Rocket engine injector rectifying device and method, propulsion system and rocket

Abstract

The invention relates to the technical field of spacecraft propulsion systems, and provides a rocket engine injector rectifying device and method, a propulsion system and a rocket. The rocket engine injector rectifying device comprises a tangential inlet, a divergent flow channel, an arc-shaped dome liquid collecting cavity and an injection hole; the two ends of the divergent flow channel are connected with the tangential inlet and one end of the arc-shaped dome liquid collecting cavity respectively; a micro-vortex generator array is arranged on the inner wall face of the divergent flow channel, and the injection hole is connected with the other end of the arc-shaped dome liquid collecting cavity to form a propellant injection unit. Bubbles are pushed to the center of a flow channel through a centrifugal force field, and bubble retention at the near wall is avoided; the inner wall face of the divergent flow channel induces turbulence pulsation through a fin array, enhances boundary layer energy exchange, inhibits local pressure sudden drop, reduces vaporization phase change of a downstream propellant, ensures normal propellant conveying flow, improves the stability of injection flow, atomization and combustion processes, and maintains normal and reliable output thrust of an engine.

Description

Technical Field

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

[0006] A rocket engine injector rectifying device according to the present invention includes a tangential inlet, a gradually expanding flow channel, an arc-shaped dome liquid collection chamber, and an injection hole; One end of the gradually expanding channel is connected to the tangential inlet, and the other end of the gradually expanding channel is connected to one end of the arc-shaped dome liquid collection cavity, forming a propellant head cavity and channel; A micro vortex generator array is disposed on the inner wall surface of the gradually expanding channel. The micro vortex generator array is located downstream of the tangential inlet. The injection hole is connected to the other end of the arc-shaped dome liquid collection cavity to form a propellant injection unit.

[0007] According to a rocket engine injector rectification method provided by the present invention, the propellant head cavity and flow channel are configured as follows: one end of the gradually expanding flow channel is connected to a tangential inlet, the other end of the gradually expanding flow channel is connected to one end of an arc-shaped dome liquid collection cavity, and a micro vortex generator array is arranged on the inner wall surface of the gradually expanding flow channel such that the micro vortex generator array is located downstream of the tangential inlet, and the injection hole is connected to the other end of the arc-shaped dome liquid collection cavity to form a propellant injection unit.

[0008] Preferably, the tangential inlet converts the propellant into a swirling flow within the channel, with a swirling angle of β, and the value of β is preferably 30° to 60°.

[0009] Preferably, the expansion angle of the gradually expanding channel is α, and the value of α is preferably 15° to 30°.

[0010] Preferably, the micro vortex generator array comprises a single row or multiple rows of fins.

[0011] Preferably, the height of a single fin of the micro vortex generator array is h, and the length is l, with the length l being 5 to 10 times the height h.

[0012] Preferably, the value of h ranges from 0.5 mm to 1.0 mm.

[0013] Preferably, the fin angle of the micro vortex generator array is θ, which differs from the swirl angle by less than 5°; The spacing between adjacent vortex fins in the same row is b, where b is 3 to 5 times the boundary layer thickness. The vortex fins in different rows are staggered, and the spacing between rows is a, which is consistent with the range of b.

[0014] According to the present invention, a rocket engine propulsion system employs the aforementioned rocket engine injector rectification 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, the gradually expanding channel is connected to the tangential inlet and the arc-shaped dome liquid collection cavity to form the propellant head cavity and channel. The centrifugal force field pushes the bubbles towards the center of the channel, avoiding bubble stagnation near the wall. The micro vortex generator array is located downstream of the tangential inlet on the inner wall of the gradually expanding channel. It induces turbulent pulsation through the fin array, enhances boundary layer energy exchange, suppresses sudden local pressure drops, reduces downstream propellant vaporization phase change, ensures normal propellant delivery flow rate, improves the stability of injection flow, atomization and combustion processes, and maintains normal and reliable 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 rocket engine injector rectifier device in this invention; Figure 2 A schematic cross-sectional view of the tangential inlet and the gradually expanding channel; Figure 3 This is a schematic diagram of bubble movement in a gradually expanding flow channel; Figure 4 This is a schematic diagram of bubble movement in a constant current channel; Figure 5 This is a schematic diagram of a micro vortex generator array. Figure 6 A schematic diagram of the flow direction and vortex generation and development in a micro vortex generator; Figure 7 This is a schematic diagram of the parameters of the micro vortex generator array.

