Combustion chamber extraction cycle liquid rocket engine

Abstract

The present invention relates to the technical field of rocket power systems, and discloses a combustion chamber exhaust circulation liquid rocket engine, comprising: an engine body, the engine body comprising a turbine pump and a combustion chamber, the turbine pump and the combustion chamber are connected by a gas conduit, and the turbine pump is also provided with an oxidant inlet, a fuel inlet, a turbine collecting ring and a turbine nozzle; a main shaft, the main shaft is rotatably installed in the turbine pump, and the main shaft is also installed with a fuel pump impeller, a turbine and an oxygen pump impeller; the present invention uses a coaxial main shaft to drive two pumps, so that the flow rate and head of the fuel and oxidant are always proportional, which is convenient for the stability and thrust adjustment of the system, and can greatly reduce the research and development and debugging cycle, directly taking out the gas from the combustion chamber to drive the turbine, greatly simplifying the structure, and only the combustion chamber of the entire system needs to be ignited, avoiding the problem of ignition asynchronism caused by multiple igniters, simplifying the system workflow and improving reliability.

Description

Technical Field

[0001] The invention belongs to the technical field of rocket power systems, and in particular to a combustion chamber exhaust cycle liquid rocket engine. Background Art

[0002] The main components of the existing gas generator cycle liquid rocket engine are: gas generator and accessories, thrust chamber and turbo pump. The gas generator is equivalent to a small liquid rocket engine, which requires special valves, pipelines, injectors, combustion chambers and igniters, etc. The structure is complex and the reliability is poor.

[0003] Therefore, in view of the above situation, there is an urgent need to provide a combustion chamber exhaust cycle liquid rocket engine to overcome the shortcomings in current practical applications. Summary of the invention

[0004] In view of the above situation, in order to overcome the defects of the prior art, the present invention provides a combustion chamber extraction circulation liquid rocket engine, which directly takes out the fuel gas from the combustion chamber and cools it down through injection to drive the turbine pump, which can greatly reduce the structural weight and production cost, and has the advantages of simple structure, light weight, high reliability, etc.

[0005] To achieve the above object, the present invention provides the following technical solution: a combustion chamber exhaust circulation liquid rocket engine, the combustion chamber exhaust circulation liquid rocket engine comprising:

[0006] An engine body, the engine body comprising a turbine pump and a combustion chamber, the turbine pump and the combustion chamber being connected via a gas conduit, and the turbine pump is also provided with an oxidant inlet, a fuel inlet, a turbine gas collecting ring and a turbine nozzle;

[0007] The main shaft is rotatably mounted in the turbine pump. A fuel pump impeller, a turbine and an oxygen pump impeller are also mounted on the main shaft. The oxygen pump impeller is located at one end of the oxidant inlet, and the fuel pump impeller is located at one end of the fuel inlet.

[0008] Preferably, a plurality of groups of injection cooling nozzles are also provided on the combustion chamber.

[0009] Preferably, it also includes a combustion chamber exhaust port for taking out high-pressure fuel gas in the combustion chamber.

[0010] Preferably, a regulating valve is provided in the combustion chamber exhaust port. By adjusting the opening of the combustion chamber exhaust regulating valve, the gas flow rate flowing through the turbine pump can be adjusted, and the rotation speed of the main shaft and the impeller also changes accordingly, thereby realizing the adjustment of the propellant flow rate and the thrust.

[0011] Preferably, both ends of the main shaft are rotatably mounted in the turbine pump via a first bearing and a second bearing respectively.

[0012] Preferably, a second dynamic sealing system and a first dynamic sealing system are further provided between the main shaft and the turbine pump, and the sealing performance can be improved by the coordinated arrangement of the second dynamic sealing system and the first dynamic sealing system.

[0013] Preferably, it also includes a turbine exhaust pipe, which is installed on the turbine pump and is connected to the turbine pump for discharging exhaust gas from the turbine 12.

[0014] Preferably, the first bearings are provided in four groups.

[0015] Preferably, the second bearings are provided in two groups.

[0016] Preferably, the main shaft is made of high-strength material.

[0017] Compared with the prior art, the present invention has the following beneficial effects:

[0018] The turbine uses a common main shaft to drive two pumps, so that the flow and head of the fuel and oxidizer are always proportional, which is convenient for system stability and thrust adjustment, and can greatly reduce the R&D and debugging cycle;

[0019] The gas is directly taken out from the combustion chamber to drive the turbine, eliminating the need for a gas generator and corresponding accessories, greatly simplifying the structure;

[0020] In the whole system, only the combustion chamber needs to be ignited, which avoids the problem of ignition asynchrony caused by multiple igniters, simplifies the system workflow and improves reliability;

[0021] The gas flow rate can be adjusted by adjusting the opening of the combustion chamber exhaust regulating valve. Since the gas flow rate flowing through the regulating valve is small, the required actuator power is also small, which is beneficial to the system's response speed and automatic control. BRIEF DESCRIPTION OF THE DRAWINGS

[0022] The accompanying drawings are used to provide further understanding of the present invention and constitute a part of the specification. They are used to explain the present invention together with the embodiments of the present invention and do not constitute a limitation of the present invention.

