Plasma torch ignition device of rocket engine and rocket engine

Abstract

The invention discloses a plasma torch ignition device of a rocket engine and the rocket engine. The problems that in an ignition mode of an existing rocket engine, the replacement cost is high due to the fact that explosive columns need to be manually filled for ignition of explosive columns of initiating explosive devices, ignition failure is easily caused by spark plug type ignition, and a torch type igniter is large in size and poor in ignition reliability are solved. The ignition device is suitable for torch ignition under the non-premixing condition, a propellant channel is formed between the cathode and the anode in the upper cavity, a propellant in the propellant channel is ionized through a high-voltage electric field formed between the cathode and the anode, and high-temperature particles are generated; and the fuel gas enters the lower cavity and is mixed with another propellant to form high-temperature fuel gas, so that the fuel gas has the advantage of high stability.

Description

Technical Field

[0001] The present invention relates to a rocket engine ignition device, and particularly to a plasma torch ignition device for a rocket engine and a rocket engine. Background Art

[0002] Plasma ignition technology forms a large number of active particles by electrolyzing the working medium between electrodes through high-energy pulses, which can penetrate the working medium, generating a high-temperature and high-speed plasma jet, and can be used as an initial fire core to achieve rapid and stable ignition of the combustion chamber. As a new type of ignition technology, plasma ignition has many advantages over traditional electrical ignition: a large ignition area, high ignition energy, better coupling of ignition energy with the fuel-air mixture, short ignition delay time, high ignition success rate, etc.

[0003] In the field of rocket engines, plasma ignition technology is currently also in the research stage and is less commonly used at present. For liquid rocket engines, the commonly used ignition methods mainly include pyrotechnic grain ignition, spark plug ignition, and torch ignition, etc. Pyrotechnic grain ignition belongs to a one-time ignition strategy, and manual loading and replacement are required each time it is used, with a relatively high replacement cost, which is not conducive to the requirement of repeated use; the ignition energy generated by spark plug ignition is relatively low, and at the same time, as the vacuum degree decreases, the Joule energy generated by the spark plug will further decrease, which may lead to ignition failure; the essence of a torch igniter is the secondary amplification of the combustion flame. By introducing a small amount of propellant, a relatively small-power spark plug can be used to form a stable high-temperature flame. The reaction medium has a relatively high combustion temperature and ignition flow rate, and can achieve repeated ignition and start-up of engines with a relatively large flow rate. However, a conventional torch igniter generally has a relatively large volume, and its ignition reliability is affected by the performance of the spark plug. Summary of the Invention

[0004] The purpose of the present invention is to solve the problems existing in the ignition methods of existing rocket engines, such as the high replacement cost caused by manual loading of pyrotechnic grains in pyrotechnic grain ignition, the easy occurrence of ignition failure in spark plug ignition, and the large volume and poor ignition reliability of torch igniters, and to provide a plasma torch ignition device for a rocket engine and a rocket engine.

[0005] To achieve the above purpose, the technical solution adopted by the present invention is as follows:

[0006] A plasma torch ignition device for a rocket engine, characterized in that it includes an igniter housing with openings at both ends, a cathode plate disposed in the igniter housing and dividing the inner cavity of the igniter housing into an upper cavity and a lower cavity, an anode coaxially disposed in the upper cavity, an insulating layer sleeved outside the anode, and a plug connector hermetically connected to one end of the igniter housing located in the upper cavity and electrically connected to the anode for connecting an external power supply;

[0007] A propellant channel that communicates with each other is formed between the inner wall of the igniter housing and the outer wall of the insulating layer, and between the bottom of the anode and the cathode plate. A plurality of first propellant inlet holes communicating with the propellant channel are provided on the igniter housing on one side of the upper cavity.

[0008] A high-temperature particle channel communicating with the upper cavity and the lower cavity is provided in the middle of the cathode plate.

[0009] A plurality of second propellant inlet holes communicating with the lower cavity are provided on the igniter housing on one side of the lower cavity, and one end of the igniter housing body located in the lower cavity serves as an ignition output end.

