Rocket engine injector pressure compensation device and method, propulsion system and rocket

Abstract

The invention relates to the technical field of spacecraft propulsion systems, and provides a pressure compensation device and method for a rocket engine injector, a propulsion system and a rocket. The pressure compensation device for the rocket engine injector comprises an equal flow channel, a pressure storage cavity and a buffer layer; the two ends of the equal flow channel are connected with a propellant upstream inlet and the head containing cavity respectively, the pressure storage cavity is located in the head containing cavity and tightly attached to the inner wall face of the head containing cavity, and the buffer layer is arranged at an inlet of the injection hole. The problems of low-temperature propellant vaporization phase change and the like are systematically solved, the atomization and combustion stability of the injector is improved, and normal output of engine thrust is maintained.

Description

Technical Field

[0001] This invention relates to the field of spacecraft propulsion system technology, specifically to a rocket engine injector pressure compensation device, method, propulsion system, and rocket. Background Technology

[0002] Propulsion systems employing cryogenic propellant dual-element engines offer numerous advantages, including simplified structure, wide adaptability, and ease of use and maintenance, meeting the operational needs of propulsion systems across all regions and seasons. The use of cryogenic propellants provides a technological possibility for reducing thermal control, power consumption, size, and weight, while improving structural efficiency.

[0003] Cryogenic propellants have low boiling points and readily vaporize in engine piping, forming a two-phase flow. This leads to increased flow resistance, reduced and unstable flow rate, causing abnormal engine thrust output and even structural ablation. Therefore, preventing two-phase flow and achieving stable flow rate in engine piping is crucial for cryogenic propellant engine design. In engineering practice, increasing engine operating pressure is commonly used to reduce propellant vaporization and its effects.

[0004] Although existing measures can suppress the vaporization of cryogenic propellants to some extent, narrow flow channels and sharp angle areas in the injector structure can still vaporize and generate bubbles, posing a risk of unstable engine flow and abnormal operation. It is difficult to systematically solve the problem of phase change during the vaporization of cryogenic propellants. Summary of the Invention

[0005] To address the shortcomings of existing technologies, the present invention aims to provide a rocket engine injector pressure compensation device, method, propulsion system, and rocket.

[0006] A rocket engine injector pressure compensation device according to the present invention includes a constant current channel, a pressure accumulator chamber, and a buffer layer. One end of the constant DC channel is connected to the upstream inlet of the propellant, and the other end of the constant DC channel is connected to the head cavity. The accumulator is located inside the head cavity and is arranged close to the inner wall of the head cavity. The buffer layer is configured at the inlet of the injection hole to form a propellant injection unit. The inner surface of the constant DC channel is coated with a composite coating.

[0007] According to the present invention, a rocket engine injector pressure compensation method is provided, wherein the two ends of a constant current channel coated with a composite coating are respectively connected to the upstream inlet of the propellant and the head cavity, so that cavitation bubbles can quickly detach from the wall and be carried away by the mainstream. A pressure accumulator is arranged inside the head cavity and closely attached to the inner wall of the head cavity, and the pressure fluctuation is absorbed by the expansion of the pressure accumulator. At the same time, a buffer layer is arranged at the injection hole inlet.

[0008] Preferably, the composite coating is made of fluorinated silane-graphene composite coating with a contact angle >160°.

[0009] Preferably, the pressure accumulator is a cavity formed by an elastic metal bellows; The diameter of the metal bellows is smaller than the inner diameter of the propellant head cavity, and the metal bellows are stacked.

[0010] Preferably, the accumulator chamber includes multiple turns of bellows, with the number of turns being 5 to 15.

[0011] Preferably, the metal bellows is made of nickel-titanium alloy with a bulk modulus of less than 5 MPa. When the vapor pressure of MON-25 rises sharply, the bellows expands to absorb the pressure fluctuation, maintaining the downstream pressure of the incompressible liquid working fluid at more than 20% higher than the saturated vapor pressure of the propellant.

[0012] Preferably, the buffer layer is made of porous metal foam.

[0013] Preferably, the buffer layer is made of aluminum foam with a porosity of 80% and a pore size of 50-100μm, which stabilizes the continuity of the liquid phase through capillary force and reduces the amplitude of flow rate changes caused by bubble rupture.

[0014] According to the present invention, a rocket engine propulsion system employs the aforementioned rocket engine injector pressure compensation device.

[0015] A rocket according to the present invention includes the aforementioned rocket engine propulsion system.

[0016] Compared with the prior art, the present invention has the following beneficial effects: In this invention, a composite coating is sprayed onto a DC channel connecting the upstream inlet of the propellant and the head cavity, allowing cavitation bubbles to quickly detach from the wall and be carried away by the mainstream. A metal bellows accumulator chamber is located inside the propellant head cavity, closely attached to the inner wall. The bellows expands to absorb pressure fluctuations, maintaining the downstream pressure at least 20% higher than the propellant's saturated vapor pressure. A porous metal foam buffer layer is located at the injection orifice inlet, forming a propellant injection unit. This unit stabilizes the liquid phase continuity through capillary action, reducing the amplitude of flow rate changes caused by bubble collapse. This invention systematically solves problems such as vaporization and phase change of cryogenic propellants, improves injector atomization and combustion stability, and maintains normal engine thrust output. Attached Figure Description

[0017] Other features, objects, and advantages of the present invention will become more apparent from the following detailed description of non-limiting embodiments with reference to the accompanying drawings: Figure 1 This is a schematic cross-sectional view of the structure of the present invention; Figure 2 This is a schematic diagram of the composite coating structure; Figure 3 This is a schematic diagram of the accumulator chamber. Figure 4 This is a schematic diagram of the buffer layer structure.

