Extruded single stage liquid rocket

Abstract

The application discloses a kind of extrusion type first-stage liquid rocket, comprising: liquid nitrogen tank, liquid oxygen tank, fuel tank, thrust chamber, liquid nitrogen booster turbine pump, vaporizer around being arranged in the outer wall of thrust chamber, relevant valve and pipeline;The top of the thrust chamber is equipped with double-group element injector.Hollow shaft blade transports liquid nitrogen to actively cool turbine end, adopt turbine end internal cavity to realize efficient heat exchange, cooperate with the micro-hole of turbine end peripheral wall and open turbine blade cooling, effectively alleviate the scouring and damage of high-temperature gas to turbine end, realize liquid nitrogen booster turbine pump super multiple test run is not damaged.Liquid nitrogen booster turbine pump only carries out booster operation to a small amount of liquid nitrogen, without bearing the high-pressure pumping task of liquid propellant, can directly use civil-grade turbocharger pump to replace traditional high-cost aerospace-grade turbine pump as the liquid nitrogen booster turbine pump of the application, so that the manufacturing cost and research and development cost of rocket are significantly controlled, and the threshold of large-scale application is greatly reduced.

Description

Technical Field

[0001] This invention specifically relates to a squeeze-type first-stage liquid rocket. Background Technology

[0002] Currently, launch vehicles include two main types: solid fuel and liquid fuel. Solid fuel rockets have disadvantages such as high cost and low payload capacity. Furthermore, the combustion of propellants is uncontrollable, making it difficult to adjust thrust. Therefore, they are not suitable for first-stage rocket missions that require high payload accuracy and thrust flexibility.

[0003] Liquid rockets mostly employ pump-fed rocket engines, which require complex turbopumps to pressurize the fuel and dedicated burners to drive them. This results in a complex overall structure, extremely high precision requirements for components, low reliability, and exorbitant manufacturing costs. Currently, aerospace-grade turbopumps can cost tens of millions of dollars. This persistently high manufacturing cost not only limits the number of rocket launches and hinders the improvement of payload capacity but also significantly impacts the rocket development process. For example, the use of expensive aerospace-grade turbopumps during test runs, coupled with frequent test runs, can easily damage the turbopumps, leading to a substantial increase in development costs and restricting the large-scale application of liquid rockets.

[0004] For example, the invention patent with patent number CN2022110167540 abandons the traditional gas generator of the pump-fed engine and makes further improvements by directly extracting gas from the combustion chamber to drive the turbine. However, this solution still has the following inherent defects and is difficult to apply to the high reliability and low cost requirements of the first-stage rocket: ① Due to the high pressure of 15-20 MPa in the combustion chamber, the existing technology uses a turbopump to directly pressurize the liquid propellant. The turbopump needs to be pressurized to an ultra-high pressure of over 20 MPa, which requires the turbopump to operate at ultra-high speeds, with speeds of 15,000-80,000 rpm. This easily leads to harsh operating conditions for the impeller pump, with the oxygen pump being more susceptible to oxidation and corrosion, which in turn causes damage to the turbopump. At the same time, the pump chambers are isolated by dynamic seals. Under high-speed rotation, the seals are prone to wear and failure, causing fuel and oxidizer to leak and premix in the pump chamber. This can easily lead to catastrophic safety accidents such as spontaneous combustion, deflagration, and pump body explosion, seriously threatening the safety of rocket launches.

[0005] ② The pressurization of liquid rocket propellant mainly relies on turbopumps. The high load of turbopumps keeps their operating speed between 15,000 and 80,000 rpm. Moreover, the turbopumps are driven by high-temperature gas extracted from the combustion chamber, which can reach temperatures as high as 3,000°C. The high temperature of the gas combined with the high speed of the turbine makes the turbine end extremely susceptible to ultra-high temperature, resulting in rapid aging and damage. This further reduces the reliability of the system and increases maintenance costs and failure risks.

[0006] While traditional extrusion-type liquid rockets rely on high-pressure gas to pressurize the propellant, allowing it to directly enter the thrust chamber for combustion, offering advantages such as low cost, simple structure, and high reliability, the need for overall pressurization of the liquid fuel tank requires the tank structure to withstand extremely high internal pressure. This leads to a significant increase in the rocket's exhaust weight during flight, reducing its effective payload capacity. Furthermore, the upper limit of the tank's internal pressure directly restricts the propellant delivery flow rate, resulting in significantly lower engine combustion chamber pressure compared to turbopump-based systems. This, in turn, leads to a substantial decrease in the engine's specific impulse, further limiting the rocket's payload capacity.

