

[0004] The object of the present invention is to provide a direct current ignition structure for a rocket engine combustion chamber, which is used to achieve stable ignition of a liquid rocket engine, ensure sufficient ignition of the fuel, and make the engine thrust chamber structure simple and compact.

[0005] To achieve the above-mentioned object, the present invention provides a direct current ignition structure for a rocket engine combustion chamber, comprising:

[0006] One or more ignition nozzles are evenly arranged along the circumference, and the combustion chamber is provided with one or more ignition nozzles evenly distributed along the circumference, and the wall surface of the combustion chamber is sealed and fixedly connected to the outer wall of the ignition nozzle;

[0007] The liquid collecting ring is fixedly connected to the wall of the combustion chamber. An annular cavity is provided in the liquid collecting ring. Each ignition nozzle is connected to the annular cavity. The annular cavity is used to connect to the ignition agent supply pipeline. The ignition nozzle is used to inject the ignition agent into the combustion chamber.

[0008] Compared with the prior art, an embodiment of the present invention provides a direct-flow ignition structure of a rocket engine combustion chamber, including one or more ignition nozzles and a liquid collecting ring. The combustion chamber is provided with one or more ignition nozzles evenly distributed along the circumference, that is, the ignition nozzles are evenly arranged on the circumferential side of the combustion chamber along the circumference of the combustion chamber. Each ignition nozzle is connected to the annular cavity of the liquid collecting ring. After the ignition agent flows into the annular cavity from the ignition agent supply pipeline, it finally flows to the ignition nozzle and is sprayed into the combustion chamber in different directions through different ignition nozzles, so that the ignition agent can be fully mixed with the incoming propellant in the combustion chamber and spontaneously combust, thereby realizing the ignition of the fuel. Based on this, an embodiment of the present invention provides a direct current ignition structure of a rocket engine combustion chamber, which sprays ignition agent into the combustion chamber through ignition nozzles evenly arranged on the circumferential side of the combustion chamber. Even when the diameter of the combustion chamber is large, the ignition agent can be fully distributed in the combustion chamber, thereby achieving full ignition of the fuel. Compared with the torch ignition structure, there is no need for a separate torch ignition chamber and other structures, so that the structure of the engine thrust chamber is simple and compact. Compared with ignition structures such as gunpowder ignition, electric ignition, and plasma ignition, the ignition agent is sprayed in different directions through one or more ignition nozzles evenly arranged along the circumference, thereby ensuring that the fuel in the combustion chamber is fully ignited, thereby achieving stable ignition of the liquid rocket engine. This structure is also more suitable for large-diameter combustion chambers.

[0009] Optionally, in the above-mentioned direct current ignition structure of the rocket engine combustion chamber, the direct current ignition structure of the rocket engine combustion chamber further includes:

[0010] The ignition agent inlet nozzle is fixedly arranged on the outer wall of the combustion chamber and is arranged between the ignition nozzles. The ignition nozzles are evenly distributed on both sides of the ignition agent inlet nozzle. The outlet of the ignition agent inlet nozzle is connected to the annular cavity. The inlet of the ignition agent inlet nozzle is used to connect with the ignition agent supply pipeline.

[0011] Optionally, in the above-mentioned direct current ignition structure of the rocket engine combustion chamber, a drainage hole and an injection hole are provided on the ignition nozzle, the drainage hole is connected to the annular cavity, and the injection hole is connected to the drainage hole.

[0012] Optionally, in the above-mentioned direct current ignition structure of the rocket engine combustion chamber, the axis of the injection hole has a certain angle with the inner wall of the combustion chamber.

[0013] Optionally, in the above-mentioned direct current ignition structure of the rocket engine combustion chamber, the wall surface of the combustion chamber and the outer wall of the ignition nozzle are sealed and welded;

[0014] And / or, the liquid collecting ring is seal-welded to the wall surface of the combustion chamber. BRIEF DESCRIPTION OF THE DRAWINGS

[0015] The drawings described herein are used to provide a further understanding of the present invention and constitute a part of the present invention. The exemplary embodiments of the present invention and their descriptions are used to explain the present invention and do not constitute an improper limitation of the present invention. In the drawings:

[0016] Figure 1 A cross-sectional view of a straight-flow ignition nozzle structure for a rocket engine combustion chamber provided in an embodiment of the present invention.

[0017] Reference numerals:

[0018] 1-ignition nozzle; 11-drainage hole; 12-injection hole; 2-combustion chamber; 3-liquid collecting ring; 31-annular cavity; 4-ignition agent inlet nozzle. DETAILED DESCRIPTION

[0019] In order to make the technical problems, technical solutions and beneficial effects to be solved by the present invention more clearly understood, the present invention is further described in detail below with reference to the accompanying drawings and embodiments. It should be understood that the specific embodiments described herein are only used to explain the present invention and are not intended to limit the present invention.

[0020] It should be noted that when a part is referred to as being “fixed to” or “disposed on” another part, it can be directly on the other part or indirectly on the other part. When a part is referred to as being “connected to” another part, it can be directly connected to the other part or indirectly connected to the other part.