[0010] Further, the end of the anode away from the plug connector is conical.

[0011] Further, a plurality of anode cooling holes communicating with the gas fuel channel are provided on the insulating layer near the conical end of the anode, and the anode cooling holes are arranged opposite to the conical inclined surface.

[0012] Further, the igniter housing is a cylindrical housing body, and the second propellant inlet holes are provided along the tangential direction of the inner wall of the igniter housing.

[0013] Further, the input voltage of the plug connector is 70 kV to 100 kV.

[0014] Further, the distance between the anode and the cathode plate is 1 mm to 3 mm.

[0015] Further, the gap distance between the inner wall of the igniter housing and the outer wall of the insulating layer is 0.5 mm to 1 mm.

[0016] Further, the number of the first propellant inlet holes and the second propellant inlet holes is 3 to 6, and they are evenly distributed along the circumferential direction on the outer side wall of the igniter housing.

[0017] Further, the aperture diameters of the first propellant inlet holes and the second propellant inlet holes are both 0.5 mm to 0.7 mm;

[0018] The aperture diameter of the anode cooling holes is 0.3 mm to 0.4 mm,

[0019] The aperture diameter of the high-temperature particle channel is 0.8 mm to 1 mm.

[0020] A rocket engine is characterized in that it includes a rocket engine body and the above-mentioned rocket engine plasma torch ignition device;

[0021] The ignition output end of the rocket engine plasma torch ignition device is fixedly connected to the combustion chamber of the rocket engine body.

[0022] The beneficial effects of the present invention are:

[0023] 1. A plasma torch ignition device for a rocket engine provided by the present invention is applicable to torch ignition under non-premixed conditions. By forming a propellant channel between the cathode and the anode in the upper cavity, the propellant in the propellant channel is ionized by the high-voltage electric field formed between the cathode and the anode to generate high-temperature particles, which enter the lower cavity to be mixed with another propellant and form high-temperature gas.

[0024] 2. For a plasma torch ignition device for a rocket engine provided by the present invention, the temperature near the anode is relatively high during operation. The propellant is shunted through the anode cooling holes, and a cooling film is formed at the end of the anode, greatly delaying the ablation rate of the anode, and thus extending the service life of the plasma torch ignition device for the rocket engine.

[0025] 3. A plasma torch ignition device for a rocket engine provided by the present invention has a simple structure and can be compatible with a variety of propellants, including liquid (gaseous) oxygen / kerosene, liquid (gaseous) oxygen / methane, etc. The mixing ratio is flexibly adjustable. During use, the propellants in the first propellant inlet hole and the second propellant inlet hole can be swapped according to actual needs, and the ignition reliability and stability are high.

[0026] 4. A rocket engine provided by the present invention has high ignition reliability and stability. Description of the Drawings

[0027] Figure 1 is a schematic structural diagram of an embodiment of a plasma torch ignition device for a rocket engine of the present invention.

[0028] In the figure, 1 - plug connector; 2 - igniter housing; 3 - anode; 4 - insulating layer; 5 - cathode plate; 6 - high-temperature particle channel; 7 - anode cooling hole; 8 - first propellant inlet hole; 9 - second propellant inlet hole; 10 - propellant channel. Detailed Embodiments

[0029] To make the objectives, advantages, and features of the present invention clearer, the following further describes in detail a plasma torch ignition device for a rocket engine and a rocket engine proposed by the present invention with reference to the accompanying drawings and specific embodiments. According to the following detailed embodiments, the advantages and features of the present invention will be clearer.

[0030] In response to the current repeated and reliable ignition requirements of rocket engines, this embodiment provides a plasma torch ignition device for a rocket engine, which can be used for multiple propellants such as liquid (gaseous) oxygen / kerosene, liquid (gaseous) oxygen / methane, etc. It has a simple structure, a flexibly adjustable mixing ratio, the media at the inlet can be swapped, it can be reused multiple times, and the ignition reliability is high. At the same time, through structural design, the convective heat transfer of the propellant is utilized to achieve the cooling film of the anode, which can greatly delay the ablation rate of the anode and extend the service life of the igniter.