[0018] The diagram shows: Tangential inlet 1; Micro-vortex generator array 2; Gradually expanding flow channel 3; Arc-shaped dome liquid collection chamber 4; Injection hole 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] This invention provides a rocket engine injector straightening device, such as... Figure 1 As shown, it includes a tangential inlet 1, a gradually expanding channel 3, an arc-shaped dome liquid collection chamber 4, and an injection hole 5. One end of the gradually expanding channel 3 is connected to the tangential inlet 1, and the other end of the gradually expanding channel 3 is connected to one end of the arc-shaped dome liquid collection chamber 4, forming a propellant head cavity and channel. A micro vortex generator array 2 is arranged on the inner wall surface of the gradually expanding channel 3. The micro vortex generator array 2 is located downstream of the tangential inlet 1. The injection hole 5 is connected to the other end of the arc-shaped dome liquid collection chamber 4, forming a propellant injection unit.

[0021] The present invention also provides a method for rectifying the flow of a rocket engine injector, wherein the propellant head cavity and flow channel are configured such that one end of the gradually expanding flow channel 3 is connected to the tangential inlet 1, the other end of the gradually expanding flow channel 3 is connected to one end of the arc-shaped dome liquid collection cavity 4, and a micro vortex generator array 2 is arranged on the inner wall surface of the gradually expanding flow channel 3 such that the micro vortex generator array 2 is located downstream of the tangential inlet 1, and the injection hole 5 is connected to the other end of the arc-shaped dome liquid collection cavity 4 to form a propellant injection unit.

[0022] like Figure 2 As shown, the tangential inlet 1 transforms the propellant into a swirling flow within the channel, with a swirling angle of β. The preferred value of β is 30°–60°, creating a stable stratification between a low-pressure zone at the center and a high-pressure zone at the wall. Centrifugal force pushes the bubbles towards the center of the channel, preventing bubble retention and aggregation near the wall. The expansion angle of the gradually diffusing channel 3 is α, preferably 15°–30°, reducing the flow velocity and minimizing flash evaporation caused by adiabatic expansion. The expansion angle of the gradually diffusing channel 3... α To adapt to the propellant swirling process, an excessively small expansion angle affects the centrifugal force field's ability to drive bubbles, while an excessively large expansion angle reduces propellant transport efficiency. Optimal expansion angle parameters can be obtained through cold flow experiments or numerical simulations. Furthermore, the swirling pattern can enhance convective heat transfer and create a uniform temperature field.

[0023] Furthermore, the tangential inlet 1 needs to be adapted to the propellant supply flow rate on the one hand, and on the other hand, it needs to ensure that the internal swirl angle is 30° to 60°. If the swirl angle is too small, it may cause central cavitation, and if the swirl angle is too large, it will affect the effect of stripping bubbles near the wall. Through cold flow tests or numerical simulations, the optimal parameters for the area and number of tangential inlets can be obtained to ensure that the static pressure at any point in the entire flow channel is higher than the saturated vapor pressure of the propellant at that point.

[0024] like Figure 3 and Figure 4 As shown, when the propellant flows in the channel, after vaporization and phase change, gas nuclei are formed at the wall of the gradually expanding channel 3. (See figure). Figure 4 In the process described above, gas nuclei (B) act as the initial carriers for bubble formation, gradually transforming into small bubbles (A). If these small bubbles are not promptly detached from the wall, they will gradually grow larger with gas accumulation, potentially causing instability in downstream propellant injection, leading to rough combustion and pressure oscillations. Furthermore, when small bubbles (A) are located at the wall, gas nuclei (B) induce more propellant vaporization phase changes. This invention employs a diffuser channel (3) and converts the propellant within the diffuser channel (3) into a swirling flow, which can reduce the bubble volume fraction by 60% at -40°C, effectively improving the stability of propellant injection.

