Integrated direct-combustion liquid rocket engine

Abstract

The present invention discloses an integrated direct-combustion liquid rocket engine, comprising: a combustion chamber, a nozzle mounted below the combustion chamber, and a rotor mounted within the combustion chamber. The rotor of the present invention is a stainless steel product integrated through 3D printing, integrating multiple key components such as an oxygen impeller, oxygen guide vanes, a fuel impeller, fuel guide vanes, an auxiliary shaft, a turbine, and a main shaft. This not only makes the overall structure of the engine compact, reduces its volume, and achieves lightweight, but also significantly improves the installation efficiency of the present invention and greatly reduces labor costs. The turbine of the present invention has an internal cavity structure and is connected to the fuel flow channel, effectively reducing the temperature of the turbine in high-temperature environments and preventing performance degradation or structural damage due to overheating.

Description

Technical Field

[0001] The invention relates to an integrated direct-combustion driven liquid rocket engine. Background Art

[0002] The main components of existing liquid rocket engines include a gas generator and accessories, a combustion chamber, and a turbopump. The gas generator, equivalent to a small liquid rocket engine, requires specialized valves, piping, injectors, a combustion chamber, and an igniter, resulting in a complex structure and poor reliability.

[0003] Patent application number 202211018754.0 eliminates the gas generator and makes further improvements, extracting the gas directly from the combustion chamber to drive the turbine. Currently, this technical solution is used in open rockets. However, in open rockets, the rich fuel gas that drives the turbine is directly discharged to the outside, resulting in energy waste and affecting the rocket's specific impulse. Furthermore, the discharge of the rich fuel gas to the outside world is prone to the risk of accidental ignition.

[0004] Furthermore, even though the aforementioned patent eliminates the independent gas generator, the presence of the independent turbopump still makes it difficult to completely simplify the rocket engine structure. This is because the turbopump itself is a complex system, with its own independent workflow and supporting facilities, and requires various pipe connections to the engine's combustion chamber and propellant tank. Therefore, this prior art fails to fundamentally address the issues of structural complexity and high cost. Summary of the Invention

[0005] In order to overcome the above-mentioned shortcomings, the present invention aims to provide a technical solution that can solve the above-mentioned problems.

[0006] An integrated direct-combustion liquid rocket engine comprises: a combustion chamber, a nozzle mounted below the combustion chamber, and a rotor mounted inside the combustion chamber; the rotor is a stainless steel product integrated by 3D printing;

[0007] The rotor is composed of an oxygen impeller, an oxygen guide vane, a fuel impeller, a fuel guide vane, an auxiliary shaft, a turbine and a main shaft; wherein the oxygen impeller, the fuel impeller, the auxiliary shaft and the turbine are sequentially arranged on the same shaft from the inside to the outside;

[0008] An oxidant flow channel is formed between the oxygen impeller and the fuel impeller, and a fuel flow channel is formed between the fuel impeller and the lower end of the auxiliary shaft;

[0009] The outer side of the turbine is rotatably connected to the bottom side wall of the combustion chamber via a first rotating bearing ring;

[0010] The turbine has an internal cavity structure, the interior of which is connected to the fuel flow channel, and the turbine is evenly provided with a plurality of oblique spray slots penetrating the upper and lower end surfaces of the turbine.

[0011] Preferably, the oblique nozzle slots are connected to the combustion chamber and the interior of the nozzle, respectively; the inner diameter of the top section of the nozzle is smaller than the outer diameter of the turbine; and the top of the nozzle is rotatably connected to the turbine via a fourth bearing ring;

[0012] Preferably, the upper end of the fuel impeller is connected to a main rotating shaft, and the main rotating shaft is an internal hollow structure and its interior is connected to the oxidant flow channel;

[0013] Preferably, the upper end of the main rotating shaft is rotatably connected to the top of the combustion chamber via a second rotating bearing ring, and the top end of the main rotating shaft passes through the top of the combustion chamber and is rotatably connected to the oxygen pipeline;

[0014] Preferably, the top end of the auxiliary shaft is rotatably connected to the inner side of the top of the combustion chamber via a third bearing ring, and the gap between the auxiliary shaft and the main shaft is a fuel inlet and is connected to the fuel flow channel;

[0015] Preferably, the oxygen guide vanes are respectively connected to the outer periphery of the oxygen impeller and the inner periphery of the fuel impeller, and the fuel guide vanes are respectively connected to the outer periphery of the fuel impeller and the inner periphery of the lower end of the auxiliary shaft;

[0016] Preferably, the connection between the upper end of the main rotating shaft and the top of the combustion chamber is sealed by a sealing ring, and the connection between the auxiliary rotating shaft and the inner side of the top of the combustion chamber is sealed by a sealing ring;

[0017] Preferably, the upper end surface of the turbine is connected to the lower end of the auxiliary shaft, and the lower end surface of the turbine is circumferentially connected to the fuel impeller;

[0018] Preferably, a fuel injection nozzle is provided on the auxiliary shaft at the rear end of the fuel guide vane, and the fuel injection nozzle is connected to the fuel flow channel and the combustion chamber;

[0019] Preferably, an oxidant injection nozzle communicating with the combustion chamber is provided on the fuel impeller at the rear end of the oxygen guide vane.

