Squeezing type first-stage rocket capable of achieving high thrust output

Abstract

The invention discloses an extrusion type first-stage rocket capable of achieving high thrust output. The extrusion type first-stage rocket comprises a liquid nitrogen storage box, a liquid nitrogen turbine pump, an oxidizing agent storage box, a fuel storage box, a thrust chamber and a vaporizer. The turbine pump only pressurizes and conveys liquid nitrogen, and does not make contact with an oxidizing agent and fuel in the whole process, so that the potential safety hazards of propellant inner leakage, premixing spontaneous combustion and pump body explosion caused by pump body dynamic sealing failure are fundamentally eradicated; meanwhile, the problem of oxidation corrosion damage of the oxygen pump impeller is avoided, and the operation safety and reliability of the system are fundamentally improved; the vaporizer is used for vaporizing liquid nitrogen and cooling the pipe wall of the thrust chamber, compared with a cooling fuel flow channel in the prior art, the vaporizer structure achieves lightweight design, the overall weight and the manufacturing cost of the thrust chamber spray pipe are remarkably reduced, the liquid nitrogen directly makes contact with the pipe wall of the thrust chamber for heat exchange, and the heat exchange efficiency is improved. And the heat exchange efficiency is also synchronously improved.

Description

Technical Field

[0001] This invention relates to a first-stage rocket, specifically a squeeze-type liquid rocket, which can achieve high thrust output. Background Technology

[0002] Liquid rocket propulsion systems are the core power units of space launch vehicles, and the main propulsion methods are divided into two categories: extrusion propulsion and pump-jet propulsion. Traditional extrusion propulsion systems rely on high-pressure helium or nitrogen cylinders to pressurize the propellant tank. They are simple in structure and highly reliable, but limited by cylinder pressure and tank pressure-bearing capacity, the propellant supply pressure is low and the thrust is small, which cannot meet the high thrust requirements of the first-stage rocket. At the same time, the high-pressure cylinders have thick walls and are heavy, which seriously reduces the launch efficiency and is only suitable for attitude control engines or low-thrust engines in the final stage.

[0003] Traditional pump-jet engines generate high-temperature gas through a gas generator and pre-combustion chamber to drive a turbopump, which directly pressurizes the oxidizer and fuel, achieving high thrust. However, they have significant drawbacks: the turbopump is in contact with both the oxidizer and fuel, making the dynamic seals prone to failure. Once internal leakage occurs, it can easily lead to propellant premixing, spontaneous combustion, or even deflagration. The system includes a pre-combustion chamber, igniter, and complex turbopump components, with many parts, high precision requirements, and high research and manufacturing costs.

[0004] For example, the invention patent with patent number CN2021112515598 discloses an open-stage combustion exhaust cycle first-stage rocket engine, in which the turbine is driven by fuel-rich gas drawn from the upper chamber of the thrust chamber, and the thrust chamber adopts a staged combustion mode with fuel-rich upper chamber and supplementary combustion in lower chamber. This scheme still has the following inherent defects.

[0005] ① The existing technology uses the same pump body to pump fuel and oxidant simultaneously, and the pump chambers are isolated only by dynamic seals. Under high-speed rotation, the seals are prone to wear and failure, which leads to leakage and premixing of fuel and oxidant in the pump chamber, causing catastrophic safety accidents such as spontaneous combustion, deflagration, and pump body explosion. Furthermore, the turbopump is directly used to pressurize liquid propellant, which easily leads to poor impeller pump operation, and the oxygen pump is more easily oxidized, which in turn leads to damage to the turbopump.

[0006] ② The existing technology uses a gunpowder starter or a high-pressure gas cylinder starter, which increases the complexity of the rocket engine structure. The high-pressure gas cylinder with high pressure and thick walls reduces the rocket's carrying efficiency.