[0007] The aforementioned drawbacks have led to traditional squeeze-type engines not being the mainstream design for current first-stage rockets, but rather being used only in a few final-stage rockets for spaceflight missions (where the requirements for thrust and payload capacity are relatively low). Therefore, this invention aims to provide a squeeze-type first-stage rocket that does not directly pump propellant and does not require high-pressure gas cylinders. This squeeze-type first-stage rocket not only possesses high thrust delivery capability but also significantly improves system reliability and drastically reduces manufacturing costs, solving many technical problems inherent in existing technologies. Summary of the Invention

[0008] To overcome the shortcomings mentioned above, the present invention aims to provide a technical solution that can solve the above problems.

[0009] A squeeze-type first-stage liquid rocket includes: a liquid nitrogen tank, a liquid oxygen tank, a fuel tank, a thrust chamber, a liquid nitrogen booster turbopump, a vaporizer surrounding the outer wall of the thrust chamber, related valves and pipes; a bicomponent injector is provided on the top of the thrust chamber. In this process, the full flow of liquid oxygen and the full flow of fuel directly enter the thrust chamber for combustion to generate thrust. The turbine end of the liquid nitrogen booster turbopump is equipped with multiple turbine blades along its peripheral wall, and a small amount of high-temperature gas is directly drawn out from the thrust chamber to drive the turbine blades. A small portion of the liquid nitrogen from the liquid nitrogen tank outlet is pressurized via a branch pipeline through the pump end of the liquid nitrogen booster turbopump and then transported to the vaporizer to exchange heat with the outer wall of the thrust chamber. The outlet of the vaporizer is connected to the top air cushion space of the liquid nitrogen tank via a pipeline. Most of the liquid nitrogen from the liquid nitrogen tank outlet enters the cooling channel inside the thrust chamber pipe wall via the main pipeline. The outlet of this cooling channel is connected to the top air cushion space of both the liquid oxygen tank and the fuel tank. The turbine end and the pump end are driven by a hollow shaft. The hollow shaft has internal blades and an internal cavity. The interior of the hollow shaft connects the liquid inlet of the pump end with the internal cavity of the turbine end. The peripheral wall of the turbine end is provided with micropores between any two turbine blades; Preferably, the air intake direction of the inner blade is completely consistent with the air intake direction of the pump end; Preferably, the outlets of the liquid oxygen tank and the fuel tank are both connected to the bi-component injector, an oxidant regulating valve is connected in series between the liquid oxygen tank and the bi-component injector, and a fuel regulating valve is connected in series between the fuel tank and the bi-component injector. Preferably, the oxidizer regulating valve and the fuel regulating valve are both electrically connected to the rocket control system to adjust the ratio and flow of liquid oxygen and fuel in real time, thereby achieving thrust adjustment under varying operating conditions. Preferably, a nitrogen pressure control valve is connected in series between the outlet of the vaporizer and the top air cushion space of the liquid nitrogen tank; an oxygen pressure control valve is connected in series between the outlet of the cooling channel and the liquid oxygen tank; and a fuel pressure control valve is connected in series between the outlet of the cooling channel and the top air cushion space of the fuel tank. Preferably, the nitrogen pressure control valve, oxygen pressure control valve, and fuel pressure control valve are all automatically adjusted by the rocket control system.

[0010] Compared with the prior art, the advantages of the present invention are: This invention utilizes the internal blades of a hollow shaft to deliver liquid nitrogen for active cooling of the turbine end, employs the internal cavity of the turbine end for efficient heat exchange, and combines this with micropores on the turbine end's peripheral wall to cool the turbine blades. This effectively mitigates the scouring and damage to the turbine end caused by high-temperature combustion gases, significantly improves the operational reliability of the liquid nitrogen booster turbopump and even the entire propulsion system, extends the equipment's service life, and ensures that the liquid nitrogen booster turbopump remains undamaged even after numerous test runs.

