A gas generator cycle liquid rocket engine system with electrically powered bypass

Abstract

This invention relates to the field of aerospace propulsion technology and discloses a gas generator circulating liquid rocket engine system with an electric bypass. This invention constructs an electric bypass supply channel by reusing the discharge pipeline, introducing the originally directly discharged high-pressure propellant into the fuel electric pump and oxidizer electric pump for secondary pressurization, and then injecting it into the main thrust chamber via a separately set bypass injector. This solves the problems of traditional open gas generator circulating liquid engines, such as the main turbopump easily entering the critical speed range and causing mechanical resonance under deep throttling conditions, and the deterioration of atomization mixing effect due to insufficient pressure drop of the main injector. It achieves stable operation of the main turbopump within the rated operating speed range and stable high-pressure atomization and combustion of propellant under extremely low thrust conditions, expands the lower limit of the engine's deep throttling capability, and improves propellant energy utilization efficiency and system operational reliability. Simultaneously, the electric bypass supply system enables rapid response and precise adjustment under low thrust conditions.

Description

Technical Field

[0001] This invention relates to the field of aerospace propulsion technology, and in particular to a gas generator circulating liquid rocket engine system with electric bypass. Background Technology

[0002] In the field of aerospace engineering, the deep throttling capability of liquid rocket engines is one of the core technologies for achieving precise spacecraft control. It directly determines the maneuverability, safety, and mission success rate of spacecraft in complex mission scenarios, and is of great significance for the smooth buffering of vertical soft landings, precise attitude adjustment during space rendezvous and docking, and flexible orbit changes during atmospheric flight. However, traditional open gas generator cycle engines generally suffer from two major technical bottlenecks when implementing deep throttling, especially under extremely low operating conditions where thrust drops to less than 10% of the rated value, which severely restricts their application scope.

[0003] First, the turbopump's operational stability is insufficient. When engine thrust decreases significantly, the main turbopump's speed also drops significantly. At this time, the turbopump rotor is prone to entering the critical speed range, triggering severe mechanical resonance. Simultaneously, the decrease in speed disrupts the internal axial force balance of the turbopump, leading to mechanical failures such as bearing wear and seal failure. In severe cases, this can cause the turbopump to be scrapped, fundamentally affecting the overall reliability of the engine. Second, the injector's atomization and mixing effect is severely degraded. The quality of propellant atomization and mixing directly determines combustion efficiency and stability, and the atomization effect is closely related to the injector pressure drop. As engine thrust decreases, the propellant flow rate decreases simultaneously, and the injector pressure drop decreases accordingly. This results in larger propellant atomized particles and uneven mixing, which can easily lead to combustion oscillations, flameout, and other problems, making stable combustion under extremely low operating conditions impossible.

[0004] To address the aforementioned technical bottlenecks, existing technologies often employ complex mechanical variable-area injectors, such as needle-type injectors, to adapt to different flow rate requirements by altering the injector's flow area, thereby achieving deep throttling. However, this approach has significant drawbacks: its complex mechanical structure and extremely high machining precision requirements lead to prolonged development cycles and substantial cost increases. Furthermore, it only improves atomization conditions at the injector end and cannot fundamentally solve the problem of main turbopump stability at extremely low speeds, leaving the lower limit of throttling strictly limited. In addition, traditional engine systems typically have a post-pump discharge line, primarily used for engine pre-cooling circulation, emergency depressurization, or pre-start pipeline filling. Under normal operating conditions, the high-pressure propellant, pressurized by the main pump, is directly discharged to the outside or reinjected into the storage tank, resulting in both propellant waste and reduced system energy utilization efficiency.

[0005] Therefore, there is an urgent need to provide a technical solution to address the above problems. Summary of the Invention

[0006] To address the aforementioned technical problems, this invention provides a gas generator cycle liquid rocket engine system with an electric bypass. The technical solution of this gas generator cycle liquid rocket engine system with an electric bypass is as follows: Includes: main pump pressure supply circulation system, discharge system and electric bypass supply system; The main pump pressure supply cycle system includes a fuel main pump, an oxidizer main pump, a turbine, a gas generator, a main thrust chamber, and a main injector; the gas generator receives propellant from the fuel main pump and the oxidizer main pump and generates gas to drive the turbine, which in turn drives the fuel main pump and the oxidizer main pump; the main injector is located at the head of the main thrust chamber and injects propellant into the main thrust chamber. The discharge system includes a fuel discharge valve, an oxidant discharge valve, a fuel discharge pipeline, and an oxidant discharge pipeline; the fuel discharge valve is located at the outlet of the main fuel pump, the oxidant discharge valve is located at the outlet of the main oxidant pump, the fuel discharge pipeline is connected to the fuel discharge valve, and the oxidant discharge pipeline is connected to the oxidant discharge valve. The electric bypass supply system includes a fuel inlet, an oxidant inlet, a fuel electric pump, an oxidant electric pump, a bypass-specific injector, a fuel check valve, and an oxidant check valve. The fuel inlet is connected to the fuel discharge pipeline downstream of the fuel discharge valve, and the oxidant inlet is connected to the oxidant discharge pipeline downstream of the oxidant discharge valve. The inlet of the fuel electric pump is connected to the fuel liquid inlet, and the outlet of the fuel electric pump is connected to the bypass dedicated injector through the fuel check valve. The inlet of the electric oxidant pump is connected to the oxidant liquid outlet, and the outlet of the electric oxidant pump is connected to the bypass dedicated injector through the oxidant one-way valve. The bypass dedicated injector is located at the head of the main thrust chamber and is independent of the main injector. The fuel check valve is located downstream of the fuel electric pump and upstream of the bypass injector, and the oxidant check valve is located downstream of the oxidant electric pump and upstream of the bypass injector.

[0007] The beneficial effects of the gas generator circulating liquid rocket engine system with electric bypass of the present invention are as follows: The system of this invention constructs an electric bypass supply channel by reusing the discharge pipeline, introducing the high-pressure propellant, which was originally directly discharged, into the fuel electric pump and oxidizer electric pump for secondary pressurization, and then injecting it into the main thrust chamber through an independently set bypass injector. This solves the problems of mechanical resonance caused by the main turbopump easily entering the critical speed range under deep throttling conditions in traditional open gas generator circulating liquid engines, and the deterioration of atomization mixing effect due to insufficient pressure drop of the main injector. It realizes stable operation of the main turbopump within the rated operating speed range and stable combustion of high-pressure atomization of propellant under extremely low thrust conditions, expands the lower limit of the engine's deep throttling capability, and improves the propellant energy utilization efficiency and system operational reliability. At the same time, the electric bypass supply system enables rapid response and precise adjustment under micro-thrust conditions.

[0008] Based on the above scheme, the gas generator circulating liquid rocket engine system with electric bypass of the present invention can be further improved as follows.

[0009] In one alternative embodiment, when the electric bypass supply system is in operation, the fuel drain valve and the oxidizer drain valve are closed, and the propellant in the fuel drain line and the oxidizer drain line enters the fuel electric pump and the oxidizer electric pump respectively through the fuel inlet and the oxidizer inlet.

[0010] The beneficial effects of adopting the above-mentioned optional method are as follows: by further closing the fuel drain valve and the oxidizer drain valve, the channel for propellant to be discharged to the outside is blocked, so that the high-pressure propellant that originally flowed out through the drain pipeline accumulates in the pipeline and flows into the electric bypass supply system, thereby realizing the secondary pressurization and reuse of the residual propellant in the drain pipeline, and avoiding the energy loss and resource waste caused by the direct discharge of propellant.

[0011] In one alternative configuration, when the main pump pressure supply circulation system is operating, the electric bypass supply system is in standby mode, and the fuel check valve and the oxidizer check valve are closed under the pressure of the main thrust chamber.