[0007] ③ The turbopump is driven by the upper chamber of the thrust chamber rich in fuel (equivalent to a burner or pre-combustion chamber), which results in a complex structure. The peripheral wall of the thrust chamber also has fuel flow channels, which leads to an excessively large engine and an excessively thick peripheral wall, resulting in an excessively heavy engine. At the same time, the engine can only be manufactured by 3D printing and cannot be mass-produced, which leads to excessively high manufacturing costs.

[0008] Therefore, the present invention aims to provide a squeeze-type first-stage rocket that does not directly pump propellant, has no high-pressure gas cylinders, and has no independent pre-combustion chamber, which significantly improves system reliability and reduces manufacturing costs while ensuring high thrust output. Summary of the Invention

[0009] To overcome the shortcomings mentioned above, the present invention aims to provide a technical solution that can solve the above problems.

[0010] A squeeze-type first-stage rocket capable of high thrust output includes: a liquid nitrogen tank, a liquid nitrogen turbopump, an oxidizer tank, a fuel tank, a thrust chamber, and a vaporizer; The vaporizer is attached to the outer wall of the thrust chamber; The liquid nitrogen turbopump includes a liquid nitrogen pump and a turbine connected to the same drive shaft. The outlet of the liquid nitrogen tank is connected to the inlet of the liquid nitrogen pump through a nitrogen tank outlet control valve. The outlet of the liquid nitrogen pump is connected to the inlet of the vaporizer. The outlet of the vaporizer is connected to an oxygen tank pressurization control valve, a fuel tank pressurization control valve, and a nitrogen tank pressurization control valve. The other end of the oxygen tank pressurization control valve is connected to the pressurization diffuser inside the oxidant storage tank, and the other end of the fuel tank pressurization control valve is connected to the pressurization diffuser inside the fuel storage tank. Liquid nitrogen is pressurized by a liquid nitrogen pump and enters the vaporizer, where it absorbs residual heat from the thrust chamber wall and partially seeps into the thrust chamber wall through the sweat holes to form a thin film for cooling. After absorbing heat, it vaporizes into high-pressure nitrogen. A portion of the high-pressure nitrogen gas compresses the oxidizer tank and the fuel tank, forcing the oxidizer in the oxidizer tank and the fuel in the fuel tank into the thrust chamber for combustion; another portion of the high-pressure nitrogen gas compresses the liquid nitrogen surface in the liquid nitrogen tank, pushing the liquid nitrogen continuously into the liquid nitrogen pump, forming a closed loop. The turbine is driven by a small portion of high-temperature gas directly drawn from the thrust chamber, and the turbine's outlet is connected to the outside. Preferably, the vaporizer is composed of a liquid inlet pipe, an air outlet pipe and multiple annular grooves. The multiple annular grooves with C-shaped cross sections are interconnected with the liquid inlet pipe and the air outlet pipe. The C-shaped opening side of each annular groove is sealed and fitted to the outer wall of the thrust chamber, forming an annular heat exchange channel. Preferably, any one layer of the annular groove sleeve is formed by welding the upper ring plate, the lower ring plate and the middle ring cylinder together by conventional argon arc welding, and the three together form a C-shaped groove cavity structure with one side open; Preferably, the inner wall of the thrust chamber is provided with a plurality of sweat holes, which are distributed circumferentially along the inner wall of the thrust chamber, and the axis of the sweat holes forms an angle of 30 to 45° with the inner wall of the thrust chamber and faces the thrust chamber outlet. Any one of the sweat holes penetrates the wall of the thrust chamber and connects the thrust chamber with the interior of the vaporizer. Preferably, the pipeline between the outlet of the oxidizer tank and the oxygen injector of the thrust chamber is equipped with an oxygen tank outlet control valve, and the pipeline between the outlet of the fuel tank and the fuel injector of the thrust chamber is equipped with a fuel tank outlet control valve. Preferably, a gas extraction control valve is provided on the gas injection line from the thrust chamber to the turbine; Preferably, check valves are provided at the outlets of the oxygen tank pressurization control valve, the fuel tank pressurization control valve, and the nitrogen tank pressurization control valve.

