Spontaneous combustion propellant space detonation engine injector

Abstract

The invention provides a spontaneous combustion propellant space detonation engine injector, which belongs to the technical field of spacecraft propulsion systems and comprises an injection surface and an injection structure. The injection surface bears an injection structure, and the injection structure is formed by three annularly arranged injection rings with radiuses gradually increased from inside to outside; aiming at the urgent demand of a space engine with more advanced performance indexes, a novel space detonation engine design configuration is provided, and detonation wave stable propagation of a spontaneous combustion propellant for a long time can be realized; through cooperative cooperation of multiple circles of injection holes and matching with a damping structure, the anti-quenching capacity is greatly improved compared with that of an annular detonation combustion chamber, stable propagation of detonation waves is achieved, and the duration time is larger than or equal to 5 s; the combustion efficiency in the detonation mode can be larger than or equal to 99.5% (the combustion efficiency of a conventional injector is larger than or equal to 95%), and the combustion efficiency of a space engine can be further improved compared with that of an existing space engine.

Description

TECHNICAL FIELD

[0001] The present application belongs to the technical field of spacecraft propulsion system, and particularly relates to a self-ignition propellant space detonation engine injector. BACKGROUND

[0002] High-performance space engines are an important support for future national defense equipment competitiveness and national defense capability, and advanced performance indicators are urgently needed for space engines to achieve the goal of a strong aerospace nation. The combustion efficiency of space engines using constant-pressure thermodynamic cycles has approached the upper limit, and it is necessary to fundamentally change the thermodynamic cycle of space engines to improve their performance. Detonation combustion can enable space engines to achieve approximately constant-volume thermodynamic cycle efficiency, thereby improving the combustion efficiency of space engines.

[0003] Considering thrust, specific impulse, on-orbit conditions, and operating temperature, self-ignition propellants are generally used as fuel and oxidizers for high-performance space engines. However, in self-ignition propellant detonation combustion, slow-burning interference can easily cause premature ignition and detonation extinction. Currently, annular detonation chambers are commonly used in engineering, and the boundary quenching effect leads to unstable circumferential propagation of detonation waves.

[0004] Currently, the annular detonation chamber of the self-ignition propellant space detonation engine injector can achieve detonation initiation and maintenance, but the short maintenance time is caused by slow-burning interference and unstable detonation propagation. A new space detonation engine design configuration is needed to achieve stable propagation of self-ignition propellant detonation waves for a long time.

[0005] Currently, related prior art includes: A combined combustion stabilizing device (patent document CN110805506A) is a combined combustion stabilizing device, which includes an injector base, an oxidizer top cover, an acoustic cavity ring, a partition nozzle, and a common nozzle. The injector base and the oxidizer top cover are welded together, and the upper surface of the injector base is a V-shaped stepped structure. The injector base and the oxidizer top cover form an oxidizer head cavity. The acoustic cavity ring is welded to the outer edge of the lower surface of the injector base, and the acoustic cavity ring and the injector base form a fuel head cavity. The partition nozzle and the V-shaped stepped structure of the injector base are connected by threads + brazing, and the partition nozzle extends a certain distance from the lower surface of the injector base. On the V-shaped stepped structure of the injector base, a plurality of common nozzles are installed with the center of the V-shaped stepped structure as the center. However, the device described in this patent document cannot achieve stable detonation combustion and can only achieve stable slow-burning. SUMMARY

[0006] In view of the defects in the prior art, the purpose of the present application is to provide a self-ignition propellant space detonation engine injector.

[0007] The self-ignition propellant space detonation engine injector provided by the application comprises a jet surface and a jet structure. The jet surface 1 bears the jet structure and comprises a first jet ring 11, a second jet ring 12 and a third jet ring 13 arranged in sequence from inside to outside, which are three rings of jet rings arranged in sequence with increasing radii. The flow strength ratio of the first jet ring, the second jet ring and the third jet ring is 5:7:4.

[0008] Preferably, the first jet ring is located at a radius R a1 =0.15R, R is the radius of the combustion chamber, and the propellant flow strength accounts for 31.25%.