[0012] The advantages of adopting the above-mentioned optional method are as follows: when the main pump-pressurized supply circulation system is working, the electric bypass supply system is put into standby mode. The pressure inside the main thrust chamber acts on the fuel check valve and the oxidizer check valve, so that the two check valves are kept closed under the action of pressure difference. This achieves physical isolation between the main pump-pressurized supply circulation system and the electric bypass supply system, prevents propellant from flowing back into the bypass pipeline under the main pipeline working state, and ensures that the two supply channels operate independently without interfering with each other.

[0013] In one alternative embodiment, both the fuel electric pump and the oxidizer electric pump are centrifugal pumps driven by high-speed motors, and the outlet pressure design values ​​of the fuel electric pump and the oxidizer electric pump are greater than the operating pressure value of the main thrust chamber under extremely low operating conditions.

[0014] The advantages of adopting the above-mentioned optional method are as follows: by further using a centrifugal pump driven by a high-speed motor to perform secondary pressurization on the propellant introduced from the discharge pipeline, and by designing the outlet pressure of the fuel electric pump and the oxidizer electric pump to be higher than the working pressure of the main thrust chamber under extremely low operating conditions, sufficient injection pressure drop is provided for the bypass supply system, solving the problem of atomization quality deterioration caused by insufficient injector pressure drop under extremely low thrust conditions, and ensuring the fineness of propellant atomized particles and the uniformity of mixing.

[0015] In one alternative embodiment, the bypass-specific injector includes a central atomizing nozzle and wall cooling nozzles facing the inner wall of the main thrust chamber.

[0016] The advantages of adopting the above-mentioned optional method are as follows: by setting a central atomizing nozzle and a wall cooling nozzle through a bypass dedicated injector, a portion of the propellant is directionally sprayed onto the inner wall surface of the main thrust chamber using the wall cooling nozzle, forming a continuous coolant film covering the thrust chamber wall surface, reducing the thermal impact and ablation effect of high combustion temperature on the chamber wall, while the central atomizing nozzle ensures good atomization mixing and flame stability in the combustion core area, achieving reliable thermal protection under extremely low operating conditions.

[0017] In one alternative embodiment, the fuel check valve is a differential pressure driven check valve, which opens when the pressure in the main thrust chamber is less than the outlet pressure of the fuel electric pump; the oxidant check valve is a differential pressure driven check valve, which opens when the pressure in the main thrust chamber is less than the outlet pressure of the oxidant electric pump.

[0018] The advantages of adopting the above-mentioned optional method are as follows: by further using differential pressure driven check valves as fuel check valves and oxidizer check valves, the fuel check valve opens when the pressure value of the main thrust chamber is less than the outlet pressure value of the fuel electric pump, and the oxidizer check valve opens when the pressure value of the main thrust chamber is less than the outlet pressure value of the oxidizer electric pump, thereby realizing automatic opening and closing control of the check valves without the need for additional electromagnetic control signals or external driving force input, simplifying the control logic and hardware configuration of the electric bypass supply system, and improving the system response speed and operational reliability.

[0019] In one alternative embodiment, the fuel check valve closes when the pressure in the main thrust chamber is greater than the outlet pressure of the fuel electric pump, and the oxidant check valve closes when the pressure in the main thrust chamber is greater than the outlet pressure of the oxidant electric pump.

[0020] The advantages of adopting the above-mentioned optional method are as follows: by further utilizing the reverse shut-off characteristic of the pressure differential driven check valve, the fuel check valve closes when the pressure value of the main thrust chamber is greater than the outlet pressure value of the fuel electric pump, and the oxidant check valve closes when the pressure value of the main thrust chamber is greater than the outlet pressure value of the oxidant electric pump. The pressure difference between the main thrust chamber and the outlet of the electric pump is used to realize the automatic closure of the check valve, cut off the bypass channel to prevent the high temperature and high pressure gas from flowing back towards the electric pump, and protect the electric pump and related pipelines from gas erosion damage.

[0021] In an alternative embodiment, an electronic controller and a chamber pressure sensor are also included. The chamber pressure sensor is located in the main thrust chamber and is used to collect the pressure signal of the main thrust chamber and transmit it to the electronic controller. The electronic controller is used to control the operation of the fuel discharge valve, the oxidizer discharge valve, the fuel electric pump, and the oxidizer electric pump according to the pressure signal.

[0022] The advantages of adopting the above-mentioned optional method are as follows: by further setting up an electronic controller and a chamber pressure sensor, the chamber pressure sensor collects the pressure change signal inside the main thrust chamber in real time and transmits it to the electronic controller. The electronic controller controls the operation of the fuel discharge valve, oxidant discharge valve, fuel electric pump and oxidant electric pump according to the pressure feedback signal, so as to realize the smooth switching and closed-loop control between the two working modes of main pump pressure supply and electric bypass supply, thereby improving the system's automation level and working condition adaptability.

[0023] In one alternative embodiment, upon receiving an extremely low operating condition command, the electronic controller partially or completely closes the main fuel valve and the main oxidant valve in the main pump pressure supply circulation system, and closes the fuel discharge valve and the oxidant discharge valve. When the chamber pressure sensor detects that the pressure value of the main thrust chamber is less than the outlet pressure value of the fuel electric pump and less than the outlet pressure value of the oxidant electric pump, the electronic controller starts the fuel electric pump and the oxidant electric pump.

[0024] The advantages of adopting the above-mentioned optional method are as follows: After receiving the extremely low operating condition command, the electronic controller will reduce or close the main fuel valve and the main oxidant valve, and close the fuel discharge valve and the oxidant discharge valve. When the chamber pressure sensor detects that the pressure value of the main thrust chamber is less than the outlet pressure value of the fuel electric pump and less than the outlet pressure value of the oxidant electric pump, the fuel electric pump and the oxidant electric pump will be started. The valve closing actions will be coordinated and controlled according to the preset timing to avoid motor overload or mechanical damage caused by starting the electric pump under high pressure differential conditions.

[0025] In one alternative embodiment, the electric bypass supply system is activated as an independent emergency power source when the main propulsion system fails. The fuel drain valve and the oxidizer drain valve are closed, and the electric fuel pump and the electric oxidizer pump are activated. Propellant enters the electric fuel pump and the electric oxidizer pump through the fuel inlet and the oxidizer inlet, and after being pressurized, it is injected into the main thrust chamber through the bypass-specific injector to generate micro-thrust.

[0026] The beneficial effects of adopting the above-mentioned optional method are as follows: In the event of a failure in the main propulsion system, the electric bypass supply system can be started as an independent emergency power source. By closing the fuel drain valve and the oxidizer drain valve and starting the fuel electric pump and the oxidizer electric pump, the high-pressure propellant in the drain pipeline enters the fuel electric pump and the oxidizer electric pump through the fuel liquid intake port and the oxidizer liquid intake port. After being pressurized, it is injected into the main thrust chamber through the bypass dedicated injector to generate micro-thrust, providing the spacecraft with the power required for emergency maneuvering, attitude adjustment or emergency landing, and improving the mission completion capability of the engine system under failure conditions.