[0011] Compared with the prior art, the advantages of the present invention are: The turbopump of this invention only pressurizes and delivers liquid nitrogen, without any contact with oxidizer or fuel throughout the process. This eliminates the safety hazards of propellant leakage, premixed spontaneous combustion, and pump body explosion caused by pump body dynamic seal failure, and also avoids the problem of oxygen pump impeller oxidation and corrosion damage. Fuel and oxidizer are directly squeezed into the thrust chamber, and neither propellant is connected by rotating components, fundamentally improving the safety and reliability of the system operation.

[0012] The three-way pressure control valve at the outlet of the vaporizer of this invention adjusts the pressure of the three storage tanks respectively, and realizes a reasonable distribution of extrusion pressure based on system feedback, thereby adjusting the flow ratio of fuel and oxidant.

[0013] The system of this invention completely eliminates the high-pressure gas cylinder, gunpowder starter and independent pre-combustion chamber structure in existing liquid rocket technology. It eliminates the need for high-pressure thick-walled gas cylinder, complex ignition components and special combustion chamber, greatly simplifying the system structure, significantly reducing the overall weight and manufacturing cost, while avoiding the safety risks and structural redundancy brought about by high-pressure gas cylinder.

[0014] The vaporizer of the present invention is used to vaporize liquid nitrogen and cool the tube wall of the thrust chamber. Compared with the prior art cooling fuel flow channel, the vaporizer has a lightweight design, which significantly reduces the overall weight and manufacturing cost of the thrust chamber nozzle. Furthermore, the liquid nitrogen directly contacts the thrust chamber tube wall for heat exchange, and its heat exchange efficiency is also improved.

[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 overall structure of the present invention.

[0018] Figure 2 This is a structural diagram of the thrust chamber of the present invention.

[0019] Figure 3 This is a perspective view of the thrust outer surface of the present invention integrated with other components. Detailed Implementation

[0020] 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.

[0021] 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.

[0022] 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.

[0023] 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.

[0024] Please see Figures 1-2 In this embodiment of the invention, a squeeze-type first-stage rocket capable of high thrust output is described. The rocket system has a simple overall structure, versatile processing technology, and is suitable for mass production while significantly reducing manufacturing costs. It includes: a liquid nitrogen tank 100, a liquid nitrogen turbopump, an oxidizer tank 300, a fuel tank 400, a thrust chamber 500, and a vaporizer 600.

[0025] In this invention, the liquid nitrogen turbopump adopts a coaxial integrated structure, with the liquid nitrogen pump 202 and the turbine 201 connected by the same drive shaft. The turbine 201 is driven by a small portion of high-temperature gas directly drawn from the thrust chamber 500. The outlet of the turbine 201 is connected to the outside. A gas injection pipeline 700 between the thrust chamber 500 and the turbine 201 is equipped with a gas extraction control valve 701, which is used to adjust the flow rate of the high-temperature gas entering the turbine 201, thereby controlling the speed of the liquid nitrogen turbopump and the overall pressurization pressure of the system.

[0026] In this invention, a nitrogen tank outlet control valve 101 is provided on the pipeline between the outlet of the liquid nitrogen storage tank 100 and the inlet of the liquid nitrogen pump 202 to control the on / off state of liquid nitrogen; an oxygen tank outlet control valve 301 is provided on the pipeline between the outlet of the oxidant storage tank 300 and the oxygen injector of the thrust chamber 500; and a fuel tank outlet control valve 401 is provided on the pipeline between the outlet of the fuel storage tank 400 and the fuel injector of the thrust chamber 500, so as to independently control the on / off state of oxidant and fuel entering the thrust chamber 500.