[0031] As Figure 1 shown, this embodiment provides a plasma torch ignition device for a rocket engine, which includes an igniter housing 2, a cathode plate 5, an anode 3, an insulating layer 4, and a plug connector 1.

[0032] The cathode plate 5 is connected inside the igniter housing 2 and divides the igniter housing 2 into an upper cavity and a lower cavity. The plug connector 1 is hermetically connected to the end of the igniter housing 2 in the upper cavity. The plug connector 1 is electrically connected to the anode 3, and the anode 3 is sleeved with the insulating layer 4; a propellant passage 10 is formed between the inner wall of the igniter housing 2 and the outer wall of the insulating layer 4 and between the bottom of the anode 3 and the cathode plate 5. A plurality of first propellant inlet holes 8 communicating with the propellant passage 10 are provided on one side of the igniter housing 2 in the upper cavity; a high-temperature particle passage 6 is provided in the middle of the cathode plate 5; the high-temperature particle passage 6 communicates the upper cavity and the lower cavity; a plurality of second propellant inlet holes 9 are provided on one side of the igniter housing 2 in the lower cavity, and one end of the igniter housing body 2 in the lower cavity serves as an ignition output end.

[0033] The plug connector 1 is connected to a 28V DC power supply externally connected by a high-voltage package. The high-voltage package converts the 28V DC power into 90kV high voltage, and a high-voltage electric field is formed between the bottom of the anode 3 and the cathode plate 5 with a distance of 1mm; Propellant A enters the upper cavity through 4 first propellant inlet holes 8 with a diameter of 0.5mm, flows in the propellant passage 10 to the bottom of the cathode plate 5 and the anode 3, and the propellant A is ionized into high-temperature active particles by the high-voltage electric field, and then enters the lower cavity through the high-temperature particle passage 6 with a diameter of 1mm, and is mixed with propellant B entering the lower cavity through 4 second propellant inlet holes 9 with a diameter of 0.7mm to form high-temperature gas, and the rocket engine combustion chamber is ignited through the ignition output end.

[0034] One end of the anode 3 away from the plug connector 1 is conical, which can form a strong local electric field, can quickly ionize the propellant A, release high temperature through local ionization, and improve the ionization efficiency.

[0035] Due to the high temperature generated during ionization, the ablation rate of the anode is relatively fast. To solve this problem, in this embodiment, a plurality of anode cooling holes 7 communicating with the propellant passage 10 are provided on the side wall of the insulating layer 4 near the conical end of the anode 3, and the anode cooling holes 7 are opposite to the conical inclined surface. When the propellant A flows in the propellant passage 10, a part enters the anode cooling holes 7 and forms a cooling film on the surface of the anode 3, which can cool the end of the anode, greatly delay the ablation rate of the anode 3, and thus extend the service life of the ignition device.

[0036] In this embodiment, for the cylindrical housing of the igniter housing 2, the second propellant inlet hole 9 is opened along the tangential direction of the inner wall of the igniter housing 2. The propellant B enters the lower cavity through the second propellant inlet hole 9, and the mixing of the propellant B and the high-temperature active particles is completed under the swirling flow action, making the mixing more uniform and increasing the output area after mixing, so as to realize the rapid ignition of the rocket engine.

[0037] Among them, the propellant A and the propellant B can adopt various propellant combinations such as liquid (gaseous) oxygen / kerosene, liquid (gaseous) oxygen / methane, etc. In the actual use process, the propellant types of the first propellant inlet hole 8 and the second propellant inlet hole 9 can be selected according to needs, and the number and size of the first propellant inlet hole 8, the second propellant inlet hole 9, and the anode cooling hole 7 can be adjusted according to the ignition flow demand to ensure successful ignition under the condition of reliable cooling.