[0027] The above description is merely an overview of the technical solution of the present invention. In order to better understand the technical means of the present invention and to implement it in accordance with the contents of the specification, and in order to make the above and other objects, features and advantages of the present invention more apparent and understandable, specific embodiments of the present invention are described below. Attached Figure Description

[0028] The accompanying drawings are for illustrative purposes only and are not intended to limit the invention. Furthermore, the same reference numerals denote the same parts throughout the drawings. In the drawings: Figure 1 This is a schematic diagram of an embodiment of a gas generator circulating liquid rocket engine system with electric bypass according to the present invention; Figure 2 This is a schematic diagram of the overall system architecture. Figure 3 This is a timing diagram showing the operating conditions and valve actions; Figure 4 This is a schematic diagram of the thrust chamber head structure; Figure 5 This diagram illustrates the advantages of the present invention compared to traditional needle-type injector solutions. Detailed Implementation

[0029] Exemplary embodiments of the invention will now be described in more detail with reference to the accompanying drawings. While exemplary embodiments of the invention are shown in the drawings, it should be understood that the invention can be implemented in various forms and should not be limited to the embodiments set forth herein.

[0030] Figure 1A schematic diagram of an embodiment of a gas generator circulating liquid rocket engine system 10 with an electric bypass provided by the present invention is shown. Figure 1 As shown, the gas generator circulating liquid rocket engine system 10 with electric bypass includes: a main pump-pressurized supply circulation system 11, a venting system 12, and an electric bypass supply system 13.

[0031] The main pump pressure supply circulation system 11 includes a fuel main pump, an oxidizer main pump, a turbine, a gas generator, a main thrust chamber, and a main injector; the gas generator is used to receive propellant from the fuel main pump and the oxidizer main pump and generate gas to drive the turbine, and the turbine is used to drive the fuel main pump and the oxidizer main pump; the main injector is disposed at the head of the main thrust chamber and is used to inject propellant into the main thrust chamber.

[0032] The main fuel pump, driven by a turbine, is a pump that draws fuel from the storage tank and pressurizes it to the required pressure level. For example, in a liquid oxygen-methane rocket engine with a rated thrust of 10 tons, the main fuel pump draws liquid methane from the storage tank, increases the outlet pressure to 12 MPa, and then delivers it to the gas generator and main injector. Similarly, the main oxidizer pump, driven by a turbine, draws oxidizer from the storage tank and pressurizes it to the required pressure level. For example, in a liquid oxygen-methane rocket engine with a rated thrust of 10 tons, the main oxidizer pump draws liquid oxygen from the storage tank, increases the outlet pressure to 12 MPa, and then delivers it to the gas generator and main injector.

[0033] In this context, a turbine refers to a rotating power unit that uses the expansion of high-temperature, high-pressure gas generated by a gas generator to convert thermal energy into mechanical energy. It drives the main fuel pump and the main oxidizer pump. For example, in a liquid oxygen-methane rocket engine with a rated thrust of 10 tons, the turbine rotates at tens of thousands of revolutions per minute driven by the high-temperature gas generated by the gas generator, synchronously driving the coaxially connected main fuel pump and main oxidizer pump. A gas generator, on the other hand, is a device that receives a portion of the propellant for combustion to generate high-temperature, high-pressure gas, which drives the turbine. For instance, in a liquid oxygen-methane rocket engine with a rated thrust of 10 tons, the gas generator receives a small portion of methane from the main fuel pump and a small portion of liquid oxygen from the main oxidizer pump. After combustion, it produces fuel-rich gas at approximately 600°C and a pressure of approximately 10 MPa, driving the turbine's rotation.

[0034] The main thrust chamber refers to the primary combustion chamber where propellant combustion generates thrust. For example, in a liquid oxygen-methane rocket engine with a rated thrust of 10 tons, the main thrust chamber has a bell-shaped structure with an internal diameter of approximately 300 mm. The walls employ a regenerative cooling channel design, and the propellant generates 10 tons of rated thrust after combustion within the chamber. The main injector is a device located at the head of the main thrust chamber, used to atomize and mix the propellant from the main system before injecting it into the main thrust chamber. For example, in a liquid oxygen-methane rocket engine with a rated thrust of 10 tons, the main injector adopts a coaxial centrifugal structure, with approximately 200 injection units distributed at the head of the thrust chamber. It atomizes methane and liquid oxygen into droplets with an average particle size of 50 μm before injecting them into the combustion chamber.

[0035] In rocket engines, propellant refers to the fuel and oxidizer that undergo a chemical reaction to produce high-temperature combustion gases. For example, in a liquid oxygen-methane rocket engine with a rated thrust of 10 tons, liquid oxygen is used as the oxidizer and liquid methane as the fuel, which are mixed in a mass ratio of 3.5:1 and then burned to generate thrust. Gas-driven operation refers to the method of using high-temperature, high-pressure combustion gases generated by a gas generator to drive a turbine. For example, in a liquid oxygen-methane rocket engine with a rated thrust of 10 tons, the high-temperature combustion gases generated by the gas generator impact the turbine blades at a speed of approximately 500 m / s, causing the turbine to generate approximately 800 kW of mechanical power to drive the main fuel pump and the main oxidizer pump.

[0036] The discharge system 12 includes a fuel discharge valve, an oxidant discharge valve, a fuel discharge pipeline, and an oxidant discharge pipeline; the fuel discharge valve is located at the outlet of the main fuel pump, the oxidant discharge valve is located at the outlet of the main oxidant pump, the fuel discharge pipeline is connected to the fuel discharge valve, and the oxidant discharge pipeline is connected to the oxidant discharge valve.

[0037] The fuel venting valve refers to a valve installed on the outlet pipeline of the main fuel pump to control the opening and closing of the fuel venting pipeline. For example, in a liquid oxygen-methane rocket engine with a rated thrust of 10 tons, the fuel venting valve is an electric ball valve installed at the outlet flange of the main fuel pump. It opens before engine start to pre-cool the pipeline and closes at extremely low operating conditions to trap high-pressure methane in the pipeline. The oxidizer venting valve refers to a valve installed on the outlet pipeline of the main oxidizer pump to control the opening and closing of the oxidizer venting pipeline. For example, in a liquid oxygen-methane rocket engine with a rated thrust of 10 tons, the oxidizer venting valve is an electric ball valve installed at the outlet flange of the main oxidizer pump. It opens before engine start to pre-cool the pipeline and closes at extremely low operating conditions to trap high-pressure liquid oxygen in the pipeline.

[0038] The fuel venting pipeline refers to a pipeline connected to the fuel venting valve for fuel pre-cooling circulation or emergency depressurization. For example, in a liquid oxygen-methane rocket engine with a rated thrust of 10 tons, the fuel venting pipeline is a stainless steel pipeline with an inner diameter of 15 mm, one end connected to the fuel venting valve, and the other end leading to the engine exhaust port. The pipeline is designed to withstand a pressure of 15 MPa. The oxidizer venting pipeline refers to a pipeline connected to the oxidizer venting valve for oxidizer pre-cooling circulation or emergency depressurization. For example, in a liquid oxygen-methane rocket engine with a rated thrust of 10 tons, the oxidizer venting pipeline is a stainless steel pipeline with an inner diameter of 15 mm, one end connected to the oxidizer venting valve, and the other end leading to the engine exhaust port. The pipeline is designed to withstand a pressure of 15 MPa.

[0039] The electric bypass supply system 13 includes a fuel inlet, an oxidant inlet, a fuel electric pump, an oxidant electric pump, a bypass-specific injector, a fuel check valve, and an oxidant check valve.

[0040] The fuel intake port refers to the interface from the fuel discharge line where propellant is led out to the fuel electric pump. For example, in a liquid oxygen-methane rocket engine with a rated thrust of 10 tons, the fuel intake port is located approximately 200 mm downstream of the fuel discharge valve on the discharge line, and high-pressure methane in the line is introduced into the inlet of the fuel electric pump through a high-pressure hose with an inner diameter of 8 mm. The oxidizer intake port refers to the interface from the oxidizer discharge line where propellant is led out to the oxidizer electric pump. For example, in a liquid oxygen-methane rocket engine with a rated thrust of 10 tons, the oxidizer intake port is located approximately 200 mm downstream of the oxidizer discharge valve on the discharge line, and high-pressure liquid oxygen in the line is introduced into the inlet of the oxidizer electric pump through a high-pressure hose with an inner diameter of 8 mm.