[0021] Furthermore, the terms "first" and "second" are used for descriptive purposes only and should not be understood to indicate or imply relative importance or implicitly specify the number of the technical features indicated. Thus, a feature specified as "first" or "second" may explicitly or implicitly include one or more of the features. In the description of the present invention, "plurality" means two or more, unless otherwise specifically defined. "Several" means one or more, unless otherwise specifically defined.

[0022] In the description of the present invention, it should be understood that the terms "up", "down", "front", "back", "left", "right", etc., indicating directions or positional relationships, are based on the directions or positional relationships shown in the accompanying drawings, and are only for the convenience of describing the present invention and simplifying the description, rather than indicating or implying that the device or element referred to must have a specific direction, be constructed and operated in a specific direction, and therefore cannot be understood as a limitation on the present invention.

[0023] In the description of the present invention, it should be noted that, unless otherwise expressly specified or limited, the terms "mounted," "connected," and "connected" should be understood in a broad sense. For example, they may refer to fixed, detachable, or integral connections; mechanical or electrical connections; direct or indirect connections through an intermediate medium; and may encompass internal communication between two components or interaction between two components. Those skilled in the art will understand the specific meanings of the above terms in the present invention based on specific circumstances.

[0024] The combustion chamber ignition nozzle is a critical component for organizing combustion in the thrust chamber of liquid rocket engines using non-hypergolic propellants. The ignition performance of the ignition nozzle is crucial for stable engine startup, and its performance directly impacts the stability of subsequent combustion. Common ignition structures include torch, gunpowder, electric, or plasma. However, torch ignition structures suffer from complex construction and limited ignition energy in high-flow thrust chambers. This also increases system and structural redundancy for engines that do not require multiple starts. Gunpowder, electric, and plasma ignition structures all suffer from limited ignition energy. For high-thrust, high-flow liquid rocket engines, the large combustion chamber diameter prevents sufficient fuel ignition.

[0025] In order to solve the above problems, Figure 1 As shown, an embodiment of the present invention provides a direct-flow ignition structure for a rocket engine combustion chamber, comprising one or more ignition nozzles 1 and a liquid collecting ring 3. One or more ignition nozzles 1 are evenly arranged along the circumference of a combustion chamber 2, and the combustion chamber 2 is sealed and fixedly connected to the outer wall of the ignition nozzles 1. The liquid collecting ring 3 is fixedly connected to the combustion chamber 2, and an annular cavity 31 is provided within the liquid collecting ring 3. Each ignition nozzle 1 is connected to the annular cavity 31, and the annular cavity 31 is used to connect to the ignition agent supply pipeline. The ignition nozzles 1 are used to inject the ignition agent into the combustion chamber 2. The number of ignition nozzles 1 is selected based on the diameter of the combustion chamber.

[0026] During the specific working process, the ignition agent in the ignition agent supply pipeline passes into the annular cavity 31 of the liquid collecting ring 3, flows into each ignition nozzle 1 through the annular cavity 31, and is finally injected into the combustion chamber 2 through the ignition nozzle 1. After mixing with the incoming oxidant in the combustion chamber 2, it spontaneously combusts, providing ignition conditions for the incoming fuel, and igniting the incoming fuel.

[0027] It can be seen from the above-mentioned direct current ignition structure and specific processing process that an embodiment of the present invention provides a direct current ignition structure of a rocket engine combustion chamber, which includes one or more ignition nozzles 1 and a liquid collecting ring 3. The combustion chamber 2 is provided with one or more ignition nozzles 1 evenly arranged along the circumference, that is, the ignition nozzles 1 are evenly arranged on the circumferential side of the combustion chamber 2 along the circumference of the combustion chamber 2. Each ignition nozzle 1 is connected to the annular cavity 31 of the liquid collecting ring 3. After the ignition agent flows into the annular cavity 31 from the ignition agent supply pipeline, it finally flows to the ignition nozzle 1 and is sprayed into the combustion chamber 2 in different directions through one or more ignition nozzles 1, so that the ignition agent can contact the incoming propellant in the combustion chamber 2 to spontaneously combust, thereby realizing the ignition of the fuel. Compared with the prior art, the embodiment of the present invention provides a direct current ignition structure of a rocket engine combustion chamber, which sprays ignition agent into the combustion chamber 2 through one or more ignition nozzles 1 evenly arranged on the side of the combustion chamber 2. When the diameter of the combustion chamber 2 is small, one ignition nozzle 1 can be set to ensure stable ignition of the fuel, and the structure is simple; when the diameter of the combustion chamber 2 is large, multiple ignition nozzles 1 can be set, which can distribute multiple ignition agent flames into the combustion chamber 2, thereby achieving stable ignition of the fuel. Compared with the torch ignition structure, there is no need for a separate torch electric ignition chamber and other structures, making the engine thrust chamber structure simple and compact. Compared with ignition structures such as gunpowder ignition, electric ignition, and plasma ignition, the ignition agent is sprayed in different directions through one or more ignition nozzles 1, thereby ensuring that the fuel in the large-diameter combustion chamber 2 is fully ignited, thereby achieving stable ignition of the liquid rocket engine.

