Integrated direct-combustion-driven liquid rocket engine

Abstract

The invention discloses an integrated direct-combustion-driven liquid rocket engine. The integrated direct-combustion-driven liquid rocket engine comprises a combustion chamber, a spray pipe installed below the combustion chamber and a rotor installed in the combustion chamber. The rotor is a stainless steel product integrated through 3D printing, multiple key components such as the oxygen impeller, the oxygen guide vane, the fuel impeller, the fuel guide vane, the auxiliary rotating shaft, the turbine and the main rotating shaft are integrated, the engine is compact in overall structure, the size is reduced, light weight is achieved, the installation efficiency is remarkably improved, and the service life of the engine is prolonged. The labor cost investment is greatly reduced; the turbine is of an internal cavity structure and communicates with the fuel runner, the temperature of the turbine in a high-temperature environment is effectively reduced, and performance reduction or structural damage caused by overheating is prevented.

Description

Technical Field

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

[0002] The main components of existing liquid rocket engines are: a gas generator and its accessories, a combustion chamber, and a turbopump. Among them, the gas generator is equivalent to a small liquid rocket engine, and special valves, pipelines, injectors, combustion chambers, igniters, etc. need to be set, with a complex structure and poor reliability.

[0003] The invention patent with the application number 202211018754.0 abandons the gas generator and makes further improvements, directly extracting gas from the combustion chamber to drive the turbine. Currently, this technical solution is applied to open rockets. However, the rich-burning gas of open rockets is directly discharged to the outside after driving the turbine, which causes energy waste and affects the improvement of the rocket's specific impulse. At the same time, when the rich-burning gas is discharged to the outside, there is also an easy risk of accidental ignition.

[0004] Moreover, even if the above-mentioned patent abandons the independent gas generator, the existence of its independent turbopump still makes it difficult to completely simplify the structure of the rocket engine. Because the turbopump itself is a complex system with an independent working process and supporting facilities, and various pipelines are required to connect it with the combustion chamber and propellant tank of the engine. Therefore, this existing technology is difficult to fundamentally solve the problems of complex structure and high cost. Summary of the Invention

[0005] The present invention aims to provide a technical solution to solve the above problems in order to overcome the above deficiencies.

[0006] An integrated direct combustion-driven liquid rocket engine, comprising: 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, oxygen guide vanes, a fuel impeller, fuel guide vanes, a secondary rotating shaft, a turbine, and a main rotating shaft; among them, the oxygen impeller, the fuel impeller, the secondary rotating shaft, and the turbine are coaxially sleeved and arranged from inside to outside in sequence; An oxidizer 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 secondary rotating shaft; The outside of the turbine is rotationally connected to the bottom side wall of the combustion chamber through a first rotating bearing ring; The turbine is of an internal cavity structure, its interior is communicated with the fuel flow channel, and a plurality of inclined spray grooves penetrating the upper and lower end faces of the turbine are evenly opened along the circumference of the turbine; Preferably, the inclined spray nozzles are respectively communicated with the combustion chamber and the inside of the nozzle. 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 rotationally connected with the turbine along the circumference through a fourth bearing ring; Preferably, the upper end of the fuel impeller is connected to the main rotating shaft. The main rotating shaft is a hollow structure inside and is communicated with the oxidant flow channel inside; Preferably, the upper end of the main rotating shaft is rotationally connected with the top of the combustion chamber through a second rotating bearing ring. The top of the main rotating shaft passes through the top of the combustion chamber and is rotationally connected with the oxygen pipeline; Preferably, the top end of the auxiliary rotating shaft is rotationally connected with the inner side of the top of the combustion chamber through a third rotating bearing ring. The gap between the auxiliary rotating shaft and the main rotating shaft is the fuel inlet and is communicated with the fuel flow channel; Preferably, the oxygen guide vanes are respectively connected to the outer circumference of the oxygen impeller and the inner circumference of the fuel impeller. The fuel guide vanes are respectively connected to the outer circumference of the fuel impeller and the inner circumference of the lower end of the auxiliary rotating shaft; Preferably, the connection between the upper end of the main rotating shaft and the top of the combustion chamber is sealed through a sealing ring. The connection between the auxiliary rotating shaft and the inner side of the top of the combustion chamber is sealed through a sealing ring; Preferably, the upper end surface of the turbine is connected to the lower end of the auxiliary rotating shaft, and the lower end surface of the turbine is connected to the fuel impeller along the circumference; Preferably, a fuel injection nozzle is arranged on the auxiliary rotating shaft at the rear end of the fuel guide vane. The fuel injection nozzle communicates the fuel flow channel and the combustion chamber; Preferably, an oxidant injection nozzle communicating with the combustion chamber is arranged on the fuel impeller at the rear end of the oxygen guide vane.