[0041] The electric fuel pump refers to a motor-driven pump used to further pressurize fuel and deliver it to a dedicated bypass injector. For example, in a liquid oxygen-methane rocket engine with a rated thrust of 10 tons, the electric fuel pump is a high-speed motor-driven centrifugal pump with a rated power of 5 kW, a speed adjustment range of 10,000 to 50,000 revolutions per minute, an inlet pressure of 12 MPa, and an outlet pressure increased to 2 MPa before being delivered to the dedicated bypass injector. The electric oxidizer pump refers to a motor-driven pump used to further pressurize oxidizer and deliver it to a dedicated bypass injector. For example, in a liquid oxygen-methane rocket engine with a rated thrust of 10 tons, the electric oxidizer pump is a high-speed motor-driven centrifugal pump with a rated power of 5 kW, a speed adjustment range of 10,000 to 50,000 revolutions per minute, an inlet pressure of 12 MPa, and an outlet pressure increased to 2 MPa before being delivered to the dedicated bypass injector.

[0042] The bypass injector refers to a spraying device located at the head of the main thrust chamber, independent of the main injector, and specifically designed for propellant atomization and mixing under extremely low operating conditions. For example, in a liquid oxygen-methane rocket engine with a rated thrust of 10 tons, the bypass injector adopts a ring-shaped layout surrounding the main injector, equipped with a central atomizing nozzle and wall cooling nozzles facing the inner wall of the combustion chamber, ensuring uniform propellant atomization even when the engine thrust drops to 50 kg. The fuel check valve refers to a valve located downstream of the fuel electric pump to prevent fuel or gas from flowing back from the thrust chamber to the fuel electric pump. For example, in a liquid oxygen-methane rocket engine with a rated thrust of 10 tons, the fuel check valve is a pressure differential driven check valve installed on the fuel electric pump outlet pipeline. It automatically opens when the thrust chamber pressure is lower than the fuel electric pump outlet pressure and automatically closes when the thrust chamber pressure is higher than the fuel electric pump outlet pressure. An oxidizer check valve is a valve located downstream of the oxidizer electric pump to prevent oxidizer or combustion gas from flowing back from the thrust chamber to the oxidizer electric pump. For example, in a liquid oxygen-methane rocket engine with a rated thrust of 10 tons, the oxidizer check valve is a differential pressure driven check valve installed on the outlet pipeline of the oxidizer electric pump. It automatically opens when the thrust chamber pressure is lower than the outlet pressure of the oxidizer electric pump and automatically closes when the thrust chamber pressure is higher than the outlet pressure of the oxidizer electric pump.

[0043] The fuel inlet is connected to the fuel discharge pipeline downstream of the fuel discharge valve, and the oxidant inlet is connected to the oxidant discharge pipeline downstream of the oxidant discharge valve.

[0044] The inlet of the fuel electric pump is connected to the fuel liquid outlet, and the outlet of the fuel electric pump is connected to the bypass dedicated injector through the fuel check valve.

[0045] The inlet of the electric oxidant pump is connected to the oxidant liquid outlet, and the outlet of the electric oxidant pump is connected to the bypass dedicated injector through the oxidant one-way valve.

[0046] The bypass-specific injector is located at the head of the main thrust chamber and is independent of the main injector.

[0047] The fuel check valve is located downstream of the fuel electric pump and upstream of the bypass injector, and the oxidant check valve is located downstream of the oxidant electric pump and upstream of the bypass injector.

[0048] The technical solution of this embodiment constructs an electric bypass supply channel by reusing the discharge pipeline. The high-pressure propellant, which was originally directly discharged, is introduced into the fuel electric pump and oxidizer electric pump for secondary pressurization, and then injected into the main thrust chamber through an independently set bypass injector. This solves the problems of mechanical resonance caused by the main turbopump easily entering the critical speed range under deep throttling conditions in traditional open gas generator circulating liquid engines, and the deterioration of atomization mixing effect due to insufficient pressure drop of the main injector. It realizes stable operation of the main turbopump in the rated operating speed range and stable combustion of high-pressure atomization of propellant under extremely low thrust conditions. It expands the lower limit of the engine's deep throttling capability and improves the propellant energy utilization efficiency and system operation reliability. At the same time, the electric bypass supply system realizes rapid response and precise adjustment under micro-thrust conditions.

[0049] In one alternative embodiment, when the electric bypass supply system 13 is in operation, the fuel drain valve and the oxidizer drain valve are closed, and the propellant in the fuel drain line and the oxidizer drain line enters the fuel electric pump and the oxidizer electric pump respectively through the fuel inlet and the oxidizer inlet.

[0050] In the above-mentioned optional methods, by further closing the fuel drain valve and the oxidizer drain valve, the channel for propellant to be discharged to the outside is blocked, so that the high-pressure propellant that originally flowed out through the drain pipeline accumulates in the pipeline and flows into the electric bypass supply system 13, thereby realizing the secondary pressurization and reuse of the residual propellant in the drain pipeline, and avoiding the energy loss and resource waste caused by the direct discharge of propellant.

[0051] In one alternative, when the main pump pressure supply circulation system 11 is operating, the electric bypass supply system 13 is in standby mode, and the fuel check valve and the oxidizer check valve are closed under the pressure of the main thrust chamber.

[0052] The standby state refers to the working mode in which the electric bypass supply system 13 is in a ready-to-work state but not in operation; for example, in a liquid oxygen methane rocket engine with a rated thrust of 10 tons, when the engine is running at a rated thrust of 10 tons, neither the fuel electric pump nor the oxidizer electric pump is started, the bypass shut-off valve is in the closed position, and the fuel check valve and the oxidizer check valve remain closed under the high pressure of the thrust chamber, and the electric bypass supply system 13 is in a standby state.

[0053] In the above-mentioned optional method, when the main pump pressure supply circulation system 11 is working, the electric bypass supply system 13 is put into standby mode. The pressure inside the main thrust chamber acts on the fuel check valve and the oxidizer check valve, so that the two check valves are kept closed under the action of pressure difference. This achieves physical isolation between the main pump pressure supply circulation system 11 and the electric bypass supply system 13, prevents propellant from flowing back into the bypass pipeline under the main pipeline working state, and ensures that the two supply channels operate independently without interfering with each other.

[0054] In one alternative embodiment, both the fuel electric pump and the oxidizer electric pump are centrifugal pumps driven by high-speed motors, and the outlet pressure design values ​​of the fuel electric pump and the oxidizer electric pump are greater than the operating pressure value of the main thrust chamber under extremely low operating conditions.

[0055] Centrifugal pumps refer to pumps that use the centrifugal force generated by the rotation of an impeller to draw propellant in from the inlet and throw it out from the outlet to achieve pressurization. For example, in a liquid oxygen methane rocket engine with a rated thrust of 10 tons, both the fuel electric pump and the oxidizer electric pump adopt a centrifugal pump structure driven by a high-speed motor. The impeller diameter is about 50 mm. At a speed of 50,000 revolutions per minute, the propellant is drawn in from the inlet and thrown out by centrifugal force to achieve a pressure increase of about 2 MPa.

[0056] The outlet pressure design value refers to the pressure that the electric pump can stably provide at the outlet under rated operating conditions. For example, in a liquid oxygen methane rocket engine with a rated thrust of 10 tons, the outlet pressure design value of both the fuel electric pump and the oxidizer electric pump is 2MPa. This value is determined based on the thrust chamber working pressure of approximately 1.5MPa under extremely low operating conditions and after considering pipeline pressure loss, to ensure that the fuel check valve and the oxidizer check valve can open normally.