[0009] Preferably, the second jet ring is located at a radius R a2 =0.5R, and the propellant flow strength accounts for 43.75%.

[0010] Preferably, the third jet ring is located at a radius R a3 =0.85R, and the propellant flow strength accounts for 25%.

[0011] Preferably, an acoustic cavity is further arranged on the jet surface. The acoustic cavity is arranged around the edge of the jet surface 1, and the depth of the acoustic cavity meets the Helmholtz resonance requirement.

[0012] Preferably, the depth h of the acoustic cavity and the frequency f D of the detonation wave satisfy:

[0013] Wherein, v c is the sound speed of the gas in the acoustic cavity hole.

[0014] Preferably, the impedance of the acoustic cavity is matched with the gain effect of the first jet ring, so that the intensity fluctuation of the detonation wave is ≤15%.

[0015] Preferably, each jet ring of the jet structure is composed of a pair of mutually hitting jet holes, and each jet ring has 6 to 96 jet hole pairs; each jet hole pair comprises an oxidizer jet hole and a fuel jet hole, and the oxidizer jet hole and the fuel jet hole are both inclined jet pairs, and the included angle between the central axis of the oxidizer jet hole and the vertical line is 1、 2=30° to 60°.

[0016] Preferably, the distance from the intersection point of the central axes of the oxidizer jet hole and the fuel jet hole to the center of the jet surface is R x , and the radius of the combustion chamber is R, which is set according to requirements. .

[0017] Preferably, the diameters of the mutually hitting jet holes satisfy:

[0018]

[0019] wherein d o , d f is the oxidizer, fuel aperture; q o , q f is the oxidizer, fuel density; n o , n f is the oxidizer, fuel number of injection holes; v o , v f is the oxidizer, fuel flow rate; o , f is the oxidizer, fuel density.

[0020] Compared with the prior art, the present application has the following beneficial effects: 1. The present application greatly improves the anti-quenching ability compared with the annular detonation combustion chamber through the synergistic cooperation of the multiple circle injection holes and the matching with the damping structure, realizes the stable propagation of the detonation wave, and the duration is greater than or equal to 5s; 2. The present application can realize the combustion efficiency greater than or equal to 99.5% in the detonation mode (the conventional injector is greater than or equal to 95%) for the self-ignition propellant space detonation engine injector, so that the space engine is improved compared with the existing combustion efficiency. BRIEF DESCRIPTION OF DRAWINGS

[0021] Other features, objects and advantages of the present application will become more apparent from the following detailed description of non-limiting embodiments with reference to the attached drawings: Fig. 1 is the axial schematic view of the self-ignition propellant space detonation engine injector of the present application; Fig. 2 is the angle schematic view of the injection structure of the self-ignition propellant space detonation engine injector of the present application.

[0022] The drawings show: DETAILED DESCRIPTION

[0023] The present application will be described in detail below with reference to specific embodiments. The following embodiments will help those skilled in the art to further understand the present application, but do not limit the present application in any form. It should be pointed out that those skilled in the art can make several changes and improvements without departing from the concept of the present application. These all belong to the protection scope of the present application.

[0024] As Figs. 1-2As shown, the embodiment of the present application provides a self-ignition propellant space detonation engine injector, comprising: an injection surface 1, an injection structure; The injection surface 1 carries the injection structure, comprising: a first injection ring 11, a second injection ring 12, and a third injection ring 13 arranged in sequence from inside to outside, which are three rings of injection rings with increasing radii; The flow intensity ratio of the first injection ring 11, the second injection ring 12, and the third injection ring 13 is 5:7:4.