[0020] Compared with the prior art, the advantages of the present invention are:

[0021] The rotor of the present invention is a stainless steel product integrated through 3D printing. It integrates multiple key components such as oxygen impeller, oxygen guide vanes, fuel impeller, fuel guide vanes, auxiliary shaft, turbine and main shaft. It not only makes the overall structure of the engine compact, the volume reduced and the weight reduced, but also significantly improves the installation efficiency of the present invention, greatly reducing the pre-launch preparation time and labor cost investment.

[0022] The turbine of the present invention has an internal cavity structure and is connected to the fuel flow channel, which effectively reduces the temperature of the turbine in a high-temperature environment and prevents performance degradation or structural damage due to overheating. In addition, the fuel absorbs part of the heat during the heat exchange process, which increases the internal energy of the fuel and helps to release energy more fully in the subsequent combustion process.

[0023] Additional aspects and advantages of the present invention will be set forth in part in the description which follows and, in part, will be obvious from the description which follows, or may be learned by practice of the present invention. BRIEF DESCRIPTION OF THE DRAWINGS

[0024] In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the following briefly introduces the drawings required for use in the embodiments or the description of the prior art. Obviously, the drawings described below are only some embodiments of the present invention. For ordinary technicians in this field, other drawings can be obtained based on these drawings without paying any creative work.

[0025] Figure 1 It is a structural schematic diagram of the present invention.

[0026] Figure 2 yes Figure 1 Enlarged view of circle A in the middle.

[0027] Figure 3 It is a perspective view of the rotor of the present invention. DETAILED DESCRIPTION

[0028] The following is a clear and complete description of the technical solutions in the embodiments of the present invention. Obviously, the embodiments described are only some embodiments of the present invention, not all embodiments. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without making any creative efforts are within the scope of protection of the present invention.

[0029] In the description of the present invention, it should be noted that the terms "center", "up", "down", "left", "right", "vertical", "horizontal", "inside", "outside", etc., indicating directions or positional relationships, are based on the directions or positional relationships shown in the accompanying drawings. They are only for the convenience of describing the present invention and simplifying the description, and do not indicate or imply that the device or element referred to must have a specific direction, be constructed and operated in a specific direction. Therefore, they cannot be understood as limiting the present invention.

[0030] Furthermore, in the description of the present invention, 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; internal communication between two components; and wireless or wired connections. Those skilled in the art will understand the specific meanings of the above terms in the present invention based on the specific circumstances.

[0031] In addition, the technical features involved in the different embodiments of the present invention described later can be combined with each other as long as they do not conflict with each other.

[0032] See also Figures 1 to 3 In an embodiment of the present invention, an integrated direct-combustion driven liquid rocket engine includes: a combustion chamber 1, a nozzle 2 installed below the combustion chamber 1, and a rotor 3 installed inside the combustion chamber 1; the present invention uses a bipropellant, and the fuel can be liquid hydrogen or methane, etc. In order to reduce costs, the present invention can also use kerosene or alcohol as fuel, and the oxidizer selects commonly used liquid oxygen.

[0033] In one embodiment, the rotor 3 is a 3D-printed stainless steel product with an integrated structure. It integrates multiple key components such as the oxygen impeller 31, oxygen guide vanes 32, fuel impeller 33, fuel guide vanes 34, auxiliary shaft 35, turbine 36, and main shaft 37. The oxygen impeller 31, fuel impeller 33, auxiliary shaft 35, and turbine 36 are sequentially arranged from the inside to the outside of the same shaft. This compact layout makes the space between the components more reasonable, reduces the radial size of the engine, and improves the integration. Therefore, the present invention greatly reduces the complex connection structure between the components, not only making the overall structure of the engine compact and reducing the volume, but also truly achieving lightweighting. Compared with traditional engines, under the same thrust output, the present invention significantly improves the rocket payload ratio.

