Direct-driven liquid rocket engine

Abstract

The invention discloses a direct-drive liquid rocket engine which comprises an engine body, a combustion chamber and a thrust chamber. The combustion chamber is installed at the upper end of the interior of the engine body, and the thrust chamber is located below the combustion chamber. According to the simplified system structure, the requirements for part precision and the manufacturing process are reduced, the number of parts is reduced, and the material cost and the manufacturing cost are reduced; the oxidant is fully gasified before reaching the nozzle, the fuel is fully gasified into fuel gas through multi-point heat exchange in the flowing process, the oxidant and the fuel are mixed with the fuel more uniformly, the combustion reaction is more sufficient, the combustion efficiency of the engine is effectively improved, and the thrust and specific impulse performance of the engine are further improved; compared with a staged combustion rocket engine, due to the fact that the structure is simple and the number of transmission links is small, a control instruction can be responded more quickly during starting, shutdown, thrust adjustment and other operations.

Description

Technical Field

[0001] The invention relates to a direct-drive liquid rocket engine. Background Art

[0002] At present, liquid rocket engines are widely used in the aerospace field. Traditional liquid rocket engines are usually equipped with independent turbine vane pumps and gas generators. The gas generator consumes additional fuel and oxidant to produce high-temperature and high-pressure gas to drive the turbine vane pump to achieve the delivery and pressurization of fuel and oxidant. This not only leads to a complex structure of the engine system, increases manufacturing and maintenance costs, but also reduces the reliability of the engine, and has poor adaptability under different working conditions. In addition, the complex system design also limits the mass production process of rocket engines. At the same time, in the process of oxidant delivery, the preheating and gasification of the oxidant in the traditional engine is not efficient enough, which affects the combustion efficiency; in terms of fuel delivery, there are also problems such as insufficient heat exchange and low combustion efficiency. In addition, the connection method between the various components of the traditional engine is prone to component wear, affecting the stability and service life of the engine.

[0003] Although the existing technology has made improvements, it still fails to achieve the best technical effect. For example, the invention patent with publication number CN114060170A relates to an open staged combustion exhaust cycle liquid rocket engine, which is composed of a set of turbo pumps, a staged combustion thrust chamber and corresponding control valves. The staged combustion thrust chamber adopts a two-stage combustion mode of upper chamber and lower chamber. Part of the oxidant and all the fuel are organized to burn rich in the upper chamber of the thrust chamber to form a relatively low temperature rich combustion zone. The generated rich combustion gas is then introduced into the lower chamber of the thrust chamber to carry out supplementary combustion with most of the oxidant introduced from the oxygen pump.

[0004] The above invention patent still retains the precombustion chamber and independent turbine pump, so the open staged combustion liquid rocket engine not only has turbine blade exhaust loss and relatively low performance, but also the system is still complex and the manufacturing cost is high. In addition, the precombustion chamber extracts air to drive the turbine, and its combustion gas has the problem of low kinetic energy.

[0005] The purpose of the present invention is to propose a direct-drive liquid rocket engine without an independent turbo pump, completely abandoning the pre-combustion chamber, and integrating the fuel and oxidant pressurization functions in the main combustion chamber. This new direct-drive liquid rocket engine has the characteristics of simple and reliable system, strong adaptability, and ultra-low manufacturing cost. It solves the problems of complex system, high cost and low reliability of existing liquid rocket engines, while improving the preheating gasification efficiency of fuel and oxidant, enhancing the combustion effect, and optimizing the fuel transportation and heat exchange process. Summary of the invention

[0006] In order to overcome the above-mentioned shortcomings, the present invention aims to provide a technical solution that can solve the above-mentioned problems.

[0007] A direct-drive liquid rocket engine comprises: a body, a combustion chamber, and a thrust chamber; the combustion chamber is installed at the upper end of the body, and the thrust chamber is located below the combustion chamber; An annular bicomponent injector is installed below the rotor assembly at the top of the combustion chamber, an oxidant flow channel is formed between the bicomponent injector and the rotor assembly, and an oxidant impeller is installed in the oxidant flow channel; A turbine is installed at the bottom of the combustion chamber, an isolation plate is installed above the turbine, a fuel flow channel is formed between the turbine and the isolation plate, and a fuel impeller is installed in the fuel flow channel; The turbine, fuel impeller and oxidant impeller are coaxially fixedly connected to the main shaft of the internal cavity; The turbine is provided with turbine blades along its circumference, the turbine and the turbine blades are both internal cavities, and the interior of the main shaft is interconnected with the interior of the turbine, the interior of the turbine blades and the fuel flow channel respectively; Preferably, a nozzle ring is installed above the turbine blade and between the combustion chamber side wall and the isolation plate, for spraying gas and driving the turbine blade; Preferably, the outer side of the rotor assembly is rotatably connected to the top of the body through a rolling bearing, and the top of the combustion chamber is circumferentially connected to the top of the body through a plane bearing; Preferably, the nozzle ring is interconnected with the combustion chamber and the thrust chamber; Preferably, the internal cavity of the main shaft is used to transport fuel, and its top end is a fuel inlet; Preferably, the outer side of the upper end of the main shaft is connected to the top of the body through a sealing ring; Preferably, the interior of the combustion chamber side wall is a cavity and forms a fuel cooling channel; Preferably, the interior of the side wall of the body is a cavity and forms an oxidant cooling channel; Preferably, the bipropellant injector is provided with a fuel nozzle and an oxidant nozzle, and the fuel cooling channel is interconnected with the fuel nozzle and the fuel flow channel respectively; Preferably, the oxidant flow channel is interconnected with the oxidant cooling channel and the oxidant nozzle respectively.