[0038] The ignition output end of the plasma torch ignition device for a rocket engine provided in this embodiment is fixedly connected to the combustion chamber of the rocket engine body to form a new type of rocket engine, and the rocket engine in this embodiment has been verified for ignition. In the ignition test, gaseous oxygen / kerosene is used as the propellant, the total flow rate of the igniter is 30 g / s, the mixing ratio is 0.2, the actual outlet gas temperature of the ignition output end of the torch ignition device is 873 K, and the length of the formed gas flame is about 1.5 m. It has been successfully applied to the ignition of a certain type of liquid oxygen / kerosene thrust chamber and has high ignition reliability.

[0039] The above embodiments are only used to illustrate the technical solutions of the present application, rather than to limit the present application; although the present application has been described in detail with reference to the foregoing embodiments, those of ordinary skill in the art should understand that they can still modify the technical solutions recorded in the foregoing embodiments, or perform equivalent replacements on some or all of the technical features; and these modifications or replacements do not make the essence of the corresponding technical solutions deviate from the scope of the technical solutions of the present application.

Claims

1. A plasma torch ignition device for a rocket engine, characterized in that: It includes an igniter housing (2) with openings at both ends, a cathode plate (5) disposed within the igniter housing (2) and dividing the inner cavity of the igniter housing (2) into an upper cavity and a lower cavity, an anode (3) coaxially arranged in the upper cavity, an insulating layer (4) sleeved outside the anode (3), and a plug connector (1) hermetically connected to one end of the igniter housing (2) located in the upper cavity and electrically connected to the anode (3) for connecting an external power source; A propellant passage (10) is formed between the inner wall of the igniter housing (2) and the outer wall of the insulating layer (4) and between the bottom of the anode (3) and the cathode plate (5). A plurality of first propellant inlet holes (8) communicating with the propellant passage (10) are provided on one side of the igniter housing (2) located in the upper cavity; A high-temperature particle passage (6) communicating with the upper cavity and the lower cavity is provided in the middle of the cathode plate (5); A plurality of second propellant inlet holes (9) communicating with the lower cavity are provided on one side of the igniter housing (2) located in the lower cavity. One end of the igniter housing body (2) located in the lower cavity serves as an ignition output end.

2. The plasma torch ignition device for a rocket engine according to claim 1, characterized in that: One end of the anode (3) away from the plug connector (1) is conical.

3. The plasma torch ignition device for a rocket engine according to claim 2, characterized in that: A plurality of anode cooling holes (7) communicating with the propellant passage (10) are provided on the side wall of the insulating layer (4) near the conical end of the anode (3), and the anode cooling holes (7) are arranged opposite to the conical slope.

4. The plasma torch ignition device for a rocket engine according to claim 1, characterized in that: The igniter housing (2) is a cylindrical housing body, and the second propellant inlet holes (9) are provided along the tangential direction of the inner wall of the igniter housing (2).

5. The plasma torch ignition device for a rocket engine according to claim 1, characterized in that: The input voltage of the plug connector (1) is 70 kV to 100 kV.

6. The plasma torch ignition device for a rocket engine according to claim 1, characterized in that: The distance between the anode (3) and the cathode plate (5) is 1 mm to 3 mm.

7. The plasma torch ignition device for a rocket engine according to claim 1, wherein: The clearance distance between the inner wall of the igniter housing (2) and the outer wall of the insulating layer (4) is 0.5 mm to 1 mm.

8. The plasma torch ignition device for a rocket engine according to claim 1, characterized in that: The number of the first propellant inlet holes (8) and the second propellant inlet holes (9) is 3 to 6, and they are evenly distributed along the circumferential direction on the outer side wall of the igniter housing (2).

9. The plasma torch ignition device for a rocket engine according to claim 3, characterized in that: The aperture diameters of the first propellant inlet holes (8) and the second propellant inlet holes (9) are both 0.5 mm to 0.7 mm; The aperture diameter of the anode cooling holes (7) is 0.3 mm to 0.4 mm; The aperture diameter of the high-temperature particle passage (6) is 0.8 mm to 1 mm.

10. A rocket engine, characterized in that: It includes a rocket engine body and the rocket engine plasma torch ignition device according to any one of claims 1 to 9; The ignition output end of the rocket engine plasma torch ignition device is fixedly connected to the combustion chamber of the rocket engine body.

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