[0023] In the attached picture:

[0024] Figure 1 It is a schematic structural diagram of an embodiment of the present invention.

[0025] Figure 2 It is a partial enlarged view of an embodiment of the present invention.

[0026] In the figure: 1-engine body, 2-combustion chamber exhaust port, 3-injection cooling nozzle, 4-oxidant inlet, 5-oxygen pump impeller, 6-first dynamic sealing system, 7-first bearing, 8-fuel inlet, 9-second dynamic sealing system, 10-fuel pump impeller, 11-second bearing, 12-turbine, 13-main shaft, 14-turbine exhaust pipe, 15-regulating valve, 16-gas duct, 17-turbine gas collecting ring, 18-turbine nozzle. DETAILED DESCRIPTION

[0027] The technical solutions in the embodiments of the present invention will be clearly and completely described below in conjunction with the drawings in the embodiments of the present invention. Obviously, the described embodiments are only part of the embodiments of the present invention, rather than all the embodiments; based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without making creative work are within the scope of protection of the present invention.

[0028] See also Figure 1 and Figure 2 The combustion chamber exhaust cycle liquid rocket engine provided by the embodiment of the present invention comprises:

[0029] An engine body 1, the engine body 1 comprises a turbopump and a combustion chamber, the turbopump and the combustion chamber are connected via a gas conduit 16, and the turbopump is also provided with an oxidant inlet 4, a fuel inlet 8, a turbine gas collecting ring 17 and a turbine nozzle 18;

[0030] The main shaft 13 is rotatably mounted in the turbine pump. The main shaft 13 is also mounted with a fuel pump impeller 10, a turbine 12 and an oxygen pump impeller 5. The oxygen pump impeller 5 is located at one end of the oxidant inlet 4, and the fuel pump impeller 10 is located at one end of the fuel inlet 8.

[0031] In an embodiment of the present invention, the oxidant and the fuel enter the turbopump through the oxidant inlet 4 and the fuel inlet 8, flow through the impeller and then through the injector, and burn in the combustion chamber. A small portion of the combustion products enter the turbopump through the side opening of the combustion chamber and the gas duct 16 to drive the turbine 12 to rotate, and drive the oxygen pump and the fuel pump to pressurize the propellant and further increase the internal pressure. Most of the high-temperature gas is finally ejected at high speed from the combustion chamber outlet to generate thrust.

[0032] In one embodiment of the present invention, see Figure 1 The combustion chamber is also provided with a plurality of groups of injection cooling nozzles 3.

[0033] In this embodiment, the gas temperature can be lowered by providing multiple groups of injection cooling nozzles 3 .

[0034] In one embodiment of the present invention, see Figure 1, and also includes a combustion chamber exhaust port 2 for taking out high-pressure gas in the combustion chamber;

[0035] A regulating valve 15 is provided in the combustion chamber air exhaust port 2.

[0036] In this embodiment, the high-pressure fuel gas extracted from the combustion chamber exhaust port 2 is introduced into the turbine collecting ring 17 through the fuel gas conduit 16, and its direction can be changed through the turbine nozzle 18. By adjusting the opening of the combustion chamber exhaust regulating valve 15, the fuel gas flow rate flowing through the turbine pump can be adjusted, and the rotation speed of the main shaft 13 and the impeller also changes accordingly, thereby realizing the adjustment of the propellant flow rate and the thrust.

[0037] In one embodiment of the present invention, see Figure 1 and Figure 2 , the two ends of the main shaft 13 are rotatably mounted in the turbine pump through the first bearing 7 and the second bearing 11 respectively;

[0038] A second dynamic sealing system 9 and a first dynamic sealing system 6 are further provided between the main shaft 13 and the turbine pump.

[0039] In this embodiment, the second dynamic sealing system 9 and the first dynamic sealing system 6 are arranged in coordination to improve the sealing performance and avoid gas leakage.

[0040] In one embodiment of the present invention, see Figure 1 , and also includes a turbine exhaust pipe 14, which is installed on the turbine pump and is connected to the turbine pump for discharging the exhaust gas of the turbine 12.

[0041] In one embodiment of the present invention, see Figure 1 and Figure 2 The first bearings 7 are provided in four groups, which can improve the stability of the main shaft 13 during rotation.

[0042] In one embodiment of the present invention, see Figure 1 and Figure 2 The second bearings 11 are provided in two groups, which can improve the stability of the main shaft 13 when rotating.

[0043] In one embodiment of the present invention, see Figure 1 The main shaft 13 is made of high-strength material, which can improve the service life of the engine, and the main shaft 13 and other components are made of high-temperature resistant materials.

[0044] In this embodiment, in order to achieve the same speed of the two pumps or a certain stable speed ratio relationship, a flexible coupling or a gear, a pulley or a magnetic coupling can be used;

[0045] In order to reduce the temperature of the combustion gas flowing through the turbine, a heat exchanger or the like may be used, which is not specifically limited here.