[0034] In one embodiment, the annular outer side of the turbine 36 is rotatably connected to the bottom side wall of the combustion chamber 1 through a first bearing ring 5, the upper end of the main shaft 37 is rotatably connected to the top of the combustion chamber 1 through a second bearing ring 6, the top of the auxiliary shaft 35 is rotatably connected to the inner side of the top of the combustion chamber 1 through a third bearing ring 7, and the top of the nozzle 2 is rotatably connected to the turbine 36 through a fourth bearing ring 8. It can be seen that the present invention realizes the rotational connection between the rotor 3 and the combustion chamber 1 and the nozzle 2 through multiple bearing rings, which can provide stable support for the rotor 3, ensure the stability and reliability of the rotor 3 during high-speed rotation, reduce vibration and wear, and enable the rotor 3 to rotate smoothly around its axis, thereby ensuring the normal operation of the engine.

[0035] In one embodiment, the integral and integrated rotor 3 of the present invention is installed inside the combustion chamber 1, which greatly simplifies the installation process. It is only necessary to install the first bearing ring 5 and the second bearing ring 6 at the bottom and top of the combustion chamber 1 respectively, install the third bearing ring 7 on the auxiliary shaft 35, and install the fourth bearing ring 8 on the top of the nozzle 2. Then, the combustion chamber 1 is hoisted and inserted into the corresponding position of the rotor 3, so that the top of the main shaft 37 passes through the top of the combustion chamber 1. At this time, the outside of the turbine 36, the upper end of the main shaft 37, and the top of the auxiliary shaft 35 are rotatably connected to the corresponding position of the combustion chamber 1. Finally, the combustion chamber 1 and the nozzle 2 are fixedly connected with bolts, and the fourth bearing ring 8 is rotated to abut against the bottom of the turbine 36, and the installation of the present invention is completed.

[0036] Therefore, the present invention significantly improves installation efficiency and greatly reduces pre-launch preparation time and labor cost investment.

[0037] In one embodiment, the turbine 36 is an internal cavity structure, with its upper end surface connected to the lower end of the auxiliary shaft 35 and its lower end surface circumferentially connected to the fuel impeller 33 , so that the interior of the turbine 36 is connected to the fuel flow channel 200 .

[0038] When the present invention is in working condition, fuel will fill the inner cavity of the turbine 36; the fuel inside the turbine 36 can cool the entire turbine 36 and the side walls of the oblique spray slot 360, effectively reducing the temperature of the turbine 36 in a high temperature environment, preventing performance degradation or structural damage due to overheating, and ensuring the stable operation of the turbine 36; on the other hand, the fuel absorbs part of the heat during the heat exchange process, causing its own temperature to increase, thereby increasing the internal energy of the fuel to a certain extent, and helping the fuel to release energy more fully in the subsequent combustion process.

[0039] The specific working principle of the present invention is as follows: When the present invention is in working state, liquid fuel flows into the fuel inlet 100 between the auxiliary rotating shaft 35 and the main rotating shaft 37 through the fuel pipeline, and the fuel inlet 100 is connected to the fuel flow channel 200 formed between the fuel impeller 33 and the lower end of the auxiliary rotating shaft 35, and the fuel then moves along the fuel flow channel 200; at the same time, low-temperature liquid oxygen enters the oxidant flow channel 400 between the oxygen impeller 31 and the fuel impeller 33 from the oxidant inlet 300 at the upper end of the main rotating shaft 37 through the oxygen pipeline.

[0040] Inside rotor 3, an oxygen impeller 31, a fuel impeller 33, an auxiliary shaft 35, and a turbine 36 are arranged in sequence from the inside outward about the same shaft. Driven by turbine 36, oxygen impeller 31 and fuel impeller 33 rotate at high speeds. The liquid oxygen in oxidizer flow channel 400 and the fuel in fuel flow channel 200 are compressed by oxygen guide vanes 32 and fuel guide vanes 34, respectively, causing the liquid oxygen and fuel to be injected into combustion chamber 1 through oxidizer injection nozzle 330 and fuel injection nozzle 350, respectively.

[0041] Within combustion chamber 1, high-pressure oxygen and fuel are thoroughly mixed in a specific ratio. After being ignited by the ignition device, they instantly and intensely combust, generating high-temperature, high-pressure combustion gas. The full flow of high-temperature combustion gas flows toward the inclined nozzle slots 360 of turbine 36. As the gas flows through inclined nozzle slots 360, its flow direction changes due to the tilt angle of inclined nozzle slots 360, generating a momentum change related to the rotational direction of turbine 36. This provides rotational power for turbine 36, driving it to rotate at high speed. Turbine 36 drives oxygen impeller 31 and fuel impeller 33 to operate continuously and stably, maintaining a steady supply of fuel and oxidant, and forming a stable operating cycle.