[0008] Compared with the prior art, the advantages of the present invention are: The present invention completely abandons the gas generator and the independent turbine blade pump, and integrates the oxidant impeller, the fuel impeller and the turbine blade into the main combustion chamber, which greatly simplifies the engine system structure, reduces the failure points, and significantly improves the reliability of the engine; The simplified system structure of the present invention reduces the requirements on component accuracy and manufacturing process, reduces the number of components, and reduces material cost and manufacturing cost; The oxidant of the present invention is fully gasified before reaching the nozzle, and the fuel is also fully gasified into fuel gas through multi-point heat exchange during the flow process. The two are mixed with the fuel more evenly, and the combustion reaction is more complete, which effectively improves the combustion efficiency of the engine, thereby improving the thrust and specific impulse performance of the engine; The present invention has a simple structure and fewer transmission links, and can respond to control instructions more quickly during operations such as starting, yaw, and thrust adjustment.

[0009] Additional aspects and advantages of the present invention will be given in part in the following description and in part will be obvious from the following description, or will be learned through practice of the present invention. BRIEF DESCRIPTION OF THE DRAWINGS

[0010] In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the drawings required for use in the embodiments or the description of the prior art will be briefly introduced below. 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 creative labor.

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

[0012] Figure 2 It is a structural schematic diagram of the combustion chamber of the present invention. DETAILED DESCRIPTION

[0013] The technical solutions in the embodiments of the present invention are described clearly and completely below. Obviously, the described embodiments are only part of the embodiments of the present invention, not all of them. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without creative work are within the scope of protection of the present invention.

[0014] In the description of the present invention, it should be noted that the terms "center", "up", "down", "left", "right", "vertical", "horizontal", "inside", "outside", etc. indicate directions or positional relationships 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 limitations on the present invention.

[0015] In addition, in the description of the present invention, unless otherwise clearly specified and limited, the terms "installed", "connected", "connected", 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 or an electrical connection; it can be a direct connection, or it can be indirectly connected through an intermediate medium, it can also be the internal communication of two components, it can be a wireless connection, or it can be a wired connection. For ordinary technicians in this field, the specific meanings of the above terms in the present invention can be understood according to specific circumstances.

[0016] 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.

[0017] See also Figure 1-2 In the embodiment of the present invention, the direct-drive liquid rocket engine proposed by the present invention is mainly composed of a body 1, a combustion chamber 2 and a thrust chamber 3. The combustion chamber 2 is installed at the upper end of the body 1, and the thrust chamber 3 is located below the combustion chamber 2. In the embodiment of the present invention, an annular bipropellant injector 5 is installed below the rotor assembly 27 at the top of the combustion chamber 2, and the bipropellant injector 5 is provided with a fuel nozzle 51 and an oxidant nozzle 52. An oxidant flow channel 100 is formed between the bipropellant injector 5 and the rotor assembly 27, and an oxidant impeller 21 is installed in the oxidant flow channel 100. A turbine 23 is installed at the bottom of the combustion chamber 2, and turbine blades 230 are installed along the circumference of the turbine 23. An isolation plate 24 is installed above the turbine 23, and a fuel flow channel 200 is formed between the turbine 23 and the isolation plate 24, and a fuel impeller 22 is installed in the fuel flow channel 200. The turbine 23, the fuel impeller 22 and the oxidant impeller 21 are coaxially fixedly connected to the main shaft 25 of the internal cavity.

[0018] The nozzle ring 26 is installed on the circumference of the isolation plate 24 above the turbine blade 230 to spray the combustion gas and drive the turbine blade 230 , and the nozzle ring 26 is communicated with the combustion chamber 2 and the thrust chamber 3 .

