direct-drive liquid rocket engine

Abstract

The present invention discloses a direct-drive liquid rocket engine, comprising: 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. The simplified system structure of the present invention reduces the requirements for component precision and manufacturing process, reduces the number of components, and reduces material cost and manufacturing cost; the oxidizer of the present invention is fully vaporized before reaching the nozzle, and the fuel is also fully vaporized into fuel gas through multi-point heat exchange during the flow process. The two are mixed more evenly with the fuel, 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; 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 startup, shutdown, and thrust adjustment.

Description

Technical Field

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

[0002] Liquid rocket engines are currently widely used in the aerospace industry. Traditional liquid rocket engines are typically equipped with independent turbine vane pumps and gas generators. The gas generator consumes additional fuel and oxidizer to produce high-temperature, high-pressure gas to drive the turbine vane pump and achieve fuel and oxidizer delivery and pressurization. This not only complicates the engine system structure, increasing manufacturing and maintenance costs, but also reduces engine reliability and poor adaptability under different operating conditions. Furthermore, the complex system design also limits the mass production of rocket engines. Furthermore, the oxidizer delivery process in traditional engines is not efficient enough during the preheating and gasification process, affecting combustion efficiency. Fuel delivery also suffers from insufficient heat exchange and low combustion efficiency. Furthermore, the connection method between the various components of traditional engines is prone to component wear, affecting engine stability and service life.

[0003] While existing technologies have made improvements, they still haven't achieved optimal technical results. For example, the invention patent with publication number CN114060170A relates to an open staged combustion pumped-cycle liquid rocket engine, consisting of a set of turbopumps, a staged combustion thrust chamber, and corresponding control valves. The staged combustion thrust chamber uses a two-stage combustion mode with an upper chamber and a lower chamber. Part of the oxidant and all of the fuel are combined in a rich combustion process in the upper chamber of the thrust chamber, forming a relatively low-temperature rich combustion zone. The resulting rich combustion gas is then introduced into the lower chamber of the thrust chamber, where it undergoes post-combustion combustion with the majority of the oxidant introduced from the oxygen pump.

[0004] The aforementioned invention patent still retains a precombustion chamber and independent turbopump. Therefore, this open staged combustion liquid rocket engine not only suffers from turbine blade exhaust losses and relatively low performance, but also remains a complex system with high manufacturing costs. Furthermore, the precombustion chamber pumps air to drive the turbine, resulting in low kinetic energy in the combustion gas.

[0005] The present invention aims to develop a direct-drive liquid rocket engine without a separate turbopump, completely eliminating the precombustion chamber, and integrating the fuel and oxidizer pressurization functions within the main combustion chamber. This new direct-drive liquid rocket engine features a simple and reliable system, strong adaptability, and ultra-low manufacturing costs. It addresses the complex, high-cost, and low-reliability issues of existing liquid rocket engines, while also improving the preheating and gasification efficiency of the fuel and oxidizer, enhancing combustion performance, and optimizing fuel delivery and heat exchange. 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 mounted at the upper end of the body, and the thrust chamber is located below the combustion chamber;

[0008] An annular bipropellant injector is installed below the rotor assembly at the top of the combustion chamber, an oxidant flow channel is formed between the bipropellant injector and the rotor assembly, and an oxidant impeller is installed in the oxidant flow channel;

[0009] 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;

[0010] The turbine, fuel impeller and oxidizer impeller are fixedly connected coaxially with the main shaft of the internal cavity;

[0011] The turbine is provided with turbine blades along its circumference, the turbine and the turbine blades both have 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;

[0012] 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;

[0013] Preferably, the outer side of the rotor assembly is rotatably connected to the top of the body via a rolling bearing, and the top of the combustion chamber is circumferentially connected to the top of the body via a plane bearing;

[0014] Preferably, the nozzle ring is in communication with the combustion chamber and the thrust chamber;

[0015] Preferably, the internal cavity of the main shaft is used to transport fuel, and the top end thereof is a fuel inlet;

[0016] Preferably, the outer side of the upper end of the main shaft is connected to the top of the body via a sealing ring;

[0017] Preferably, the interior of the combustion chamber side wall is a cavity and forms a fuel cooling channel;

[0018] Preferably, the interior of the side wall of the machine body is a cavity and forms an oxidant cooling channel;

[0019] 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;

[0020] Preferably, the oxidant flow channel is interconnected with the oxidant cooling channel and the oxidant nozzle respectively.