[0007] Compared with the prior art, the advantages of the present invention are as follows: The rotor of the present invention is an integrated stainless steel product by 3D printing, which integrates multiple key components such as an oxygen impeller, oxygen guide vanes, a fuel impeller, fuel guide vanes, an auxiliary rotating shaft, a turbine, and a main rotating shaft. It not only makes the overall structure of the engine compact, reduces the volume, realizes lightweight, but also significantly improves the installation efficiency of the present invention, and greatly reduces the preparation time and labor cost input before launch.

[0008] The turbine of the present invention is an internal cavity structure and is communicated with the fuel flow channel, which effectively reduces the temperature of the turbine in a high-temperature environment, prevents performance degradation or structural damage caused by overheating, and during the heat exchange process of the fuel, part of the heat is absorbed, increasing the internal energy of the fuel, which helps to release energy more fully in the subsequent combustion process.

[0009] The additional aspects and advantages of the present invention will be partially given in the following description, partially become obvious from the following description, or be understood through the practice of the present invention. Description of the Drawings

[0010] To more clearly illustrate the technical solutions in the embodiments of the present invention or the prior art, the following will briefly introduce the drawings required in the description of the embodiments or the prior art. Obviously, the drawings in the following description are only some embodiments of the present invention. For those of ordinary skill in the art, without creative efforts, other drawings can also be obtained based on these drawings.

[0011] Figure 1 It is a schematic structural diagram of the present invention.

[0012] Figure 2 is Figure 1 an enlarged view of the A circle in

[0013] Figure 3 It is a three-dimensional view of the rotor of the present invention. Specific Embodiments

[0014] The following will clearly and completely describe the technical solutions in the embodiments of the present invention. Obviously, the described embodiments are only some embodiments of the present invention, rather than all embodiments. Based on the embodiments of the present invention, all other embodiments obtained by those of ordinary skill in the art without creative efforts belong to the scope of protection of the present invention.

[0015] In the description of the present invention, it should be noted that the orientation or positional relationship indicated by the terms "center", "upper", "lower", "left", "right", "vertical", "horizontal", "inner", "outer", etc. is based on the orientation or positional relationship shown in the drawings. It is only for the convenience of describing the present invention and simplifying the description, rather than indicating or implying that the device or element referred to must have a specific orientation, be constructed and operated in a specific orientation, and therefore cannot be understood as a limitation of the present invention.

[0016] In addition, in the description of the present invention, unless otherwise clearly specified and limited, the terms "installation", "connection", "connection", etc. should be understood in a broad sense. For example, it can be a fixed connection, a detachable connection, or an integral connection; it can be a mechanical connection, an electrical connection; it can be directly connected, or indirectly connected through an intermediate medium, and can also be the communication inside two elements. It can be a wireless connection or a wired connection. For those of ordinary skill in the art, the specific meanings of the above terms in the present invention can be understood according to specific circumstances.

[0017] In addition, the technical features involved in different embodiments of the present invention described subsequently can be combined with each other as long as they do not conflict with each other.