[0027] In this invention, the vaporizer 600 is attached to the outer wall of the thrust chamber 500. The vaporizer 600 consists of a liquid inlet pipe 601, an outlet pipe 602, and multiple layers of annular grooves 603. The multiple layers of C-shaped annular grooves 603 are interconnected with the liquid inlet pipe 601 and the outlet pipe 602. The C-shaped opening side of each annular groove 603 is sealed and fitted to the outer wall of the thrust chamber 500, forming an annular heat exchange channel 604. When liquid nitrogen flows through this annular heat exchange channel 604, the liquid nitrogen can directly contact the outer wall of the thrust chamber 500, increasing the heat exchange efficiency and enabling the liquid nitrogen to be completely vaporized.

[0028] In this invention, the outlet pipe 602 of the vaporizer 600 is divided into three branches, which are respectively connected to the oxygen tank booster control valve 302, the fuel tank booster control valve 402 and the nitrogen tank booster control valve 102; each booster control valve is equipped with a check valve (not shown) at its outlet end to prevent the medium in the storage tank from flowing back into the vaporizer 600 channel; the other end of each booster control valve is connected to the booster diffuser inside the corresponding storage tank, and high-pressure nitrogen is used to supply constant pressure to the storage tank. Example 1

[0029] Before startup, open the oxygen tank outlet control valve 301, fuel tank outlet control valve 401, and nitrogen tank outlet control valve 101 respectively. Under the original pressure of each tank, the fuel is sprayed into the thrust chamber 500 through the fuel injector, and the oxidizer is sprayed into the thrust chamber 500 through the oxidizer injector. At this time, the two propellants are mixed. At the same time, liquid nitrogen fills the liquid nitrogen pump 202.

[0030] At startup, the propellant inside the thrust chamber 500 is ignited, causing the thrust chamber 500 to rapidly expand at high temperature. Simultaneously, the exhaust gas control valve 701 is opened, and a small portion of the high-temperature gas is led to the gas ejector pipe 700. The gas is then driven to rotate by the ejector nozzle of the turbine 201, and then the high-temperature gas is directly ejected to the outside.

[0031] After startup, the turbine 201 drives the liquid nitrogen pump 202. Under the pressure of the tank and the suction of the liquid nitrogen pump 202, the liquid nitrogen enters the inlet pipe 601 of the vaporizer 600 and then flows into the annular heat exchange channel 604 formed by the U-shaped cross-section annular sleeves 603 of each layer. The liquid nitrogen absorbs the residual heat of the thrust chamber 500 wall. Some of the low-temperature liquid nitrogen seeps into the inner wall surface of the thrust chamber 500 through the sweat hole 50 to form a continuous thin film cooling layer, which provides efficient thermal protection for the thrust chamber 500. The remaining liquid nitrogen completely absorbs heat and vaporizes into high-pressure nitrogen gas in the heat exchange channel. The pressure of this high-pressure nitrogen gas can reach 20-50 MPa.

[0032] High-pressure nitrogen gas, via outlet pipe 602 and three pressure control valves, a check valve (not shown), and a pressure diffuser, pressurizes the oxidizer tank 300, fuel tank 400, and liquid nitrogen tank 100 through gas compression, maintaining stable internal pressure in each tank. Under the pressure of the high-pressure nitrogen gas, the oxidizer and fuel continuously enter the thrust chamber 500 at a high flow rate for mixing and combustion, producing stable high-temperature, high-pressure combustion gas, which is then injected outwards to generate high thrust output.

[0033] Under the pressure of the storage tank and the suction of the liquid nitrogen pump 202, the liquid nitrogen continuously circulates, providing high-pressure nitrogen compression to ensure the rocket engine can achieve high thrust output.

[0034] In this embodiment of the invention, the turbine 201 pump only pressurizes and delivers liquid nitrogen, without any contact with oxidizer or fuel throughout the process. This eliminates the safety hazards of propellant leakage, premixed spontaneous combustion, and pump body explosion caused by pump body dynamic seal failure, and also avoids the problem of oxygen pump impeller oxidation and corrosion damage. The fuel and oxidizer are directly squeezed into the thrust chamber 500, and neither propellant is connected by rotating components, thus fundamentally improving the safety and reliability of the system operation.