[0011] The liquid nitrogen booster turbopump of this invention only boosts a small amount of liquid nitrogen, without having to undertake the high-pressure pumping task of liquid propellant. This significantly reduces the workload and operating speed of the turbopump, effectively simplifies the overall structure of the pump body and reduces the precision requirements of the components. As a result, it enables the application of civilian-grade turbopumps. Civilian-grade turbopumps can be directly used to replace the traditional, expensive aerospace-grade turbopumps as the liquid nitrogen booster turbopump of this invention, which significantly controls the manufacturing and R&D costs of rockets and greatly lowers the threshold for large-scale application.

[0012] Furthermore, this invention utilizes the synergistic effect of a liquid nitrogen booster turbopump and a vaporizer to compress and boost liquid nitrogen through the superposition of boosting pressure and vaporization expansion force, forming high-pressure liquid nitrogen. Most of the high-pressure liquid nitrogen enters the cooling channel of the thrust chamber tube wall, where it undergoes heat exchange and vaporization expansion to form ultra-high-pressure nitrogen gas. This ultra-high-pressure nitrogen gas is used to boost the pressure of both liquid oxygen and fuel propellants, ensuring that the two propellants can be forced into the thrust chamber at sufficient pressure and flow rate for thorough mixing and combustion, thereby achieving high thrust output of the rocket. This perfectly matches the core requirements of the first-stage rocket for high thrust, high reliability, and low cost.

[0013] This invention completely eliminates the complex propellant pumping structure of traditional pump-fed rockets, and does not directly pump liquid oxygen and fuel at high pressure. This fundamentally avoids the wear and damage to the equipment caused by ultra-high pressure and ultra-high speed operation. At the same time, it effectively avoids safety hazards such as propellant leakage in the pump chamber and spontaneous combustion and deflagration caused by premixing, and significantly improves the operational safety of the rocket propulsion system.

[0014] Meanwhile, this invention eliminates the need for high-pressure gas cylinders required in traditional squeeze-type rockets. It innovatively utilizes the heat exchange and vaporization process between liquid nitrogen and the outer wall of the thrust chamber to provide stable pressure for the liquid nitrogen tank, liquid oxygen tank, and fuel tank, achieving squeeze-type propellant delivery. This design reduces the rocket's flight weight due to high-pressure gas cylinders, increasing effective payload capacity, and effectively improves propellant delivery flow rate, ensuring combustion efficiency in the thrust chamber and engine specific impulse, thus compensating for the performance shortcomings of traditional squeeze-type rockets.

[0015] Additional aspects and advantages of the invention will be set forth in part in the description which follows, and in part will be obvious from the description, or may be learned by practice of the invention. Attached Figure Description

[0016] To more clearly illustrate the technical solutions in the embodiments of the present invention or the prior art, the drawings used in the description of the embodiments or the prior art will be briefly introduced below. Obviously, the drawings described below are only some embodiments of the present invention. For those skilled in the art, other drawings can be obtained based on these drawings without creative effort.

[0017] Figure 1 This is a schematic diagram of the structure of the present invention.

[0018] Figure 2 This is a cross-sectional view of the turbine end and the hollow shaft of the present invention.

[0019] Figure 3 This is a three-dimensional schematic diagram of the turbine end and hollow shaft of the present invention.

[0020] In the diagram: 1-Liquid nitrogen tank; 2-Liquid oxygen tank; 3-Fuel tank; 4-Thrust chamber; 5-Liquid nitrogen booster turbopump; 51-Turbine end; 511-Turbine blade; 512-Micropore; 52-Pump end; 53-Hollow shaft; 531-Inner blade; 6-Carburetor; 7-Bi-component injector; 8-Branch pipe; 9-Main pipe; 10-Cooling passage; 11-Oxidant regulating valve; 12-Fuel regulating valve; 13-Nitrogen pressure control valve; 14-Oxygen pressure control valve; 15-Fuel pressure control valve; 16-Extraction pipe. Detailed Implementation

[0021] The technical solutions in the embodiments of the present invention will be clearly and completely described below. Obviously, the described embodiments are only some embodiments of the present invention, and not all embodiments. Based on the embodiments of the present invention, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of the present invention.

[0022] In the description of this invention, it should be noted that the terms "center", "upper", "lower", "left", "right", "vertical", "horizontal", "inner", "outer", etc., indicate the orientation or positional relationship based on the orientation or positional relationship shown in the accompanying drawings. They are only for the convenience of describing this invention and simplifying the description, and do not indicate or imply that the device or element referred to must have a specific orientation, or be constructed and operated in a specific orientation. Therefore, they should not be construed as limitations on this invention.