[0025] The first injection ring 11 is located at a radius R a1 0.15R, R is the radius of the combustion chamber, and the propellant flow intensity accounts for 31.25%, which is used to organize the center area combustion of the injector and realize the gain of the detonation wave maintenance in the center area of the injector; Specifically, the center area combustion of the injector is organized, and a stable high-temperature fuel backflow area is formed under the interaction of the detonation wave, the combustion intensity is enhanced, and the stability of the detonation wave is improved; The second injection ring 12 is located at a radius R a2 0.5R, the propellant flow intensity accounts for 43.75%, which is used to maintain the circumferential propagation of the detonation wave; Specifically, the propellant required for maintaining the detonation wave front reaction area maintains the detonation wave pressure peak value at an intensity of ≥3MPa; The third injection ring 13 is located at a radius R a3 0.85R, the propellant flow intensity accounts for 25%, which is used to isolate the detonation wave and the combustion chamber wall; Specifically, by changing the mixing ratio between the oxidizer and the fuel, the combustion intensity is reduced, and a thermal boundary layer isolation area is constructed to isolate the detonation wave and the combustion chamber wall, preventing the thermal flow overload of the combustion chamber wall; Further, the injection surface 1 is further provided with an acoustic cavity 14: the acoustic cavity 14 is arranged at the edge of the injection surface 1, the depth of the acoustic cavity 14 meets the Helmholtz resonance requirement, and the energy of the detonation wave frequency is attenuated; the impedance of the acoustic cavity 14 is matched with the gain effect of the first injection ring 11, so that the intensity fluctuation of the detonation wave is ≤15%; Further, the depth h of the acoustic cavity 14 and the frequency f D meet:

[0026] Wherein, v c is the sound speed of the gas in the acoustic cavity hole.

[0027] The oscillation is attenuated by the resonance of the gas in the acoustic cavity 14, and the 3000-4500Hz high-frequency pressure oscillation is absorbed.

[0028] Furthermore, each injection ring of the injection structure is composed of pairs of mutually striking injection holes, with 6 to 96 injection pairs per ring; each injection pair includes: an oxidizer injection hole 121 and a fuel injection hole 122, wherein both the oxidizer injection hole 121 and the fuel injection hole 122 are inclined injection pairs, such as... Fig. 2 As shown, the angle between their central axes and perpendicular lines is respectively 1. 2 = 30° to 60°; the radius of the intersection of the central axes of the injection holes from the center of the injection surface 1 is R. x The combustion chamber radius is R, which is set as needed. ; Furthermore, the mutual impact injection orifice diameter satisfies:

[0029]

[0030] Where d o d f For oxidant and fuel pore size; q o q f For oxidant, fuel density; n o n f The number of oxidizer and fuel injection holes; o , f Oxidant, fuel flow rate; o , f Oxidizing agent, fuel density; Furthermore, in a more specific embodiment, the first injection ring 11 has 50 injection pairs, the second injection ring 12 has 70 injection pairs, and the third injection ring 13 has 40 injection pairs. Both the oxidizer injection hole 121 and the fuel injection hole 122 are inclined injection pairs, with their central axes forming angles with the vertical. 1. 2=45°; the radius of the intersection of the central axes of the injection holes from the center of the injection surface 1 is R. x The combustion chamber radius is R, and the first injection ring 11, the second injection ring 12, and the third injection ring 13 have a radius of R. x The values ​​are 0.15R, 0.5R, and 0.85R respectively; the mutual impact injection orifice diameters satisfy:

[0031]

[0032] Where d o d f For oxidant and fuel pore size; q o qf is the oxidizer, fuel density; n o is the oxidizer, fuel injection hole number; v f is the oxidizer, fuel injection hole number; v o is the oxidizer, fuel injection hole number; v f is the oxidizer, fuel flow rate; o , f is the oxidizer, fuel density; cavity 14 depth h and the blast wave frequency f D satisfies:

[0033] wherein v c is the sound speed of the gas in the cavity hole, used to absorb high frequency pressure oscillation of 3000 to 4500 Hz.

[0034] In summary, the embodiment of the present application provides a self-ignition propellant space detonation engine injector, aiming at the urgent demand of more advanced performance index space engine, a new space detonation engine design configuration is proposed, which can realize stable propagation of detonation wave of self-ignition propellant for a long time; through the cooperation of multiple circle injection holes and the matching with the damping structure, the quenching resistance is greatly improved compared with the annular detonation combustion chamber, the stable propagation of detonation wave is realized, and the duration is ≥5s; the combustion efficiency in detonation mode can be ≥99.5% (conventional injector ≥95%), so that the space engine can be further improved compared with the existing combustion efficiency.