[0025] like Figure 5 As shown, the micro-vortex generator array 2 is arranged on the inner wall of the gradually expanding channel 3, downstream of the tangential inlet 1. The micro-vortex generator array 2 includes a single row or multiple rows of rectangular fins, as shown in the figure. Figure 5 In the C section, the micro-vortex generator array 2 can induce turbulent pulsations, enhance boundary layer energy exchange, suppress local pressure drops, and reduce downstream propellant vaporization phase changes.

[0026] It should be noted that the fin height, length, spacing and number of rows of the micro vortex generator array 2 all affect the induction and enhancement of downstream turbulent pulsation. The intensity and number of flow vortices determine the energy exchange. The optimal fin parameters of the array can be obtained through jet cold atomization test and engine hot test chamber pressure curve measurement.

[0027] like Figure 6 As shown, the height of a single fin in the micro-vortex generator array 2 is h, ranging from 0.5 mm to 1.0 mm, and its length is l, which is 5 to 10 times the height h. When installed at a specific angle, it generates a flow vortex with a small aspect ratio and high intensity. The vortex injects high-speed liquid flow energy into the downstream boundary layer through mixing, delaying the occurrence of separation and suppressing the propellant vaporization phase change process. When the propellant flows on a smooth wall, the boundary layer near the wall is a viscous sublayer, such as... Figure 6 In the diagram, velocity D exhibits a gradient distribution, with the central region showing approximately laminar flow. Figure 6 In the case of E, when there are structures such as sudden contraction and sudden expansion downstream, the local static pressure decreases, and the propellant is very prone to vaporization and phase change.

[0028] like Figure 7 As shown, the fin angle of the micro-vortex generator array 2 is θ, which differs from the vortex angle by less than 5°, i.e., -5° < θ - β < 5°. The spacing between adjacent vortex fins in the same row is b, where b is 3 to 5 times the boundary layer thickness. The vortex fins in different rows are staggered, and the spacing between rows is a, which is consistent with the range of b. The arc-shaped dome liquid collection cavity 4 and the gradually expanding flow channel 3 transition smoothly. The propellant swirls along the inner wall surface into the liquid collection cavity, avoiding right-angle bends or sudden contraction / expansion structures, reducing flow separation and backflow zones, and avoiding propellant vaporization phase change caused by sudden drops in local pressure.

[0029] The propellant head cavity and flow channel formed in this invention push bubbles towards the center of the flow channel through a centrifugal force field, avoiding bubble stagnation near the wall. Furthermore, the fin array induces turbulent pulsation, enhancing boundary layer energy exchange, suppressing sudden local pressure drops, and reducing downstream propellant vaporization phase changes. This systematically solves problems such as low-temperature propellant vaporization phase changes, ensuring normal propellant delivery flow rate, improving the stability of injection flow, atomization, and combustion processes, and maintaining normal and reliable engine thrust output. Simultaneously, the injector rectification device in this invention utilizes the flow channel structure design and micro-vortex generator array to suppress propellant vaporization phase changes without the need for additional devices and controls. It is easily implemented in engineering through separate machining and welding or additive manufacturing.

[0030] The present invention also provides a rocket engine propulsion system, which includes a rocket engine injector rectifier. The technical advantages and effects achieved by the rocket engine propulsion system also include the technical advantages and effects achieved by the injector rectifier, which will not be elaborated here.

[0031] It should be noted that the rocket engine propulsion system can be adapted to rocket engines using low-freezing-point propellants.

[0032] The present invention also provides a rocket, including a 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.