[0023] Furthermore, in the description of this invention, unless otherwise explicitly specified and limited, the terms "installation," "connection," "linking," etc., should be interpreted broadly. For example, they can refer to a fixed connection, a detachable connection, or an integral connection; they can refer to a mechanical connection or an electrical connection; they can refer to a direct connection or an indirect connection through an intermediate medium; they can refer to the internal connection of two components; they can refer to a wireless connection or a wired connection. Those skilled in the art can understand the specific meaning of the above terms in this invention based on the specific circumstances.

[0024] Furthermore, the technical features involved in the different embodiments of the present invention described below can be combined with each other as long as they do not conflict with each other.

[0025] Please see Figures 1-2 In this embodiment of the invention, a squeeze-type first-stage liquid rocket, such as... Figure 1 As shown, its core components include a liquid nitrogen tank 1, a liquid oxygen tank 2, a fuel tank 3, a thrust chamber 4, a liquid nitrogen booster turbopump 5, a carburetor 6, a bi-component injector 7, and related valves and connecting pipes such as an oxidizer regulating valve 11, a fuel regulating valve 12, a nitrogen pressure control valve 13, an oxygen pressure control valve 14, and a fuel pressure control valve 15. The connection relationships and working principles of each component are as follows: The liquid nitrogen tank 1, liquid oxygen tank 2, and fuel tank 3 are all made of stainless steel to ensure stable operation under the preset air cushion pressure. The thrust chamber 4 is made of high-temperature resistant 304 stainless steel. The dual-component injector 7 on the top of the thrust chamber 4 adopts a centrifugal injection structure to ensure that liquid oxygen and fuel are fully atomized and mixed, thereby improving combustion efficiency. The turbine end 51 and pump end 52 of the liquid nitrogen booster turbopump 5 are coaxially arranged and driven by a hollow shaft 53. The hollow shaft 53 has an integrally formed inner blade 531. The air intake direction of the inner blade 531 is completely consistent with the air intake direction of the impeller of the pump end 52, ensuring that when the pump end rotates to boost liquid nitrogen, the liquid nitrogen in the inner cavity of the turbine end 51 can also be boosted. Multiple turbine blades 511 are evenly arranged on the circumferential wall of the turbine end 51. The turbine blades 511 are inclined in the opposite direction to the stationary blades below them, and the two change the flow direction of the airflow to generate driving force.

[0026] like Figures 2-3 As shown, micro-holes 512 are formed on the peripheral wall of the turbine end 51 between any two turbine blades 511, with a diameter of 0.8mm to 1.5mm. When the turbine end 51 is working, a very small amount of liquid nitrogen in its inner cavity is ejected through the micro-holes 512 and forms a cooling gas film to cool the turbine blades 511. The vaporizer 6 adopts a spiral coil structure, tightly wound and fixed to the outer wall of the thrust chamber 4. A thermally conductive insulating pad is set between the vaporizer 6 pipeline and the outer wall of the thrust chamber 4 to balance heat exchange efficiency and structural thermal insulation protection, preventing high temperature from being conducted to other components. It should be noted that the liquid nitrogen booster turbopump 5 in this embodiment is a civilian-grade turbopump, whose manufacturing or procurement cost is only less than 1% of that of a traditional aerospace-grade turbopump, and its structure is simple and easy to maintain, fully meeting the working requirements of this invention.

[0027] During the launch phase, the rocket control system issues a launch command, first opening the relevant valves at the outlet of liquid nitrogen tank 1. Under the influence of gravity, a small amount of liquid nitrogen (approximately 10%–20%) enters the pump end 52 of the liquid nitrogen booster turbopump 5 through the manifold pipe 8, while the majority of liquid nitrogen (approximately 80%–90%) enters the cooling channel 10 within the thrust chamber 4 through the main pipe 9. Simultaneously, the valves at the outlets of liquid oxygen tank 2 and fuel tank 3 are opened. Under the influence of gravity, liquid oxygen and fuel are respectively delivered to the bicomponent injector 7 through the oxidizer regulating valve 11 and the fuel regulating valve 12. After atomization, they enter the thrust chamber 4 and are ignited by the ignition device, starting combustion to generate high-temperature, high-pressure gas. The pressure inside the combustion chamber is maintained at 6.0 MPa–7.4 MPa.