[0057] Extremely low operating conditions refer to the operating state in which the engine thrust drops to less than 10% of the rated thrust. For example, in a liquid oxygen methane rocket engine with a rated thrust of 10 tons, when the engine thrust drops to less than 1 ton, it is considered to have entered extremely low operating conditions, and can achieve stable operation at a minimum of 0.5% of the rated thrust, or 50 kg.

[0058] In the above-mentioned optional methods, a centrifugal pump driven by a high-speed motor is further used to perform secondary pressurization on the propellant introduced from the discharge pipeline. By designing the outlet pressure of the fuel electric pump and the oxidizer electric pump to be higher than the working pressure of the main thrust chamber under extremely low operating conditions, sufficient injection pressure drop is provided for the bypass supply system, solving the problem of atomization quality deterioration caused by insufficient injector pressure drop under extremely low thrust conditions, and ensuring the fineness of propellant atomized particles and the uniformity of mixing.

[0059] In one alternative embodiment, the bypass-specific injector includes a central atomizing nozzle and wall cooling nozzles facing the inner wall of the main thrust chamber.

[0060] The central atomizing nozzle refers to a nozzle located at the center of the bypass injector, used to atomize the propellant into fine droplets. For example, in a liquid oxygen-methane rocket engine with a rated thrust of 10 tons, the bypass injector has multiple central atomizing nozzles with a diameter of 0.3 mm, which spray the propellant at a speed of approximately 20 m / s and break it into droplets with an average particle size of 30 μm, ensuring full contact with the flame inside the combustion chamber. The wall cooling nozzle refers to a nozzle located at the edge of the bypass injector, spraying propellant towards the inner wall of the main thrust chamber to form a coolant film. For example, in a liquid oxygen-methane rocket engine with a rated thrust of 10 tons, 24 wall cooling nozzles are evenly arranged along the outer edge of the bypass injector, each with a diameter of 0.2 mm, spraying a portion of the fuel onto the inner wall surface of the combustion chamber to form a continuous liquid film with a thickness of approximately 0.1 mm, protecting the wall surface from high-temperature erosion.

[0061] In the above-mentioned optional methods, a central atomizing nozzle and a wall cooling nozzle are further set by a bypass dedicated injector. The wall cooling nozzle is used to directionally spray part of the propellant onto the inner wall surface of the main thrust chamber, forming a continuous coolant film covering the thrust chamber wall surface. This reduces the heat flow impact and ablation effect of the high combustion temperature on the chamber wall. At the same time, the central atomizing nozzle ensures that the combustion core area maintains good atomization mixing and flame stability, achieving reliable thermal protection under extremely low operating conditions.

[0062] In one alternative embodiment, the fuel check valve is a differential pressure driven check valve, which opens when the pressure in the main thrust chamber is less than the outlet pressure of the fuel electric pump; the oxidant check valve is a differential pressure driven check valve, which opens when the pressure in the main thrust chamber is less than the outlet pressure of the oxidant electric pump.

[0063] Among them, the differential pressure driven check valve refers to a check valve that automatically opens and closes by utilizing the pressure difference across the valve body. For example, in a liquid oxygen methane rocket engine with a rated thrust of 10 tons, the differential pressure driven check valve adopts a spring preload structure. When the pressure difference between the pressure before and after the valve reaches 0.2 MPa, it opens by overcoming the spring force and automatically closes when the pressure after the valve is higher than the pressure before the valve, without the need for an external control signal.

[0064] In the above-mentioned optional methods, a differential pressure driven check valve is further used as both the fuel check valve and the oxidizer check valve. The fuel check valve opens when the pressure in the main thrust chamber is less than the outlet pressure of the fuel electric pump, and the oxidizer check valve opens when the pressure in the main thrust chamber is less than the outlet pressure of the oxidizer electric pump. This achieves automatic opening and closing control of the check valves without the need for additional electromagnetic control signals or external driving force input, simplifying the control logic and hardware configuration of the electric bypass supply system 13 and improving the system response speed and operational reliability.

[0065] In one alternative embodiment, the fuel check valve closes when the pressure in the main thrust chamber is greater than the outlet pressure of the fuel electric pump, and the oxidant check valve closes when the pressure in the main thrust chamber is greater than the outlet pressure of the oxidant electric pump.

[0066] In the above-mentioned optional methods, the reverse shut-off characteristic of the pressure differential driven check valve is further adopted, so that the fuel check valve closes when the pressure value of the main thrust chamber is greater than the pressure value of the fuel electric pump outlet, and the oxidant check valve closes when the pressure value of the main thrust chamber is greater than the pressure value of the oxidant electric pump outlet. The pressure difference between the main thrust chamber and the outlet of the electric pump is used to realize the automatic closure of the check valve, cut off the bypass channel to prevent the high temperature and high pressure gas from flowing back towards the electric pump, and protect the electric pump and related pipelines from gas erosion damage.

[0067] In an alternative embodiment, an electronic controller and a chamber pressure sensor are also included. The chamber pressure sensor is located in the main thrust chamber and is used to collect the pressure signal of the main thrust chamber and transmit it to the electronic controller. The electronic controller is used to control the operation of the fuel discharge valve, the oxidizer discharge valve, the fuel electric pump, and the oxidizer electric pump according to the pressure signal.

[0068] The electronic controller refers to an electronic control unit that receives sensor signals and outputs control commands according to preset logic. For example, in a liquid oxygen-methane rocket engine with a rated thrust of 10 tons, the electronic controller uses a redundant embedded computer to receive pressure signals from the chamber pressure sensor in real time and automatically control the switching and speed regulation of the fuel venting valve, oxidizer venting valve, fuel electric pump, and oxidizer electric pump according to mission requirements. The chamber pressure sensor is a sensor installed on the wall of the main thrust chamber to detect the internal gas pressure in real time. For example, in a liquid oxygen-methane rocket engine with a rated thrust of 10 tons, the chamber pressure sensor is a piezoresistive pressure sensor installed on the side wall of the main thrust chamber 50 mm from the injector panel, with a measurement range of 0 to 10 MPa, outputting a 4 to 20 mA current signal to the electronic controller.

[0069] In the above-mentioned optional methods, an electronic controller and a chamber pressure sensor are further set up. The chamber pressure sensor collects the pressure change signal inside the main thrust chamber in real time and transmits it to the electronic controller. The electronic controller controls the operation of the fuel discharge valve, oxidant discharge valve, fuel electric pump and oxidant electric pump according to the pressure feedback signal, so as to realize the smooth switching and closed-loop control between the two working modes of main pump pressure supply and electric bypass supply, thereby improving the system's automation level and working condition adaptability.

[0070] In one alternative embodiment, upon receiving an extremely low operating condition command, the electronic controller reduces or closes the main fuel valve and the main oxidant valve in the main pump pressure supply circulation system 11, and closes the fuel discharge valve and the oxidant discharge valve. When the chamber pressure sensor detects that the pressure value of the main thrust chamber is less than the outlet pressure value of the fuel electric pump and less than the outlet pressure value of the oxidant electric pump, the electronic controller starts the fuel electric pump and the oxidant electric pump.

[0071] Among them, the extremely low operating condition command refers to the control command issued by the spacecraft mission system, which requires the engine to enter the extremely low thrust operation mode. For example, in a liquid oxygen methane rocket engine with a rated thrust of 10 tons, when the spacecraft needs to further reduce the thrust from 1 ton to 50 kg in the final stage of vertical soft landing, the spacecraft flight control computer sends an extremely low operating condition command to the electronic controller, triggering the start-up procedure of the electric bypass supply system 13.