[0019] In an embodiment of the present invention, the internal cavity of the main shaft 25 is used to transport fuel, and its top end is a fuel inlet. A fuel flow channel 200 is formed between the turbine 23 and the isolation plate 24. The turbine 23 and the turbine blade 230 are both internal cavities. The interior of the main shaft 25 is respectively interconnected with the interior of the turbine 23, the interior of the turbine blade 230 and the fuel flow channel 200, and the interior of the side wall of the combustion chamber 2 is a cavity and forms a fuel cooling channel 300. The fuel cooling channel 300 is respectively interconnected with the fuel nozzle 51 and the fuel flow channel 200.

[0020] In the above embodiment of the present invention, the present invention delivers fuel through the internal cavity of the main shaft 25, and the top end serves as the fuel inlet, and an efficient fuel delivery channel is formed inside the main shaft 25. When the fuel flows in the main shaft 25, it exchanges heat with the main shaft 25 affected by the high temperature of the combustion chamber 2. The inside of the main shaft 25 is respectively connected with the inside of the turbine 23, the inside of the turbine blade 230 and the fuel flow channel 200, so that the fuel can quickly fill the inside of the turbine 23, the turbine blade 230 and the fuel flow channel 200 under the action of the fuel impeller 22; inside the turbine 23 and the turbine blade 230, the fuel not only cools them to prevent damage due to the impact of high-temperature combustion gas, but also absorbs heat to achieve fuel gasification; at the same time, the fuel flow channel 200 is interconnected with the fuel cooling channel 300, and the fuel absorbs the heat of the isolation plate 24 and the side wall of the combustion chamber 2 in the fuel flow channel 200 and the fuel cooling channel 300, respectively, so that the fuel is further gasified and heated, and then the gas formed after multi-point heat exchange and gasification flows to the fuel nozzle 51. This integrated design of fuel transmission and cooling allows the fuel to complete the sufficient preheating and gasification process before reaching the nozzle, greatly improving the efficiency of the mixed combustion of fuel and oxidant.

[0021] Therefore, the fuel realizes multi-point heat exchange during the flow process, and this series of heat exchange processes fully gasifies the fuel into fuel gas. Compared with the rocket engine in which the liquid fuel directly enters the combustion chamber 2, the gaseous fuel and the oxidant of the present invention are mixed more evenly, which can greatly improve the combustion efficiency; compared with the staged combustion rocket engine, the present invention has a simple structure and fewer transmission links, and can respond to control instructions more quickly during operations such as starting, shutting down, and adjusting thrust.

[0022] The present invention abandons the staged combustion system such as the turbo pump and the pre-combustion chamber, and directly utilizes the gas generated by the main combustion chamber 2 to drive the turbine 23, thereby reducing a large number of pipelines, valves and complex connection structures. It not only realizes the simplification and lightweight of the structure, but also the fuel gasification process fully absorbs the heat of various engine components, realizes the efficient conversion of heat, reduces energy loss, and thus significantly improves the engine performance.

[0023] In the embodiment of the present invention, the interior of the side wall of the body 1 is a cavity and forms an oxidant cooling channel 400, and the oxidant flow channel 100 is respectively interconnected with the oxidant cooling channel 400 and the oxidant nozzle 52. Through this design, before the oxidant reaches the oxidant nozzle 52, it can absorb the heat of the oxidant flow channel 100 and the oxidant cooling channel 400, achieve full gasification, convert into oxygen, and be transported to the oxidant nozzle 52, creating conditions for efficient combustion.

[0024] The fuel gas and oxygen are injected into the main combustion chamber 2 at the same time and ignited, and the combustible gas and the combustion-supporting gas are mixed and burned to achieve the most efficient combustion mode.

[0025] Working principle of the present invention: When the engine is working, the fuel and the oxidant are transported from the storage tank to the engine respectively. The oxidant enters the oxidant flow channel 100 and flows under the action of the oxidant impeller 21. In this process, the heat generated by the operation of the engine in the oxidant flow channel 100 is absorbed. At the same time, the oxidant absorbs the heat transmitted from the side wall of the body 1 through the interconnection structure between the oxidant flow channel 100 and the oxidant cooling channel 400. As the heat is absorbed, the oxidant gradually gasifies to form oxygen, which is transported to the oxidant nozzle 52 of the bicomponent injector 5.