[0028] During takeoff, a small amount of high-temperature, high-pressure combustion gas (approximately 5%–8%) generated in the thrust chamber 4 is directly extracted and transported via the extraction pipe 16 to the turbine end 51 of the liquid nitrogen booster turbopump 5. Under the action of the turbine blades 511 and the stationary blades, this small amount of high-temperature combustion gas drives the turbine blades 511 to rotate, which in turn drives the impeller of the pump end 52 to rotate via the hollow shaft 53, thus pressurizing the small amount of liquid nitrogen. At this time, the pump end 52 pressurizes the small amount of liquid nitrogen transported from the branch pipe 8 (to 1.5 MPa–2.0 MPa). The pressurized liquid nitrogen is transported to the vaporizer 6 and exchanges heat with the outer wall of the thrust chamber 4. After absorbing the heat from the thrust chamber 4, the liquid nitrogen vaporizes to form high-pressure nitrogen gas. The pressure of the high-pressure nitrogen gas is regulated by the nitrogen pressure control valve 13 (maintained at 0). After being compressed to pressures of 8MPa to 1.2MPa, the liquid nitrogen is delivered to the top air cushion space of the liquid nitrogen tank 1, providing a stable air cushion pressure and ensuring continuous liquid nitrogen output. Simultaneously, most of the liquid nitrogen delivered to the cooling channel 10 via the main pipeline 9 absorbs heat from the inner wall of the thrust chamber 4, achieving cooling protection for the thrust chamber 4 and preventing damage due to high temperatures. The heated liquid nitrogen partially vaporizes, forming a gas-liquid mixture. After pressure regulation by the oxygen pressure control valve 14 and the fuel pressure control valve 15, this mixture is delivered to the top air cushion spaces of the liquid oxygen tank 2 and the fuel tank 3, respectively. This provides extremely high compression pressure for the delivery of liquid oxygen and fuel, ensuring a stable, full-flow delivery of liquid oxygen and fuel to the bicomponent injector 7, where it continuously burns with the high-temperature combustion gases to generate thrust. During this process, the ultra-high-pressure nitrogen gas formed by the vaporization of high-pressure liquid nitrogen effectively ensures a large-flow delivery of propellant, enabling the first-stage rocket to output high thrust and launch.

[0029] During the sustained operation phase of the rocket, the working principle is the same as that of the high-thrust output of the compression-type first-stage rocket, so it will not be elaborated on here.

[0030] During this process, the rocket control system collects pressure signals from the gas cushion spaces of liquid nitrogen tank 1, liquid oxygen tank 2, and fuel tank 3 in real time, as well as the flow rates of liquid oxygen and fuel. It automatically adjusts the valve openings of nitrogen pressure control valve 13, oxygen pressure control valve 14, and fuel pressure control valve 15 to ensure that the pressure of each gas cushion is maintained within the preset range. At the same time, by adjusting the openings of oxidizer regulating valve 11 and fuel regulating valve 12, it adjusts the ratio and flow of liquid oxygen and fuel in real time to achieve thrust adjustment under different operating conditions, adapting to the thrust requirements of different flight stages such as rocket takeoff and climb.

[0031] Furthermore, when the turbine end 51 is working, the inner blades 531 inside the hollow shaft 53 rotate with the hollow shaft 53, drawing a portion of the liquid nitrogen from the inlet of the pump end 52 into the hollow shaft 53 and delivering it to the inner cavity of the turbine end 51 to cool the inside of the turbine end 51, absorbing the heat transferred by the high-temperature combustion gas, and further reducing the temperature of the turbine end 51; the cavity inside the turbine end 51 forms a heat exchange space, further improving the heat exchange efficiency between the liquid nitrogen and the turbine end; at the same time, the micropores 512 on the peripheral wall of the turbine end 51 can discharge a very small amount of liquid nitrogen from the inner cavity of the turbine end 51, forming a cooling gas film above the turbine blades 511, carrying away some heat, ensuring the working stability of the turbine end 51, preventing it from being damaged due to high temperature and high pressure, and significantly improving the reliability of the system.