[0018] Please refer to Figures 1 to 3, in the embodiments 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, methane, etc. To reduce costs, the present invention can also use kerosene or alcohol as the fuel, and the oxidizer is selected as the commonly used liquid oxygen.

[0019] In one embodiment, the rotor 3 is an integral stainless steel structure by 3D printing, which integrates multiple key components such as an oxygen impeller 31, oxygen guide vanes 32, a fuel impeller 33, fuel guide vanes 34, a secondary rotating shaft 35, a turbine 36, and a main rotating shaft 37. Moreover, the oxygen impeller 31, fuel impeller 33, secondary rotating shaft 35, and turbine 36 are coaxially sleeved from inside to outside in sequence. This compact layout makes the space utilization between components more reasonable, reduces the radial size of the engine, and improves the integration degree. Therefore, the present invention greatly reduces the complex connection structure between components, not only making the overall structure of the engine compact and the volume reduced, but also truly achieving lightweight. Compared with traditional engines, under the same thrust output, the present invention significantly improves the effective payload ratio of the rocket.

[0020] In one embodiment, the outer ring 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 rotating shaft 37 is rotatably connected to the top of the combustion chamber 1 through a second rotating bearing ring 6, the top end of the secondary rotating shaft 35 is rotatably connected to the inner side of the top of the combustion chamber 1 through a third rotating bearing ring 7, and the top of the nozzle 2 is rotatably connected to the turbine 36 along the circumference through a fourth bearing ring 8. Thus, 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, enable the rotor 3 to rotate smoothly around its axis, and ensure the normal operation of the engine.

[0021] 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. Only the first bearing ring 5 and the second bearing ring 6 need to be installed at the bottom and top of the combustion chamber 1 respectively, the third bearing ring 7 is installed on the secondary rotating shaft 35, and the fourth bearing ring 8 is installed at the top of the nozzle 2. Then, the combustion chamber 1 is hoisted and sleeved into the corresponding position of the rotor 3, so that the top end of the main rotating shaft 37 passes through the top of the combustion chamber 1. At this time, the outer side of the turbine 36, the upper end of the main rotating shaft 37, and the top end of the secondary rotating shaft 35 are rotatably connected to the corresponding positions of the combustion chamber 1. Finally, the combustion chamber 1 and the nozzle 2 are fixedly connected by bolts, and the fourth bearing ring 8 is rotationally abutted below the turbine 36, and the installation of the present invention is completed.

[0022] Therefore, the present invention significantly improves the installation efficiency and greatly reduces the preparation time and human cost input before launch.

[0023] In one embodiment, the turbine 36 has an internal cavity structure. Its upper end surface is connected to the lower end of the secondary rotating shaft 35, and its lower end surface is circumferentially connected to the fuel impeller 33. Therefore, the interior of the turbine 36 communicates with the fuel flow passage 200.

[0024] When the present invention is in the working state, the fuel fills 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 inclined injection grooves 360, effectively reducing the temperature of the turbine 36 in a high-temperature environment, preventing performance degradation or structural damage caused by overheating, and ensuring the stable operation of the turbine 36. On the other hand, during the heat exchange process, the fuel absorbs some heat, increasing its own temperature, which to a certain extent increases the internal energy of the fuel and helps the fuel release energy more fully in the subsequent combustion process.

[0025] Specific working principle of the present invention: When the present invention is in the working state, the liquid fuel flows in through the fuel inlet 100 between the secondary rotating shaft 35 and the main rotating shaft 37 via the fuel pipeline. The fuel inlet 100 is connected to the fuel flow passage 200 formed between the fuel impeller 33 and the lower end of the secondary rotating shaft 35, and the fuel then travels along the fuel flow passage 200. At the same time, the low-temperature liquid oxygen enters the oxidizer flow passage 400 between the oxidizer impeller 31 and the fuel impeller 33 through the oxidizer inlet 300 at the upper end inside the main rotating shaft 37 via the oxygen pipeline.