[0042] When the mixed chemical combustion in combustion chamber 1 of the present invention generates high-temperature, high-pressure combustion gas through inclined nozzle slot 360, the gas velocity reaches the speed of sound. Upon entering nozzle 2, the gas reaches supersonic speeds and is ejected from nozzle 2. According to Newton's third law, the supersonic gas exerts a reaction force on the rocket, propelling it forward. Nozzle 2 accelerates and directs the gas, ejecting it at the appropriate angle and velocity to maximize thrust output, providing the rocket with the necessary power for liftoff and flight, enabling it to overcome gravity and air resistance and achieve its intended flight trajectory.

[0043] It will be apparent to those skilled in the art that the present invention is not limited to the details of the exemplary embodiments described above and that the invention can be embodied in other specific forms without departing from the spirit or essential characteristics of the invention. Therefore, the embodiments should be considered in all respects as illustrative and non-restrictive, and the scope of the invention is defined by the appended claims rather than the foregoing description, and it is intended that all variations that come within the meaning and range of equivalents of the claims be embraced therein.

Claims

1. An integrated direct combustion liquid rocket engine, characterized in that: include: A combustion chamber, a nozzle installed below the combustion chamber, and a rotor installed inside the combustion chamber; the rotor is a stainless steel product integrated by 3D printing; The rotor is composed of an oxygen impeller, an oxygen guide vane, a fuel impeller, a fuel guide vane, an auxiliary shaft, a turbine and a main shaft; wherein the oxygen impeller, the fuel impeller, the auxiliary shaft and the turbine are sequentially arranged on the same shaft from the inside to the outside; An oxidant flow channel is formed between the oxygen impeller and the fuel impeller, and a fuel flow channel is formed between the fuel impeller and the lower end of the auxiliary shaft. The top end of the auxiliary shaft is rotatably connected to the inner side of the top of the combustion chamber via a third bearing ring. The gap between the auxiliary shaft and the main shaft serves as a fuel inlet and is connected to the fuel flow channel. A fuel injection nozzle is provided on the auxiliary shaft at the rear end of the fuel guide vane, and the fuel injection nozzle connects the fuel flow channel and the combustion chamber. The outer side of the turbine is rotatably connected to the bottom side wall of the combustion chamber via a first rotating bearing ring; The turbine has an internal cavity structure, the interior of which is communicated with the fuel flow channel, and a plurality of oblique spray slots penetrating the upper and lower end surfaces of the turbine are evenly arranged along the circumference of the turbine.

2. The integrated direct combustion liquid rocket engine according to claim 1, characterized in that: The oblique nozzle slots are connected to the combustion chamber and the inside of the nozzle respectively. The inner diameter of the top section of the nozzle is smaller than the outer diameter of the turbine. The top of the nozzle is connected to the turbine in a circumferential rotation through a fourth bearing ring.

3. The integrated direct combustion liquid rocket engine according to claim 1, characterized in that: The upper end of the fuel impeller is connected to the main rotating shaft, and the main rotating shaft is an internal hollow structure and the interior thereof is communicated with the oxidant flow channel.

4. The integrated direct combustion liquid rocket engine according to claim 1, characterized in that: The upper end of the main rotating shaft is rotatably connected to the top of the combustion chamber through a second rotating bearing ring, and the top end of the main rotating shaft passes through the top of the combustion chamber and is rotatably connected to the oxygen pipeline.

5. The integrated direct combustion liquid rocket engine according to claim 1, characterized in that: The oxygen guide vanes are respectively connected to the outer periphery of the oxygen impeller and the inner periphery of the fuel impeller, and the fuel guide vanes are respectively connected to the outer periphery of the fuel impeller and the inner periphery of the lower end of the auxiliary rotating shaft.

6. The integrated direct combustion liquid rocket engine according to claim 1, characterized in that: The connection between the upper end of the main rotating shaft and the top of the combustion chamber is sealed by a sealing ring, and the connection between the auxiliary rotating shaft and the inner side of the top of the combustion chamber is sealed by a sealing ring.

7. The integrated direct combustion liquid rocket engine according to claim 1, characterized in that: The upper end surface of the turbine is connected to the lower end of the auxiliary shaft, and the lower end surface of the turbine is circumferentially connected to the fuel impeller.

8. The integrated direct combustion liquid rocket engine according to claim 1, characterized in that: An oxidant injection nozzle communicating with the combustion chamber is provided on the fuel impeller at the rear end of the oxygen guide vane.

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