[0021] Compared with the prior art, the advantages of the present invention are:

[0022] The present invention completely abandons the gas generator and independent turbine vane pump, and integrates the oxidizer impeller, fuel impeller and turbine vanes into the main combustion chamber, which greatly simplifies the engine system structure, reduces failure points, and significantly improves the reliability of the engine;

[0023] The simplified system structure of the present invention reduces the requirements on component precision and manufacturing process, reduces the number of components, and reduces material and manufacturing costs;

[0024] 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 more evenly with the fuel, and the combustion reaction is more complete, which effectively improves the combustion efficiency of the engine and further improves the thrust and specific impulse performance of the engine.

[0025] Due to the simple structure and fewer transmission links, the present invention can respond to control instructions more quickly during operations such as starting, yaw and thrust adjustment.

[0026] 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

[0027] 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.

[0028] Figure 1 It is a structural schematic diagram of the engine of the present invention.

[0029] Figure 2 It is a structural schematic diagram of the combustion chamber of the present invention. DETAILED DESCRIPTION

[0030] 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.

[0031] 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.

[0032] 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.

[0033] 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.

[0034] See also Figures 1 and 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.

[0035] In this embodiment of the present invention, an annular bipropellant injector 5 is mounted below the rotor assembly 27 at the top of the combustion chamber 2. The bipropellant injector 5 is equipped with a fuel nozzle 51 and an oxidizer nozzle 52. An oxidizer flow channel 100 is formed between the bipropellant injector 5 and the rotor assembly 27, and an oxidizer impeller 21 is mounted within the oxidizer flow channel 100. A turbine 23 is mounted at the bottom of the combustion chamber 2. Turbine blades 230 are mounted along the circumference of the turbine 23. A separator plate 24 is mounted above the turbine 23. A fuel flow channel 200 is formed between the turbine 23 and the separator plate 24, and a fuel impeller 22 is mounted within the fuel flow channel 200. The turbine 23, fuel impeller 22, and oxidizer impeller 21 are coaxially and fixedly connected to the main shaft 25 of the internal cavity.

[0036] The nozzle ring 26 is installed around the isolation plate 24 above the turbine blade 230 for ejecting combustion gas and driving the turbine blade 230 , and the nozzle ring 26 is in communication with the combustion chamber 2 and the thrust chamber 3 .

[0037] 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.

[0038] In the above-described embodiment of the present invention, fuel is delivered through the internal cavity of the main shaft 25, with the top end serving as the fuel inlet. This creates an efficient fuel delivery channel within the main shaft 25. As the fuel flows within the main shaft 25, it undergoes heat exchange with the main shaft 25, which is subject to the high temperatures of the combustion chamber 2. The interior of the main shaft 25 communicates with the interior of the turbine 23, the interior of the turbine blades 230, and the fuel flow channel 200, allowing the fuel to rapidly fill the turbine 23, the turbine blades 230, and the fuel flow channel 200 under the action of the fuel impeller 22. Within the turbine 23 and turbine blades 230, the fuel not only cools them to prevent damage from the impact of high-temperature combustion gas but also absorbs heat, causing the fuel to vaporize. Simultaneously, the fuel flow channel 200 communicates with the fuel cooling channel 300. Within the fuel flow channel 200 and the fuel cooling channel 300, the fuel absorbs heat from the isolation plate 24 and the sidewalls of the combustion chamber 2, further vaporizing and increasing the temperature of the fuel. Subsequently, after multi-point heat exchange, the vaporized combustion gas flows to the fuel nozzle 51. This integrated design of fuel transmission and cooling enables the fuel to complete sufficient preheating and gasification processes before reaching the nozzle, greatly improving the efficiency of the mixed combustion of fuel and oxidizer.