[0026] Inside the rotor 3, the oxidizer impeller 31, the fuel impeller 33, the secondary rotating shaft 35, and the turbine 36 are coaxially sleeved in sequence from the inside to the outside. Driven by the turbine 36, the oxidizer impeller 31 and the fuel impeller 33 rotate at high speed. The liquid oxygen in the oxidizer flow passage 400 and the fuel in the fuel flow passage 200 are respectively compressed by the oxidizer guide vanes 32 and the fuel guide vanes 34, causing the liquid oxygen and the fuel to be injected into the combustion chamber 1 from the oxidizer injection nozzle 330 and the fuel injection nozzle 350 respectively.

[0027] Inside the combustion chamber 1, the high-pressure oxygen and fuel are fully mixed in a specific proportion and are instantaneously and violently combusted after being ignited by the ignition device, generating high-temperature and high-pressure combustion gas. The full-flow high-temperature combustion gas flows towards the inclined injection grooves 360 of the turbine 36. When the combustion gas flows through the inclined injection grooves 360, the flow direction of the combustion gas changes due to the inclination angle of the inclined injection grooves 360, thereby generating a momentum change related to the rotation direction of the turbine 36, providing rotational power for the turbine 36, driving the turbine 36 to rotate at high speed, and the turbine 36 drives the oxidizer impeller 31 and the fuel impeller 33 to operate continuously and stably, maintaining the stable supply of fuel and oxidizer, and forming a stable working cycle. When the mixed chemical combustion in the combustion chamber 1 of the present invention generates high-temperature and high-pressure gas and passes through the inclined spray groove 360, the gas velocity has reached the speed of sound. After the gas enters the nozzle 2, it can reach supersonic speed, and finally the gas is ejected from the nozzle 2. According to Newton's third law, the supersonic gas generates a reaction force on the rocket, pushing the rocket forward. The nozzle 2 will accelerate and guide the gas, so that the gas is ejected at a suitable angle and speed, maximizing the thrust output, providing the power required for the rocket to lift off and fly, enabling it to overcome the earth's gravity and air resistance, and achieving the predetermined flight trajectory.

[0028] For those skilled in the art, it is obvious that the present invention is not limited to the details of the above exemplary embodiments, and the present invention can be implemented in other specific forms without departing from the spirit or basic characteristics of the present invention. Therefore, from any point of view, the embodiments should be regarded as exemplary and non-limiting. The scope of the present invention is defined by the appended claims rather than the above description. Therefore, all changes falling within the meaning and scope of the equivalent elements of the claims are intended to be included in the present invention.

Claims

1. An integrated direct combustion driven 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 rotating shaft, a turbine and a main rotating shaft; wherein the oxygen impeller, the fuel impeller, the auxiliary rotating shaft and the turbine are sequentially sleeved on the rotating 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 rotating shaft; The outer side of the turbine is rotatably connected to the bottom side wall of the combustion chamber via a first swivel bearing ring; The turbine is 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 driven liquid rocket engine according to claim 1, characterized in that: The oblique spray 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 rotatably connected to the turbine via a fourth bearing ring.

3. The integrated direct combustion driven 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 its interior is connected to the oxidant flow channel.

4. The integrated direct combustion driven 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 driven liquid rocket engine according to claim 1, characterized in that: The top end of the auxiliary shaft is rotatably connected to the inner side of the top of the combustion chamber through a third bearing ring. The gap between the auxiliary shaft and the main shaft is a fuel inlet and is connected to the fuel flow channel.

6. The integrated direct combustion driven 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.

7. The integrated direct combustion driven 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.

8. The integrated direct combustion driven 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.

9. The integrated direct combustion driven liquid rocket engine according to claim 1, characterized in that: A fuel injection nozzle is provided on the secondary rotating shaft at the rear end of the fuel guide vane, and the fuel injection nozzle is connected with the fuel flow channel and the combustion chamber.

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

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