[0028] As a possible implementation, the aforementioned direct-flow ignition structure for a rocket engine combustion chamber further includes an ignition agent inlet nozzle 4 fixedly mounted on combustion chamber 2. The outlet of ignition agent inlet nozzle 4 communicates with annular cavity 31, and the inlet of ignition agent inlet nozzle 4 is configured to communicate with an ignition agent supply line. The ignition agent supply line is connected to annular cavity 31 of liquid collecting ring 3 via ignition agent inlet nozzle 4, thereby enabling the circulation of the ignition agent.

[0029] In one alternative embodiment, in the above-described direct-flow ignition structure of the rocket engine combustion chamber, the ignition nozzle 1 is provided with a guide hole 11 and an injection hole 12. The guide hole 11 communicates with the annular cavity 31, and the injection hole 12 communicates with the guide hole 11. The ignition agent passes through the annular cavity 31 and the guide hole 11 and enters the injection hole 12. Finally, it is injected into the combustion chamber 2 through the injection hole 12, achieving smooth flow of the ignition agent.

[0030] As a possible implementation, in the above-mentioned direct current ignition structure of the rocket engine combustion chamber, the axis of the injection hole 12 and the inner wall surface of the combustion chamber 2 have a certain angle; for example, Figure 1As shown, the axis of the injection hole 12 is perpendicular to the inner wall surface of the combustion chamber 2. Such an arrangement ensures the injection depth and uniform distribution of the ignition agent and ensures stable ignition of the combustion chamber 2.

[0031] In some embodiments, in the above-described straight-through ignition structure of the rocket engine combustion chamber, the combustion chamber 2 is welded to the outer wall of the ignition nozzle 1; and / or the liquid collector ring 3 is welded to the outer wall of the combustion chamber 2. For example, the combustion chamber 2 is welded to the outer wall of the ignition nozzle 1, or the liquid collector ring 3 is welded to the outer wall of the combustion chamber 2, or the combustion chamber 2 is welded to the outer wall of the ignition nozzle 1 and the liquid collector ring 3 is welded to the outer wall of the combustion chamber 2. This welding method is simple and reliable, and can ensure the reliability of the connection between the ignition nozzle 1 and the combustion chamber 2, and the liquid collector ring 3 and the combustion chamber 2.

[0032] In the description of the above embodiments, specific features, structures, materials or characteristics may be combined in an appropriate manner in any one or more embodiments or examples.

[0033] The above description is merely a specific embodiment of the present invention, but the scope of protection of the present invention is not limited thereto. Any modifications or substitutions that can be easily conceived by a person skilled in the art within the technical scope disclosed in the present invention should be included in the scope of protection of the present invention. Therefore, the scope of protection of the present invention should be based on the scope of protection of the claims.

Claims

1. A direct current ignition structure for a rocket engine combustion chamber, characterized in that: Applicable to high-thrust, high-flow liquid rocket engines, the direct-flow ignition structure includes: A plurality of ignition nozzles evenly distributed along the circumference are provided on the circumferential side of the combustion chamber, and the wall surface of the combustion chamber is sealed and fixedly connected to the outer wall of the ignition nozzle; a liquid collecting ring, the liquid collecting ring being fixedly connected to the outer wall of the combustion chamber, an annular cavity being provided in the liquid collecting ring, each of the ignition nozzles being in communication with the annular cavity, the annular cavity being configured to be in communication with an ignition agent supply line, the ignition nozzles being configured to inject ignition agent into the combustion chamber so that the ignition agent spontaneously ignites upon contact with the incoming oxidant, and multiple ignition agent flames are evenly distributed within the combustion chamber, thereby achieving stable ignition of the fuel and, in turn, stable ignition of the liquid rocket engine; The ignition nozzle is provided with a drainage hole and an injection hole. The drainage hole is communicated with the annular cavity, and the injection hole is communicated with the drainage hole. The diameter of the drainage hole is larger than the diameter of the injection hole.

2. The direct current ignition structure of the rocket engine combustion chamber according to claim 1, characterized in that: The direct current ignition structure of the rocket engine combustion chamber also includes: An ignition agent inlet nozzle is fixedly arranged on the liquid collecting ring, an outlet of the ignition agent inlet nozzle is communicated with the annular cavity, and an inlet of the ignition agent inlet nozzle is used to communicate with the ignition agent supply pipeline.

3. The direct current ignition structure of the rocket engine combustion chamber according to claim 1, characterized in that: The axis of the injection hole is perpendicular to the inner wall surface of the combustion chamber.

4. The direct current ignition structure of a rocket engine combustion chamber according to claim 1, characterized in that: The wall surface of the combustion chamber and the outer wall of the ignition nozzle are sealed and welded; And / or, the liquid collecting ring is sealed and welded to the wall surface of the combustion chamber.

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