[0039] Therefore, the fuel undergoes multi-point heat exchange during its flow. This series of heat exchange processes fully vaporizes the fuel into combustion gas. Compared to rocket engines where liquid fuel enters the combustion chamber 2 directly, the present invention's gaseous fuel and oxidizer mix more evenly, significantly improving combustion efficiency. Compared to staged combustion rocket engines, the present invention's simpler structure and fewer transmission links enable faster response to control commands during operations such as startup, shutdown, and thrust adjustment.

[0040] 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, reducing a large number of pipes, valves and complex connection structures. It not only achieves structural simplification and lightweight, but also the fuel gasification process fully absorbs the heat of various engine components, realizes efficient heat conversion, reduces energy loss, and significantly improves engine performance.

[0041] In this embodiment of the present invention, the interior of the sidewall of the engine body 1 is hollow and forms an oxidant cooling channel 400. The oxidant flow channel 100 is interconnected with the oxidant cooling channel 400 and the oxidant nozzle 52. Through this design, the oxidant can absorb heat from the oxidant flow channel 100 and the oxidant cooling channel 400 before reaching the oxidant nozzle 52, achieving full gasification and conversion into oxygen, which is then transported to the oxidant nozzle 52, creating conditions for efficient combustion.

[0042] 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.

[0043] Working principle of the present invention:

[0044] During engine operation, fuel and oxidizer are delivered from their respective storage tanks into the engine. The oxidizer enters the oxidizer flow channel 100 and flows under the influence of the oxidizer impeller 21. During this process, it absorbs heat generated by the engine's operation from the oxidizer flow channel 100. Simultaneously, the oxidizer absorbs heat from the sidewalls of the engine body 1 through the interconnection between the oxidizer flow channel 100 and the oxidizer cooling channel 400. As this heat is absorbed, the oxidizer gradually vaporizes to form oxygen, which is then delivered to the oxidizer nozzle 52 of the bipropellant injector 5.

[0045] The fuel enters the internal cavity of the main shaft 25 through the inlet at the top end. As it flows through the main shaft 25, it exchanges heat with the main shaft 25, which is affected by the high temperature of the combustion chamber 2. Under the action of the fuel impeller 22, the fuel rapidly 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 damage from the impact of the high-temperature combustion gas, but also absorbs heat to achieve fuel vaporization. At the same time, the fuel flow channel 200 is interconnected with the cooling channel. The fuel absorbs heat from the isolation plate 24 and the side walls of the combustion chamber 2 in the fuel flow channel 200 and the cooling channel, respectively, causing the fuel to further vaporize and increase its temperature. After multi-point heat exchange, the vaporized combustion gas flows to the fuel nozzle 51.

[0046] After being ejected from the fuel nozzle 51 and the oxidizer nozzle 52, respectively, the fuel and oxidizer are thoroughly mixed and ignited within the combustion chamber 2, producing high-temperature, high-pressure combustion gas. The combustion gas is ejected through the nozzle ring 26, impacting the turbine blades 230, and driving the turbine 23 to rotate. Because the turbine 23, the fuel impeller 22, and the oxidizer impeller 21 are coaxially fixedly connected to the main shaft 25, the rotation of the turbine 23 drives the fuel impeller 22 and the oxidizer impeller 21 to rotate synchronously, continuously providing power for the transportation of fuel and oxidizer. The combustion gas then enters the thrust chamber 3, where it is ejected at high speed to generate thrust, propelling the rocket into flight.

[0047] In the next embodiment of the present invention, a rolling bearing 6 is installed on the outside of the rotor assembly 27 and is rotatably connected 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.

[0048] 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.

[0049] 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. A direct-drive liquid rocket engine comprising: airframe, 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 bipropellant injector is installed below the rotor assembly at the top of the combustion chamber, an oxidant flow channel is formed between the bipropellant 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 oxidizer impeller are fixedly connected coaxially with the main shaft of the internal cavity; The turbine is provided with turbine blades along its circumference, the turbine and the turbine blades both have 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; The interior of the combustion chamber side wall is a cavity and forms a fuel cooling channel. The two-component injector is provided with a fuel nozzle and an oxidant nozzle. The fuel cooling channel is interconnected with the fuel nozzle and the fuel flow channel respectively.

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

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 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.

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 transporting 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 side wall of the machine body is a cavity and forms an oxidant cooling channel.

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

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