[0072] The main fuel valve refers to a valve located downstream of the main fuel pump, used to control the on / off state and flow rate of fuel entering the main injector. For example, in a liquid oxygen-methane rocket engine with a rated thrust of 10 tons, the main fuel valve is an electrically controlled regulating valve installed between the outlet of the main fuel pump and the inlet of the main injector. Under extremely low operating conditions, it is gradually closed by an electronic controller, reducing the methane flow rate from 3 kg / s under rated conditions to below 0.15 kg / s. The main oxidizer valve refers to a valve located downstream of the main oxidizer pump, used to control the on / off state and flow rate of oxidizer entering the main injector. For example, in a liquid oxygen-methane rocket engine with a rated thrust of 10 tons, the main oxidizer valve is an electrically controlled regulating valve installed between the outlet of the main oxidizer pump and the inlet of the main injector. Under extremely low operating conditions, it is gradually closed by an electronic controller, reducing the liquid oxygen flow rate from 1 kg / s under rated conditions to below 0.05 kg / s.

[0073] In the above-mentioned optional methods, after receiving the extremely low operating condition command, the electronic controller further closes or closes the main fuel valve and the main oxidizer valve, and closes the fuel discharge valve and the oxidizer discharge valve. When the chamber pressure sensor detects that the pressure value of the main thrust chamber is less than the outlet pressure value of the fuel electric pump and less than the outlet pressure value of the oxidizer electric pump, the fuel electric pump and the oxidizer electric pump are started. The valve closing actions are coordinated and controlled according to the preset timing sequence to avoid motor overload or mechanical damage caused by starting the electric pump under high pressure differential conditions.

[0074] In one alternative embodiment, the electric bypass supply system is activated as an independent emergency power source when the main propulsion system fails. The fuel drain valve and the oxidizer drain valve are closed, and the electric fuel pump and the electric oxidizer pump are activated. Propellant enters the electric fuel pump and the electric oxidizer pump through the fuel inlet and the oxidizer inlet, and after being pressurized, it is injected into the main thrust chamber through the bypass-specific injector to generate micro-thrust.

[0075] The main propulsion system refers to the main thrust supply unit that uses a gas generator cycle to generate rated thrust and high and medium-load thrust. For example, in a liquid oxygen-methane rocket engine with a rated thrust of 10 tons, the main propulsion system includes a main fuel pump, a main oxidizer pump, a turbine, a gas generator, a main thrust chamber, a main injector, a main fuel valve, and a main oxidizer valve. During normal operation, the main fuel valve and the main oxidizer valve are fully open. The main turbine pump operates at high speed under the drive of the high-temperature gas generated by the gas generator, which draws in liquid oxygen and liquid methane, pressurizes them, and delivers them to the main injector. After atomization and mixing, the mixture is injected into the main thrust chamber for combustion, generating a rated thrust of 10 tons.

[0076] Independent emergency power sources refer to backup power units that can provide thrust to spacecraft when the main propulsion system fails and cannot operate normally. For example, in a liquid oxygen-methane rocket engine with a rated thrust of 10 tons, when the main propulsion system fails and cannot continue to operate, the electric bypass supply system is activated as an independent emergency power source. The fuel discharge valve and oxidizer discharge valve are closed, and the fuel electric pump and oxidizer electric pump are activated. The high-pressure methane and high-pressure liquid oxygen trapped in the discharge pipeline are pressurized to 2MPa and then injected into the main thrust chamber through a bypass dedicated injector for combustion, generating 50kg of micro-thrust to provide the power required for emergency maneuvering, attitude adjustment, or emergency landing of the spacecraft.

[0077] Micro-thrust refers to the small thrust output of an engine when its thrust drops to less than 10% of its rated thrust, typically used for precise control or emergency maneuvers of spacecraft. For example, in a liquid oxygen-methane rocket engine with a rated thrust of 10 tons, when the electric bypass supply system operates alone, the propellant flow rate is precisely controlled at about 0.5% of the rated flow rate by adjusting the speed of the electric fuel pump and the electric oxidizer pump. The pressure in the main thrust chamber is maintained at 1.5 MPa, and the engine outputs 50 kg of thrust. This thrust value is far below 10% of the rated thrust and is used for precise speed control in the final stage of a spacecraft's vertical soft landing or for emergency maneuvers in case of a main propulsion system failure.

[0078] In the aforementioned optional methods, in the event of a failure in the main propulsion system, the electric bypass supply system can be activated as an independent emergency power source. By closing the fuel drain valve and oxidizer drain valve and starting the fuel electric pump and oxidizer electric pump, the high-pressure propellant in the drain pipeline enters the fuel electric pump and oxidizer electric pump through the fuel intake port and oxidizer intake port. After being pressurized, it is injected into the main thrust chamber through the bypass dedicated injector to generate micro-thrust, providing the spacecraft with the power required for emergency maneuvering, attitude adjustment or emergency landing, and improving the mission completion capability of the engine system under failure conditions.

[0079] To achieve stable thrust regulation across the entire operating range of the engine, this embodiment employs a master-slave dual-supply architecture. The master system is responsible for high thrust output, while the slave system is responsible for low thrust output. Without disrupting the original structure and efficiency of the traditional open-loop gas generator cycle, an independent electric bypass supply system 13 is connected in parallel to the existing high-pressure discharge pipeline after the engine's main pump. Stable thrust regulation across the entire operating range is achieved through the coordinated operation of these two systems. The system in this embodiment mainly consists of three parts: the main propulsion system, the discharge system 12, and the electric bypass supply system 13. These parts work together to ensure stable system operation.

[0080] like Figure 2 As shown, the main propulsion system is a gas generator cycle, serving as the engine's main thrust supply unit and responsible for stable operation under high and medium operating conditions. Its structure is consistent with traditional open-loop gas generator cycle systems. The main propulsion system includes a main fuel pump, a main oxidizer pump, a turbine, a gas generator, a main thrust chamber, a main injector, a main fuel valve, and a main oxidizer valve. Both the main fuel pump and the main oxidizer pump are driven by the turbine and are used to pressurize the propellant in the tank to the pressure level required for combustion. The gas generator receives a portion of the propellant pressurized by the main fuel pump and the main oxidizer pump, and generates fuel-rich or oxygen-rich gas through combustion to provide driving power for the turbine. The main injector atomizes and mixes the main system propellant before injecting it into the main thrust chamber, where it generates rated thrust through combustion. The main fuel valve and the main oxidizer valve control the on / off state and flow rate of the main system propellant, enabling coarse adjustment of the main system thrust.

[0081] The venting system 12, as a core reuse structure, provides a high-pressure propellant source for the bypass system, and its original functions remain unchanged. The venting system 12 includes a fuel venting valve, an oxidizer venting valve, a fuel venting line, and an oxidizer venting line. The fuel venting valve is located at the outlet of the main fuel pump, and the oxidizer venting valve is located at the outlet of the main oxidizer pump. The fuel venting valve and the oxidizer venting valve are used to control the opening and closing of the fuel venting line and the oxidizer venting line, respectively. The fuel venting line and the oxidizer venting line are used for pre-cooling circulation before engine start-up, emergency depressurization during operation, or line filling, and the propellant in the lines is always kept at high pressure.