[0026] The fuel enters the internal cavity of the main shaft 25 through the inlet at the top of the main shaft 25. When flowing in the main shaft 25, it exchanges heat with the main shaft 25 affected by the high temperature of the combustion chamber 2. Under the action of the fuel impeller 22, the fuel quickly fills the interior of the turbine 23, the turbine blades 230 and the fuel flow channel 200. Inside the turbine 23 and the turbine blades 230, the fuel not only cools them to prevent them from being damaged by the impact of high-temperature combustion gas, but also absorbs heat to achieve fuel gasification. At the same time, the fuel flow channel 200 is interconnected with the cooling channel, and the fuel absorbs the heat of the isolation plate 24 and the side wall of the combustion chamber 2 in the fuel flow channel 200 and the cooling channel respectively, so that the fuel is further gasified and heated. After multi-point heat exchange, the gasified combustion gas flows to the fuel nozzle 51. After the fuel and oxidant are ejected from the fuel nozzle 51 and the oxidant nozzle 52 respectively, they are fully mixed and ignited in the combustion chamber 2 to produce high-temperature and high-pressure combustion gas. The combustion gas is ejected through the nozzle ring 26, impacts the turbine blade 230, and drives the turbine 23 to rotate. Since the turbine 23, the fuel impeller 22 and the oxidant impeller 21 are coaxially fixedly connected to the main shaft 25, the rotation of the turbine 23 drives the fuel impeller 22 and the oxidant impeller 21 to rotate synchronously, continuously providing power for the transportation of fuel and oxidant. Subsequently, the combustion gas enters the thrust chamber 3, is ejected at high speed to generate thrust, and drives the rocket to fly.

[0027] In the next embodiment of the present invention, the present invention installs a rolling bearing 6 on the outside of the rotor assembly 27 and rotatably connects it to the top of the body; a plane bearing 7 is installed around the top of the combustion chamber to achieve a stable connection with the top of the body.

[0028] In the next embodiment of the present invention, the sealing ring 8 between the outer side of the upper end of the main shaft 25 and the top of the engine body 1 effectively prevents the leakage of the oxidant and ensures the safe operation of the engine.

[0029] 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 present invention can be implemented in other specific forms without departing from the spirit or essential features of the present invention. Therefore, the embodiments should be considered exemplary and non-restrictive in all respects, and the scope of the present invention is defined by the appended claims rather than the above description, and it is intended that all changes falling within the meaning and scope of the equivalent elements of the claims be included in the present invention.

Claims

1. A direct-drive liquid rocket engine, comprising: fuselage, combustion chamber, and thrust chamber; The combustion chamber is installed at the upper end of the body, and the thrust chamber is located below the combustion chamber, and is characterized in that: An annular bicomponent injector is installed below the rotor assembly at the top of the combustion chamber, an oxidant flow channel is formed between the bicomponent injector and the rotor assembly, and an oxidant impeller is installed in the oxidant flow channel; A turbine is installed at the bottom of the combustion chamber, an isolation plate is installed above the turbine, a fuel flow channel is formed between the turbine and the isolation plate, and a fuel impeller is installed in the fuel flow channel; The turbine, fuel impeller and oxidant impeller are coaxially fixedly connected to the main shaft of the internal cavity; The turbine is circumferentially provided with turbine blades, the turbine and the turbine blades are both internal cavities, and the interior of the main shaft is respectively interconnected with the interior of the turbine, the interior of the turbine blades and the fuel flow channel.

2. The direct-drive liquid rocket engine according to claim 1, characterized in that: A nozzle ring is installed above the turbine blade and between the combustion chamber side wall and the isolation plate for spraying combustion gas and driving the turbine blade.

3. The direct-drive liquid rocket engine according to claim 1, characterized in that: The outer side of the rotor assembly is rotatably connected to the top of the machine body through a rolling bearing, and the top of the combustion chamber is circumferentially connected to the top of the machine body through a plane bearing.

4. The direct-drive liquid rocket engine according to claim 2, characterized in that: The nozzle ring is communicated with the combustion chamber and the thrust chamber.

5. The direct-drive liquid rocket engine according to claim 1, characterized in that: The inner cavity of the main shaft is used for conveying fuel, and the top end of the main shaft is a fuel inlet.

6. The direct-drive liquid rocket engine according to claim 1, characterized in that: The outer side of the upper end of the main shaft is connected to the top of the machine body through a sealing ring.

7. The direct-drive liquid rocket engine according to claim 1, characterized in that: The interior of the combustion chamber side wall is a cavity and forms a fuel cooling channel.

8. The direct-drive liquid rocket engine according to claim 1, characterized in that: The interior of the side wall of the machine body is a cavity and forms an oxidant cooling channel.

9. The direct-drive liquid rocket engine according to claim 1, characterized in that: The two-component injector is provided with a fuel nozzle and an oxidant nozzle, and the fuel cooling channel is communicated with the fuel nozzle and the fuel flow channel respectively.

10. The direct-drive liquid rocket engine according to claim 1, characterized in that: The oxidant flow channel is interconnected with the oxidant cooling channel and the oxidant nozzle respectively.

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