[0035] In this embodiment of the invention, the three-way pressure control valve at the outlet of the vaporizer 600 adjusts the pressure of the three storage tanks respectively, and realizes a reasonable distribution of extrusion pressure based on system feedback, thereby adjusting the flow ratio of fuel and oxidant.

[0036] The system of this invention completely eliminates the high-pressure gas cylinder, gunpowder starter and independent pre-combustion chamber structure in existing liquid rocket technology. It eliminates the need for high-pressure thick-walled gas cylinder, complex ignition components and special combustion chamber, greatly simplifying the system structure, significantly reducing the overall weight and manufacturing cost, while avoiding the safety risks and structural redundancy brought about by high-pressure gas cylinder.

[0037] The vaporizer 600 of the present invention is used to vaporize liquid nitrogen and cool the tube wall of the thrust chamber 500. Compared with the prior art which sets a cooling fuel flow channel in the nozzle, the structure of the vaporizer 600 achieves a lightweight design, which significantly reduces the overall weight and manufacturing cost of the thrust chamber 500 nozzle. Furthermore, the liquid nitrogen directly contacts the tube wall of the thrust chamber 500 for heat exchange, and its heat exchange efficiency is also improved. Example 2

[0038] The overall system structure and working principle of this embodiment are basically the same as those of Embodiment 1. The difference lies in the material selection and manufacturing process of the core structural components, which are further optimized for industrial mass production. In particular, the forming method of the carburetor 600 is specially designed to completely abandon high-cost processes such as additive manufacturing and precision integral milling. It relies entirely on general sheet metal and welding processes to achieve large-scale manufacturing, thus solving the problems of existing technologies being unable to mass-produce and high manufacturing costs from an engineering application perspective.

[0039] In this embodiment, core structural components such as the vaporizer 600, thrust chamber 500, oxidizer tank 300, fuel tank 400, and liquid nitrogen tank 100 are all made of common engineering materials such as 304 stainless steel or aluminum alloy. These materials are widely available and have excellent formability, eliminating the need for high-cost materials such as high-temperature alloys and special composite materials, thus significantly reducing raw material costs. Auxiliary components such as the liquid nitrogen turbopump, oxygen tank outlet control valve 301, fuel tank outlet control valve 401, nitrogen tank outlet control valve 101, exhaust gas control valve 701, and various booster control valves are all industrial-grade standardized products. Alternatively, suitable products can be directly purchased from a mature supply chain, eliminating the need for customized high-precision design and processing, further reducing R&D and manufacturing costs.

[0040] The vaporizer 600 is composed of multiple layers of annular grooved sleeves 603 stacked sequentially along the axial direction of the thrust chamber 500 and interconnected with each other. (See reference...) Figures 2-3 Each annular groove sleeve 603 is formed by welding an upper annular plate, an intermediate annular cylinder, and a lower annular plate using conventional argon arc welding. The upper and lower annular plates are made from flat sheet metal through shearing, stamping, or turning to form standard annular thin sheets, while the intermediate annular cylinder is made from rolled or extruded profiles. After assembly, the upper and lower annular plates are welded to the upper and lower end faces of the intermediate annular cylinder, respectively, to form a C-shaped groove cavity structure with one side opening. During assembly, the open side of the C-shaped groove cavity is tightly fitted to the outer wall of the thrust chamber 500 and sealed by welding, so that the annular groove sleeve 603 and the outer wall of the thrust chamber 500 together form a closed annular heat exchange channel 604. Adjacent annular groove sleeves 603 are fixed in structure and connected to the flow channel by welding.

[0041] The manufacturing process described above only requires shearing machines, punch presses, ordinary lathes, rolling machines, and conventional argon arc welding equipment. It eliminates the need for expensive equipment such as CNC precision machining centers, 3D printers, and specialized forging dies. Single-piece processing is time-efficient, material utilization is high, and process consistency is strong, enabling assembly line-style mass production. The overall system assembly process is simple, with no complex assembly tolerance requirements, and subsequent maintenance and component replacement are convenient. While meeting the high thrust output requirements of the first-stage rocket, it achieves a balance between high reliability and extremely low manufacturing costs.