[0082] The electric bypass supply system 13, as a core innovation, serves as the thrust supply unit for the engine under extremely low operating conditions. It specifically addresses the issues of injection atomization and supply stability in low-flow scenarios. Its compact design ensures physical isolation from the main system. The electric bypass supply system 13 includes a fuel inlet, an oxidizer inlet, an electric fuel pump, an electric oxidizer pump, a dedicated bypass injector, a fuel check valve, an oxidizer check valve, and a control valve assembly. The fuel inlet is located on the fuel discharge line downstream of the fuel discharge valve, and the oxidizer inlet is located on the oxidizer discharge line downstream of the oxidizer discharge valve. The fuel electric pump and oxidizer electric pump are two independent, low-power, high-speed electric pumps. The fuel electric pump inlet is connected to the fuel inlet via a pipeline, and the oxidizer electric pump inlet is connected to the oxidizer inlet via a pipeline. The outlets of the fuel electric pump and oxidizer electric pump are respectively connected to the thrust chamber to further pressurize the high-pressure propellant in the discharge pipeline, ensuring that the propellant pressure meets the combustion requirements under extremely low operating conditions. Simultaneously, the propellant flow rate is precisely controlled by adjusting the electric pump speed. A dedicated bypass injector is located at the head of the main thrust chamber, independent of the main injector. It is specifically designed for atomizing and mixing extremely low-flow-rate propellant, with optimized nozzle size and layout to ensure thorough atomization and mixing of the propellant at micro-flow rates. A fuel check valve is located downstream of the fuel electric pump and upstream of the dedicated bypass injector, and an oxidizer check valve is located downstream of the oxidizer electric pump and upstream of the dedicated bypass injector. This provides physical isolation between the main system and the bypass system, preventing high-pressure combustion gases or propellant in the thrust chamber from flowing back into the bypass system during main system operation. The control valve assembly includes a bypass shut-off valve and a flow regulating valve. The bypass shut-off valve is used to control the on / off of the bypass system, and the flow regulating valve is used to assist in regulating the propellant flow rate. Together with the electric pump speed regulation, it achieves precise thrust control under extremely low operating conditions.

[0083] It should be noted that the fluid logic and control principle in this embodiment is as follows: In the fuel circuit, after being pressurized by the main fuel pump, part of the fuel is delivered to the main injector via the main fuel valve, and the other part enters the fuel discharge pipeline via the fuel discharge valve. A fuel inlet is located downstream of the fuel discharge pipeline, connected to the inlet of the electric fuel pump. The outlet of the electric fuel pump is connected to the bypass injector via a fuel check valve. In the oxidant circuit, after being pressurized by the main oxidant pump, part of the oxidant is delivered to the main injector via the main oxidant valve, and the other part enters the oxidant discharge pipeline via the oxidant discharge valve. An oxidant inlet is located downstream of the oxidant discharge pipeline, connected to the inlet of the electric oxidant pump. The outlet of the electric oxidant pump is connected to the bypass injector via an oxidant check valve. The signal connection relationships between the control system and each actuator are clearly defined. The electronic controller receives the main thrust chamber pressure signal collected by the chamber pressure sensor and controls the actions of the fuel discharge valve, oxidant discharge valve, main fuel valve, main oxidant valve, electric fuel pump, and electric oxidant pump based on the pressure signal, realizing the fluid control logic and collaborative working mechanism between the main system and the bypass system.

[0084] This embodiment achieves a seamless transition of the engine from high and medium operating conditions to extremely low operating conditions through the coordinated switching of the main system and the bypass system. The operating logic of the two systems is independent and does not interfere with each other. Figure 3 As shown, the main system operates under high and medium operating conditions. When the engine requires rated thrust or higher thrust, the main propulsion system is activated, the main fuel valve and main oxidizer valve are fully opened, and the main turbopump operates at high speed driven by the high-temperature, high-pressure gas generated by the gas generator, always maintaining operation within the rated operating speed range and avoiding the critical speed area. It draws in the propellant from the fuel tank and oxidizer tank and pressurizes it to the set pressure, then delivers it to the main injector. After being atomized and mixed by the main injector, the propellant is injected into the main thrust chamber for vigorous combustion, generating the required thrust. At this time, the electric bypass supply system 13 is in standby mode, the electric fuel pump and the electric oxidizer pump are not working, and the bypass shut-off valve is closed. Because the pressure in the thrust chamber is much higher than the outlet pressure of the electric pump when the main system is working, the fuel check valve and the oxidizer check valve located upstream of the bypass dedicated injector are in a tightly closed state under the action of the pressure difference, realizing the physical isolation between the main system and the bypass system. Meanwhile, the discharge valve selectively opens according to the engine's operating requirements, and the propellant is discharged to the outside or reinjected into the storage tank through the fuel discharge pipeline and the oxidizer discharge pipeline to maintain the stable operation of the main system.

[0085] In extremely low operating conditions, the bypass system operates. When a space mission requires the engine to enter extremely low operating conditions, the bypass system is activated. The main fuel valve and main oxidizer valve are partially or completely closed, gradually reducing the propellant supply to the main system. The thrust generated by the main system decreases accordingly, and the pressure in the thrust chamber gradually decreases simultaneously. The fuel drain valve and oxidizer drain valve are closed, trapping the high-pressure propellant that would normally be discharged through the fuel drain line and oxidizer drain line within the drain line. The electric fuel pump and electric oxidizer pump are activated, and the bypass shut-off valve is opened. The high-pressure propellant trapped in the drain line is drawn out from the fuel intake port downstream of the fuel drain valve and the oxidizer intake port downstream of the oxidizer drain valve. After being further pressurized by the electric fuel pump and electric oxidizer pump to the pressure required for combustion under extremely low operating conditions, it is delivered to the dedicated bypass injector. When the thrust chamber pressure drops below the outlet pressure of both the fuel electric pump and the oxidizer electric pump, the fuel check valve and the oxidizer check valve automatically open due to the pressure difference. The propellant, pressurized by the electric pump, is then continuously injected into the thrust chamber through a dedicated bypass injector. The residual flame in the thrust chamber during the closure of the main fuel valve and main oxidizer valve ignites the propellant injected by the dedicated bypass injector, achieving a continuous and uninterrupted combustion transition from the main system to the bypass system. By adjusting the speeds of the fuel electric pump and the oxidizer electric pump using an electronic controller, the propellant flow rate is precisely controlled, thereby achieving stable thrust regulation under extremely low operating conditions.

[0086] When exiting the extremely low operating mode, and needing to return to high or medium operating modes, the above steps are executed in reverse. The main fuel valve and main oxidizer valve are slowly opened to start the main system, gradually increasing the propellant supply, causing the thrust chamber pressure to rise. Once the thrust chamber pressure rises above the outlet pressure of both the fuel electric pump and the oxidizer electric pump, the fuel check valve and oxidizer check valve automatically close under the pressure difference, achieving a re-isolation between the main system and the bypass system. Subsequently, the fuel electric pump, oxidizer electric pump, and bypass shut-off valve are closed, and the fuel drain valve and oxidizer drain valve are opened, restoring the drain system 12 to standby mode. The main system continues to provide thrust to the engine, completing the operating mode switch.

[0087] like Figure 4 As shown, the specific arrangement of the bypass injector in the thrust chamber head structure is clearly illustrated. The main injector is located in the central region of the thrust chamber head, while the bypass injector is located at the edge of the thrust chamber head and arranged around the main injector. The bypass injector has a central atomizing nozzle and a wall cooling nozzle, with the wall cooling nozzle facing the inner wall of the main thrust chamber to inject a portion of the bypass fuel onto the wall to form a coolant film. The main injector and the bypass injector are independent of each other, with the bypass injector serving both atomization and cooling functions.