[0018] The diagram shows: DC channel 1; Composite coating 2; accumulator chamber 3; Head cavity 4; Buffer layer 5. Detailed Implementation

[0019] The present invention will now be described in detail with reference to specific embodiments. These embodiments will help those skilled in the art to further understand the present invention, but do not limit the invention in any way. It should be noted that those skilled in the art can make several changes and improvements without departing from the concept of the present invention. These all fall within the protection scope of the present invention.

[0020] To address the technical problems caused by the vaporization of low-boiling-point cryogenic propellants in narrow flow channels and sharp-angled regions of rocket engine injectors, resulting in unstable flow rates, abnormal thrust output, and even structural ablation, this invention provides a rocket engine injector pressure compensation device, such as... Figure 1 As shown, it includes a constant current channel 1, a pressure accumulator 3, and a buffer layer 5. The inner surface of the constant current channel 1 is coated with a composite coating 2, which has superhydrophobic and / or hydrophilic functions. One end of the constant current channel 1 is connected to the upstream inlet of the propellant, and the other end of the constant current channel 1 is connected to the head cavity 4. The pressure accumulator 3 is located inside the head cavity 4 and is arranged close to the inner wall of the head cavity 4. The head cavity 4 is connected to the inlet of the injection hole. The buffer layer 5 is disposed in the inlet of the injection hole to form a propellant injection unit. The pressure accumulator 3 is formed by an elastic metal bellows, and the buffer layer 5 is made of porous metal foam.

[0021] The present invention also provides a method for pressure compensation of rocket engine injectors, wherein the two ends of the constant current channel 1 coated with composite coating 2 are respectively connected to the upstream inlet of propellant and the head cavity 4, so that cavitation bubbles can be quickly detached from the wall and carried away by the mainstream. The pressure accumulator 3 is arranged inside the head cavity 4 and closely attached to the inner wall of the head cavity 4, and the pressure fluctuation is absorbed by the expansion of the pressure accumulator 3. At the same time, the buffer layer 5 is arranged at the injection hole inlet.

[0022] like Figure 2 As shown, composite coating 2 is a fluorinated silane-graphene composite coating with a contact angle >160°, which effectively reduces bubble wettability, allowing cavitation bubbles to quickly detach from the wall and be carried away by the mainstream. Composite coating 2 must withstand the strong oxidation of MON-25 and be evaluated through 500 hours of ASTM B117 salt spray test.

[0023] Specifically, the accumulator chamber 3 consists of multiple turns of bellows. The diameter of the metal bellows is slightly smaller than the inner diameter of the propellant head cavity 4. The bellows are stacked, with their diameters corresponding to their stacking positions, taking into account the increase in size after expansion and deformation. The number of bellows ranges from 5 to 15 turns, depending on the pressure compensation requirements and the change in deformation volume. The greater the pressure compensation requirement, the more bellows are needed; the greater the change in the elastic metal deformation volume, the fewer bellows are needed.

[0024] It should be noted that the flexible metal bellows are stacked inside the head cavity 4, mainly located upstream of the inner wall, filling the traditional flow backflow area, forming an arc dome, reducing local eddies and sudden pressure changes, and avoiding flash evaporation caused by excessive pressure drop at a single point.

[0025] like Figure 3 As shown, the material of the elastic metal bellows is preferably nickel-titanium alloy with a bulk modulus of less than 5 MPa. When the vapor pressure of MON-25 rises sharply, the bellows expands to absorb the pressure fluctuation, maintaining the downstream pressure at more than 20% higher than the saturated vapor pressure of the propellant for incompressible liquid working fluid.

[0026] like Figure 4 As shown, the material of the buffer layer 5 is preferably aluminum foam with a porosity of 80% and a pore size of 50-100μm. It stabilizes the continuity of the liquid phase through capillary force and reduces the amplitude of flow rate changes caused by bubble rupture.

[0027] The injector pressure compensation device in this invention uses an elastic metal bellows accumulator chamber to suppress the vaporization phase change of the propellant, and uses a superhydrophobic / hydrophilic composite coating and a porous metal foam buffer layer to manage cavitation bubbles. It does not require the introduction of additional components and controls, and is easy to implement in engineering through separate processing and welding or additive manufacturing.

[0028] This embodiment also provides a rocket engine propulsion system. The rocket engine propulsion system provided in this embodiment includes the aforementioned injector pressure compensation device. The rocket engine propulsion system can be adapted to low-freezing-point propellant rocket engines. The technical advantages and effects achieved by the rocket engine propulsion system also include the technical advantages and effects achieved by the injector pressure compensation device, which will not be elaborated here.