[0033] The working principle of this invention is as follows: The propellant enters sequentially through the tangential inlet 1 into the diffuser channel 3, the arc-shaped dome collecting chamber 4, and the injection orifice 5. As the propellant flows through the channel, it undergoes vaporization and phase change, forming gas nuclei on the wall of the diffuser channel 3. These gas nuclei act as the initial carriers for bubble formation, gradually transforming into small bubbles. If these small bubbles are not promptly detached from the wall, they will gradually grow larger with gas accumulation, potentially causing instability in downstream propellant injection, leading to rough combustion and pressure oscillations. Furthermore, when small bubbles are located at the wall, the gas nuclei induce more propellant vaporization and phase change. The tangential inlet 1 transforms the propellant into a swirling flow within the channel, creating a stable stratification between a low-pressure zone at the center and a high-pressure zone at the wall. Centrifugal force pushes the bubbles towards the center of the channel, preventing bubble retention and aggregation near the wall. The diffuser channel 3 has an expansion angle, which reduces flow velocity and minimizes flash evaporation caused by adiabatic expansion.

[0034] The micro vortex generator array 2 is arranged on the inner wall of the gradually expanding channel 3, located downstream of the tangential inlet 1. The micro vortex generator array 2 consists of a single row or multiple rows of rectangular fins, which can induce turbulent pulsation, enhance boundary layer energy exchange, suppress local pressure drop, and reduce downstream propellant vaporization phase change.

[0035] 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.

[0036] 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 rectifying device, characterized in that, It includes a tangential inlet (1), a gradually expanding flow channel (3), an arc-shaped dome liquid collection chamber (4), and a spray hole (5); One end of the gradually expanding channel (3) is connected to the tangential inlet (1), and the other end of the gradually expanding channel (3) is connected to one end of the arc-shaped dome liquid collection cavity (4) to form a propellant head cavity and channel; The inner wall of the gradually expanding channel (3) is provided with a micro vortex generator array (2), which is located downstream of the tangential inlet (1). The injection hole (5) is connected to the other end of the arc-shaped dome liquid collection chamber (4) to form a propellant injection unit.

2. A method for rectifying the flow of a rocket engine injector, characterized in that, The propellant head cavity and flow channel are configured such that one end of the gradually expanding flow channel (3) is connected to the tangential inlet (1), the other end of the gradually expanding flow channel (3) is connected to one end of the arc-shaped dome liquid collection cavity (4), and a micro vortex generator array (2) is arranged on the inner wall surface of the gradually expanding flow channel (3) and the micro vortex generator array (2) is located downstream of the tangential inlet (1), and the injection hole (5) is connected to the other end of the arc-shaped dome liquid collection cavity (4) to form a propellant injection unit.

3. The rocket engine injector rectification device according to claim 1 or the rocket engine injector rectification method according to claim 2, characterized in that, The tangential inlet (1) converts the propellant into a swirling flow within the channel, with a swirling angle of β, preferably 30° to 60°.

4. The rocket engine injector rectification device according to claim 1 or the rocket engine injector rectification method according to claim 2, characterized in that, The expansion angle of the gradually expanding channel (3) is α, and the value of α is preferably 15° to 30°.

5. The rocket engine injector rectification device according to claim 1 or the rocket engine injector rectification method according to claim 2, characterized in that, The micro vortex generator array (2) includes a single row or multiple rows of fins.

6. The rocket engine injector rectification device or the rocket engine injector rectification method according to claim 5, characterized in that, The height of a single fin of the micro vortex generator array (2) is h, and the length is l, with the length l being 5 to 10 times the height h.

7. The rocket engine injector rectification device or the rocket engine injector rectification method according to claim 6, characterized in that, The value of h ranges from 0.5 mm to 1.0 mm.

8. The rocket engine injector rectification device or the rocket engine injector rectification method according to claim 5, characterized in that, The fin angle of the micro vortex generator array (2) is θ, which is within 5° different from the vortex angle; The spacing between adjacent vortex fins in the same row is b, where b is 3 to 5 times the boundary layer thickness. The vortex fins in different rows are staggered, and the spacing between rows is a, which is consistent with the range of b.

9. A rocket engine propulsion system, characterized in that, The rocket engine injector rectifier 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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