[0032] During the shutdown phase, the rocket control system issues a shutdown command. First, it closes the valves at the outlets of liquid oxygen tank 2 and fuel tank 3, stopping the delivery of liquid oxygen and fuel. The combustion gases in thrust chamber 4 are gradually burned out. Then, it closes the relevant valves at the outlet of liquid nitrogen tank 1, stopping the output of liquid nitrogen. The liquid nitrogen booster turbopump 5 gradually stops operating due to the loss of driving combustion gases and liquid nitrogen supply. Finally, it opens all pressure relief valves to slowly release the high-pressure gas in the gas cushion spaces of liquid nitrogen tank 1, liquid oxygen tank 2, and fuel tank 3 to atmospheric pressure, completing the shutdown process.

[0033] In this embodiment, the liquid nitrogen booster turbopump 5 only boosts a small amount of liquid nitrogen, and its speed is controlled at 3000-8000 rpm, which is much lower than the turbopump speed of existing pump-driven rockets. This significantly reduces the workload of the liquid nitrogen booster turbopump 5, avoids problems such as dynamic seal wear and impeller oxidation, and improves system reliability. At the same time, by utilizing the application of civilian-grade turbopumps, such as the manufacturing process of general industrial-grade cryogenic turbopumps, the production and manufacturing of this invention can be carried out, which can reduce the manufacturing cost of the liquid nitrogen booster turbopump 5 to less than 1%.

[0034] 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 implemented 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 exemplary and non-limiting, and the scope of the invention is defined by the appended claims rather than the foregoing description. Thus, it is intended that all variations falling within the meaning and scope of equivalents of the claims be included within the present invention.

Claims

1. A squeeze-type first-stage liquid rocket, characterized in that, include: Liquid nitrogen tank, liquid oxygen tank, fuel tank, thrust chamber, liquid nitrogen booster turbopump, vaporizer surrounding the outer wall of the thrust chamber, related valves and pipelines; a dual-component injector is provided on the top of the thrust chamber; In this process, the full flow of liquid oxygen and the full flow of fuel directly enter the thrust chamber for combustion to generate thrust. The turbine end of the liquid nitrogen booster turbopump is equipped with multiple turbine blades along its peripheral wall, and a small amount of high-temperature gas is directly drawn out from the thrust chamber to drive the turbine blades. A small portion of the liquid nitrogen from the liquid nitrogen tank outlet is pressurized via a branch pipeline through the pump end of the liquid nitrogen booster turbopump and then transported to the vaporizer to exchange heat with the outer wall of the thrust chamber. The outlet of the vaporizer is connected to the top air cushion space of the liquid nitrogen tank via a pipeline. Most of the liquid nitrogen from the liquid nitrogen tank outlet enters the cooling channel inside the thrust chamber pipe wall via the main pipeline. The outlet of this cooling channel is connected to the top air cushion space of both the liquid oxygen tank and the fuel tank. The turbine end and the pump end are driven by a hollow shaft. The hollow shaft has internal blades and an internal cavity. The interior of the hollow shaft connects the liquid inlet of the pump end with the internal cavity of the turbine end. The peripheral wall of the turbine end is provided with micropores between any two turbine blades.

2. The extrusion-type first-stage liquid rocket according to claim 1, characterized in that, The air intake direction of the inner blades is completely consistent with the air intake direction of the pump end.

3. The extrusion-type first-stage liquid rocket according to claim 1, characterized in that, The outlets of the liquid oxygen tank and the fuel tank are both connected to the bi-component injector. An oxidant regulating valve is connected in series between the liquid oxygen tank and the bi-component injector, and a fuel regulating valve is connected in series between the fuel tank and the bi-component injector.

4. The extrusion-type first-stage liquid rocket according to claim 3, characterized in that, The oxidizer regulating valve and the fuel regulating valve are both electrically connected to the rocket control system and are used to adjust the ratio and flow of liquid oxygen and fuel in real time to achieve thrust adjustment under different operating conditions.

5. The extrusion-type first-stage liquid rocket according to claim 1, characterized in that, A nitrogen pressure control valve is connected in series between the outlet of the vaporizer and the top air cushion space of the liquid nitrogen tank; an oxygen pressure control valve is connected in series between the outlet of the cooling channel and the liquid oxygen tank; and a fuel pressure control valve is connected in series between the outlet of the cooling channel and the top air cushion space of the fuel tank.

6. The extrusion-type first-stage liquid rocket according to claim 5, characterized in that, The nitrogen pressure control valve, oxygen pressure control valve, and fuel pressure control valve are all automatically adjusted by the rocket control system.

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