[0029] This embodiment also provides a rocket, which includes the aforementioned rocket engine propulsion system. The technical advantages and effects achieved by the rocket also include the technical advantages and effects achieved by the rocket engine propulsion system, which will not be repeated here.

[0030] The working principle of this invention is as follows: In this invention, the constant current channel 1, coated with a composite coating 2, is connected at both ends to the upstream inlet of the propellant and the head cavity 4, respectively. This allows cavitation bubbles to quickly detach from the wall and be carried away by the mainstream. The metal bellows accumulator 3 is located inside the propellant head cavity 4, arranged close to the inner wall. The bellows expands to absorb pressure fluctuations, maintaining the downstream pressure at least 20% higher than the saturated vapor pressure of the propellant. The porous metal foam buffer layer 5 is located at the injection port inlet, forming a propellant injection unit. It stabilizes the liquid phase continuity through capillary force, reduces the amplitude of flow oscillations caused by bubble collapse, systematically solves problems such as the vaporization phase change of cryogenic propellants, improves the atomization and combustion stability of the injector, and maintains normal engine thrust output.

[0031] In the description of this application, it should be understood that the terms "upper", "lower", "front", "back", "left", "right", "vertical", "horizontal", "top", "bottom", "inner", "outer", etc., indicate the orientation or positional relationship based on the orientation or positional relationship shown in the accompanying drawings. They are only for the convenience of describing this application and simplifying the description, and do not indicate or imply that the device or element referred to must have a specific orientation, or be constructed and operated in a specific orientation. Therefore, they should not be construed as limitations on this application.

[0032] Specific embodiments of the present invention have been described above. It should be understood that the present invention is not limited to the specific embodiments described above, and those skilled in the art can make various changes or modifications within the scope of the claims, which do not affect the essence of the present invention. Unless otherwise specified, the embodiments and features described in this application can be arbitrarily combined with each other.

Claims

1. A rocket engine injector pressure compensation device, characterized in that, It includes a DC channel (1), a accumulator (3), and a buffer layer (5); One end of the constant DC channel (1) is connected to the upstream inlet of the propellant, and the other end of the constant DC channel (1) is connected to the head cavity (4). The accumulator (3) is located inside the head cavity (4) and is arranged close to the inner wall of the head cavity (4). The buffer layer (5) is arranged at the inlet of the injection hole to form a propellant injection unit. The inner surface of the constant DC channel (1) is coated with a composite coating (2).

2. A method for compensating the pressure of a rocket engine injector, characterized in that, The two ends of the constant current channel (1) coated with composite coating (2) are connected to the upstream inlet of the propellant and the head cavity (4) respectively, so that the cavitation bubbles can quickly detach from the wall and be carried away by the mainstream. The pressure accumulator (3) is arranged inside the head cavity (4) and closely attached to the inner wall of the head cavity (4). The pressure accumulator (3) expands to absorb pressure fluctuations. At the same time, the buffer layer (5) is arranged at the injection hole inlet.

3. The rocket engine injector pressure compensation device according to claim 1 or the rocket engine injector pressure compensation method according to claim 2, characterized in that, The composite coating (2) is made of fluorinated silane-graphene composite coating with a contact angle >160°.

4. The rocket engine injector pressure compensation device according to claim 1 or the rocket engine injector pressure compensation method according to claim 2, characterized in that, The pressure storage chamber (3) is a cavity formed by an elastic metal bellows; The diameter of the metal bellows is smaller than the inner diameter of the propellant head cavity (4), and the metal bellows are stacked.

5. The rocket engine injector pressure compensation device or the rocket engine injector pressure compensation method according to claim 4, characterized in that, The accumulator chamber (3) includes multiple turns of bellows, with the number of turns being 5 to 15.

6. The rocket engine injector pressure compensation device or the rocket engine injector pressure compensation method according to claim 4, characterized in that, The metal bellows is made of nickel-titanium alloy with a bulk modulus of less than 5 MPa. When the vapor pressure of MON-25 rises sharply, the bellows expands to absorb the pressure fluctuation, maintaining the downstream pressure of the incompressible liquid working fluid at more than 20% higher than the saturated vapor pressure of the propellant.

7. The rocket engine injector pressure compensation device according to claim 1 or the rocket engine injector pressure compensation method according to claim 2, characterized in that, The buffer layer (5) is made of porous metal foam.

8. The rocket engine injector pressure compensation device according to claim 1 or the rocket engine injector pressure compensation method according to claim 2, characterized in that, The buffer layer (5) is made of aluminum foam with a porosity of 80% and a pore size of 50-100μm. It stabilizes the continuity of the liquid phase through capillary force and reduces the amplitude of flow rate changes caused by bubble rupture.

9. A rocket engine propulsion system, characterized in that, The rocket engine injector pressure compensation device according to any one of claims 1 to 8 is adopted.

10. A rocket, characterized in that, Includes the rocket engine propulsion system as described in claim 9.

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