[0042] 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 compression-type first-stage rocket capable of achieving high thrust output, characterized in that, include: Liquid nitrogen storage tank, liquid nitrogen turbopump, oxidizer storage tank, fuel storage tank, thrust chamber and vaporizer; The vaporizer is attached to the outer wall of the thrust chamber; The liquid nitrogen turbopump includes a liquid nitrogen pump and a turbine connected to the same drive shaft. The outlet of the liquid nitrogen tank is connected to the inlet of the liquid nitrogen pump through a nitrogen tank outlet control valve. The outlet of the liquid nitrogen pump is connected to the inlet of the vaporizer. The outlet of the vaporizer is connected to an oxygen tank pressurization control valve, a fuel tank pressurization control valve, and a nitrogen tank pressurization control valve. The other end of the oxygen tank pressurization control valve is connected to the pressurization diffuser inside the oxidant storage tank, and the other end of the fuel tank pressurization control valve is connected to the pressurization diffuser inside the fuel storage tank. Liquid nitrogen is pressurized by a liquid nitrogen pump and enters the vaporizer, where it absorbs residual heat from the thrust chamber wall and partially seeps into the thrust chamber wall through the sweat holes to form a thin film for cooling. After absorbing heat, it vaporizes into high-pressure nitrogen. A portion of the high-pressure nitrogen gas compresses the oxidizer tank and the fuel tank, forcing the oxidizer in the oxidizer tank and the fuel in the fuel tank into the thrust chamber for combustion; another portion of the high-pressure nitrogen gas compresses the liquid nitrogen surface in the liquid nitrogen tank, pushing the liquid nitrogen continuously into the liquid nitrogen pump, forming a closed loop. The turbine is driven by a small portion of high-temperature gas drawn directly from the thrust chamber, and its outlet is connected to the outside.

2. The compression-type first-stage rocket capable of achieving high thrust output according to claim 1, characterized in that, The vaporizer consists of an inlet pipe, an outlet pipe, and multiple layers of annular grooves. The multiple layers of annular grooves with C-shaped cross sections are interconnected with the inlet pipe and the outlet pipe. The C-shaped opening side of each annular groove is sealed and fitted to the outer wall of the thrust chamber, forming an annular heat exchange channel.

3. The compression-type first-stage rocket capable of achieving high thrust output according to claim 2, characterized in that, Any layer of annular groove sleeve is formed by welding the upper ring plate, the lower ring plate and the middle ring cylinder together by conventional argon arc welding, and the three together form a C-shaped groove cavity structure with one side open.

4. The compression-type first-stage rocket capable of achieving high thrust output according to claim 3, characterized in that, The inner wall of the thrust chamber is provided with multiple sweat holes, which are distributed circumferentially along the inner wall of the thrust chamber. The axis of the sweat holes forms an angle of 30 to 45 degrees with the inner wall of the thrust chamber and faces the outlet of the thrust chamber. Any one of the sweat holes penetrates the wall of the thrust chamber and connects the thrust chamber with the interior of the vaporizer.

5. The compression-type first-stage rocket capable of achieving high thrust output according to claim 1, characterized in that, An oxygen tank outlet control valve is installed in the pipeline between the outlet of the oxidizer tank and the oxygen injector in the thrust chamber, and a fuel tank outlet control valve is installed in the pipeline between the outlet of the fuel tank and the fuel injector in the thrust chamber.

6. The compression-type first-stage rocket capable of achieving high thrust output according to claim 1, characterized in that, A gas extraction control valve is installed on the gas ejector line from the thrust chamber to the turbine.

7. The compression-type first-stage rocket capable of achieving high thrust output according to claim 1, characterized in that, Check valves are provided at the outlets of the oxygen tank pressurization control valve, the fuel tank pressurization control valve, and the nitrogen tank pressurization control valve.

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