[0035] In the description of the present application, it should be understood that the terms "upper", "lower", "front", "back", "left", "right", "vertical", "horizontal", "top", "bottom", "inner", "outer" and the like indicate the orientation or positional relationship based on the orientation or positional relationship shown in the drawings, and are only for the convenience of describing the present application and simplifying the description, and do not indicate or imply that the device or element referred to must have a particular orientation, be constructed and operated in a particular orientation, therefore it cannot be understood as a limitation to the present application.

[0036] The specific embodiments of the present application are described above. It should be understood that the present application is not limited to the above specific embodiments, and those skilled in the art can make various changes or modifications within the scope of the claims, which does not affect the essential content of the present application. In the case of no conflict, the embodiments of the present application and the features in the embodiments can be combined with each other arbitrarily.

Claims

1. A self-igniting propellant space detonation engine injector characterized by, Comprise: Injection surface (1), injection structure; The injection surface (1) carries the injection structure, comprising: the first injection ring (11), the second injection ring (12) and the third injection ring (13) are sequentially annularly arranged, which are three annularly arranged injection rings with increasing radii; The flow intensity ratio of the first injection ring (11), the second injection ring (12) and the third injection ring (13) is 5:7:

4.

2. The self-ignition propellant space detonation engine injector according to claim 1, wherein, The first injection circle (11) is located at a radius R a1 =0.15R, R is the radius of the combustion chamber, and the propellant flow intensity accounts for 31.25%.

3. The self-ignition propellant space detonation engine injector according to claim 1, wherein, The second injection circle (12) is located at a radius R a2 =0.5R, and the propellant flow intensity accounts for 43.75%.

4. The self-ignition propellant space detonation engine injector according to claim 1, wherein, The third injection circle (13) is located at a radius R a3 = 0.85R, and the propellant flow intensity accounts for 25%.

5. The self-ignition propellant space detonation engine injector according to claim 2, wherein, The injection surface (1) is further provided with an acoustic cavity (14); The acoustic cavity (14) is annularly arranged at the edge of the injection surface 1, and the depth of the acoustic cavity (14) meets the Helmholtz resonance requirement.

6. The self-ignition propellant space detonation engine injector according to claim 5, wherein, The depth h of the acoustic cavity (14) and the frequency f of the detonation wave D satisfies: where v c is the speed of sound of the gas in the bore of the acoustic cavity.

7. The self-igniting propellant space detonation engine injector of claim 5 wherein, The impedance of the acoustic cavity (14) is matched with the gain effect of the first injection ring (11), so that the detonation wave intensity fluctuation is ≤15%.

8. The self-igniting propellant space detonation engine injector of claim 1 wherein, Each injection circle of the injection structure is composed of pairs of intersecting injection holes, 6 to 96 injection hole pairs per circle; each injection hole pair comprises an oxidizer injection hole (121) and a fuel injection hole (122), both being inclined injection pairs, the included angle between the central axis of the oxidizer injection hole (121) and the vertical line and the included angle between the central axis of the fuel injection hole (122) and the vertical line are respectively 1、 2=30° to 60°.

9. The self-igniting propellant space detonation engine injector of claim 8, wherein, The intersection distance between the center axis of the oxidizer injection hole (121) and the fuel injection hole (122) and the center of the injection surface (1) is a radius R x , and the radius of the combustion chamber is R, which is set as needed .

10. The self-igniting propellant space detonation engine injector of claim 8, wherein, The mutual impact injection aperture satisfies: where d o , d f is oxidizer, fuel pore size; q o , q f is oxidizer, fuel density; n o , n f is oxidizer, fuel number of injection holes; v o , v f is oxidizer, fuel flow rate; o , f is oxidizer, fuel density.

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