[0046] Working principle: oxidant and fuel enter the turbo pump through the oxidant inlet 4 and the fuel inlet 8, flow through the impeller and then pass through the injector, and burn in the combustion chamber. A small part of the combustion products enters the turbo pump through the side opening of the combustion chamber and the gas duct to drive the turbine 12 to rotate (this small part of the gas drives the turbine 12 and is discharged through the turbine exhaust pipe 14), and drives the oxygen pump and the fuel pump to pressurize the propellant and further increase the pressure in the combustion chamber. Most of the high-temperature gas is finally ejected at high speed from the combustion chamber outlet to generate thrust; the turbine uses a common main shaft to drive two pumps, so that the flow rate and lift of the fuel and oxidant are The process is always proportional, which is convenient for the stability and thrust adjustment of the system and can greatly reduce the R&D and debugging cycle; the gas-driven turbine 12 is directly taken out from the combustion chamber, eliminating the gas generator and corresponding accessories, greatly simplifying the structure; only the combustion chamber needs to be ignited in the entire system, avoiding the problem of ignition asynchronism caused by multiple igniters, simplifying the system workflow and improving reliability; the gas flow rate can be adjusted by adjusting the opening of the combustion chamber exhaust regulating valve. Since the gas flow rate flowing through the regulating valve is small, the required actuator power is also small, which is beneficial to the response speed and automatic control of the system.

[0047] It should be noted that, in this article, relational terms such as first and second, etc. are only used to distinguish one entity or operation from another entity or operation, and do not necessarily require or imply any such actual relationship or order between these entities or operations. Moreover, the terms "include", "comprise" or any other variants thereof are intended to cover non-exclusive inclusion, so that a process, method, article or device including a series of elements includes not only those elements, but also other elements not explicitly listed, or also includes elements inherent to such process, method, article or device.

[0048] Although embodiments of the present invention have been shown and described, it will be appreciated by those skilled in the art that various changes, modifications, substitutions and variations may be made to the embodiments without departing from the principles and spirit of the present invention, and that the scope of the present invention is defined by the appended claims and their equivalents.

Claims

1. A combustion chamber exhaust cycle liquid rocket engine, comprising an engine body (1), the engine body (1) comprising a turbopump and a combustion chamber, the turbopump and the combustion chamber being connected via a gas conduit (16), and the turbopump is also provided with an oxidant inlet (4), a fuel inlet (8), a turbine gas collecting ring (17) and a turbine nozzle (18), It is characterized in that Also includes: A main shaft (13), the main shaft (13) is rotatably mounted in the turbine pump, and a fuel pump impeller (10), a turbine (12) and an oxygen pump impeller (5) are also mounted on the main shaft (13), the oxygen pump impeller (5) is located at one end of the oxidant inlet (4), and the fuel pump impeller (10) is located at one end of the fuel inlet (8); The combustion chamber is also provided with a plurality of groups of injection cooling nozzles (3); It also includes a combustion chamber exhaust port (2) for taking out high-pressure gas in the combustion chamber; A regulating valve (15) is provided in the combustion chamber air exhaust port (2); The two ends of the main shaft (13) are rotatably mounted in the turbine pump via a first bearing (7) and a second bearing (11) respectively; A second dynamic sealing system (9) and a first dynamic sealing system (6) are also provided between the main shaft (13) and the turbine pump; It also includes a turbine exhaust pipe (14), wherein the turbine exhaust pipe (14) is mounted on the turbine pump and is in communication with the turbine pump for discharging exhaust gas from the turbine (12); The first bearings (7) are provided in four groups; The second bearings (11) are provided in two groups; The main shaft (13) is made of high-strength material; The oxidant and the fuel enter the turbo pump through the oxidant inlet (4) and the fuel inlet (8), flow through the impeller and then pass through the injector, and burn in the combustion chamber. A small part of the combustion products enter the turbo pump through the side opening of the combustion chamber and the gas conduit to drive the turbine (12) to rotate. This small part of the gas drives the turbine (12) and is discharged through the turbine exhaust pipe (14), and drives the oxygen pump and the fuel pump to pressurize the propellant and further increase the pressure in the combustion chamber. Most of the high-temperature gas is finally ejected at high speed from the combustion chamber outlet to generate thrust. The turbine uses a common main shaft to drive the two pumps, so that the fuel and oxygen The flow rate and lift of the chemical agent are always proportional, which is convenient for the stability and thrust adjustment of the system and reduces the R&D and debugging cycle; the gas-driven turbine (12) is directly taken out from the combustion chamber, eliminating the gas generator and corresponding accessories and simplifying the structure; in the entire system, only the combustion chamber needs to be ignited, avoiding the ignition asynchronism problem caused by multiple igniters, simplifying the system workflow and improving reliability; the gas flow is adjusted by adjusting the opening of the combustion chamber exhaust regulating valve. Since the gas flow passing through the regulating valve is small, the required actuator power is also small, which is beneficial to the response speed and automatic control of the system.

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