[0088] like Figure 5As shown, the technical solution of this embodiment contrasts with the traditional needle-type injector solution. This embodiment employs a dual-path parallel architecture, physically separating the supply systems for high-thrust and low-thrust conditions, achieving decoupled control of thrust regulation. The main system focuses on the high-efficiency high-thrust range, ensuring the engine's efficiency and reliability under high and medium operating conditions. The bypass system is specifically optimized for low-flow scenarios, solving the problems of injection atomization and supply stability under low-flow conditions, allowing the engine's total thrust to be as low as 1%, breaking through the throttling limit of traditional engines. The bypass system's intake port is located downstream of the main pump's discharge valve. The propellant entering the electric pump is already pressurized by the main pump and is under high pressure, completely eliminating the risk of cavitation at the electric pump inlet, significantly reducing the requirements for the electric pump's suction performance, and simplifying the electric pump's structural design. The existing post-pump drain valve and drain line of the engine are directly reused as the propellant source for the bypass system, eliminating the need for additional independent high-pressure liquid intake lines and interfaces. This efficiently solves the liquid source problem for the bypass system, recovers and reuses the high-pressure propellant that was originally directly discharged, improves propellant utilization and system energy efficiency, and makes the engine system structure more compact. The pressure difference characteristics of the one-way valve enable automatic switching between the main system and the bypass system, eliminating the need for complex control timing and linkage logic, simplifying control system design, reducing control difficulty, and improving the reliability and stability of the transition process. The main injector can use a traditional coaxial or centrifugal injection unit, eliminating the need for a complex, high-precision needle-type variable area mechanism, significantly reducing the development difficulty, processing cost, and cycle time of the main combustion chamber, while improving the operational reliability of the main injector. During bypass system operation, there is no need for a separate low-condition igniter. The bypass propellant is ignited by the residual flame in the thrust chamber during the main system shutdown, achieving uninterrupted combustion transition, increasing system ignition redundancy, and improving the engine's starting reliability under extremely low operating conditions. In an emergency, if the main propulsion system fails to function properly, the bypass system can serve as an independent emergency power source to provide micro-thrust to the spacecraft, enabling emergency maneuvers, attitude adjustments, or emergency landings, thereby significantly improving the redundancy and reliability of the engine system.

[0089] The above description is merely a preferred embodiment of the present invention and an explanation of the technical principles employed. Those skilled in the art should understand that the scope of disclosure in this invention is not limited to technical solutions formed by specific combinations of the above-described technical features, but should also cover other technical solutions formed by arbitrary combinations of the above-described technical features or their equivalents without departing from the above-disclosed concept. For example, technical solutions formed by substituting the above features with (but not limited to) technical features with similar functions disclosed in this invention.

[0090] It should be noted that the terms "first," "second," etc., used in the specification and claims of this application are used to distinguish similar objects and represent a limitation on a specific order or sequence. Where appropriate, the order of use for similar objects can be interchanged so that the embodiments of this application described herein can be implemented in an order other than that shown or described.

[0091] Although embodiments of the present invention have been shown and described above, it is understood that the above embodiments are exemplary and should not be construed as limiting the present invention. Those skilled in the art can make changes, modifications, substitutions and variations to the above embodiments within the scope of the present invention.

Claims

1. A gas generator circulating liquid rocket engine system with electric bypass, characterized in that, include: Main pump pressure supply circulation system, discharge system and electric bypass supply system; The main pump pressure supply cycle system includes a fuel main pump, an oxidizer main pump, a turbine, a gas generator, a main thrust chamber, and a main injector; the gas generator receives propellant from the fuel main pump and the oxidizer main pump and generates gas to drive the turbine, which in turn drives the fuel main pump and the oxidizer main pump; the main injector is located at the head of the main thrust chamber and injects propellant into the main thrust chamber. The discharge system includes a fuel discharge valve, an oxidant discharge valve, a fuel discharge pipeline, and an oxidant discharge pipeline; the fuel discharge valve is located at the outlet of the main fuel pump, the oxidant discharge valve is located at the outlet of the main oxidant pump, the fuel discharge pipeline is connected to the fuel discharge valve, and the oxidant discharge pipeline is connected to the oxidant discharge valve. The electric bypass supply system includes a fuel inlet, an oxidant inlet, a fuel electric pump, an oxidant electric pump, a bypass-specific injector, a fuel check valve, and an oxidant check valve. The fuel inlet is connected to the fuel discharge pipeline downstream of the fuel discharge valve, and the oxidant inlet is connected to the oxidant discharge pipeline downstream of the oxidant discharge valve. The inlet of the fuel electric pump is connected to the fuel liquid inlet, and the outlet of the fuel electric pump is connected to the bypass dedicated injector through the fuel check valve. The inlet of the electric oxidant pump is connected to the oxidant liquid outlet, and the outlet of the electric oxidant pump is connected to the bypass dedicated injector through the oxidant one-way valve. The bypass dedicated injector is located at the head of the main thrust chamber and is independent of the main injector. The fuel check valve is located downstream of the fuel electric pump and upstream of the bypass injector, and the oxidant check valve is located downstream of the oxidant electric pump and upstream of the bypass injector.

2. The gas generator circulating liquid rocket engine system with electric bypass according to claim 1, characterized in that, When the electric bypass supply system is in operation, the fuel discharge valve and the oxidant discharge valve are closed, and the propellant in the fuel discharge pipeline and the oxidant discharge pipeline enters the fuel electric pump and the oxidant electric pump respectively through the fuel liquid intake port and the oxidant liquid intake port.

3. The gas generator circulating liquid rocket engine system with electric bypass according to claim 2, characterized in that, When the main pump pressure supply circulation system is working, the electric bypass supply system is in standby mode, and the fuel check valve and the oxidant check valve are closed under the pressure of the main thrust chamber.

4. The gas generator circulating liquid rocket engine system with electric bypass according to claim 1, characterized in that, Both the fuel electric pump and the oxidant electric pump are centrifugal pumps driven by high-speed motors. The design values ​​of the outlet pressure of the fuel electric pump and the oxidant electric pump are greater than the working pressure of the main thrust chamber under extremely low operating conditions.

5. The gas generator circulating liquid rocket engine system with electric bypass according to claim 4, characterized in that, The bypass-specific injector includes a central atomizing nozzle and a wall cooling nozzle, with the wall cooling nozzle facing the inner wall of the main thrust chamber.

6. The gas generator circulating liquid rocket engine system with electric bypass according to claim 5, characterized in that, The fuel check valve is a differential pressure driven check valve, which opens when the pressure in the main thrust chamber is less than the outlet pressure of the fuel electric pump; the oxidant check valve is a differential pressure driven check valve, which opens when the pressure in the main thrust chamber is less than the outlet pressure of the oxidant electric pump.

7. The gas generator circulating liquid rocket engine system with electric bypass according to claim 6, characterized in that, The fuel check valve closes when the pressure in the main thrust chamber is greater than the outlet pressure of the fuel electric pump, and the oxidant check valve closes when the pressure in the main thrust chamber is greater than the outlet pressure of the oxidant electric pump.

8. The gas generator circulating liquid rocket engine system with electric bypass according to claim 1, characterized in that, It also includes an electronic controller and a chamber pressure sensor. The chamber pressure sensor is located in the main thrust chamber and is used to collect the pressure signal of the main thrust chamber and transmit it to the electronic controller. The electronic controller is used to control the operation of the fuel discharge valve, the oxidizer discharge valve, the fuel electric pump and the oxidizer electric pump according to the pressure signal.

9. The gas generator circulating liquid rocket engine system with electric bypass according to claim 8, characterized in that, Upon receiving an extremely low operating condition command, the electronic controller reduces or closes the main fuel valve and the main oxidant valve in the main pump pressure supply circulation system, and closes the fuel discharge valve and the oxidant discharge valve. When the chamber pressure sensor detects that the pressure value of the main thrust chamber is less than the outlet pressure value of the fuel electric pump and less than the outlet pressure value of the oxidant electric pump, the electronic controller starts the fuel electric pump and the oxidant electric pump.

10. The gas generator circulating liquid rocket engine system with electric bypass according to claim 1, characterized in that, The electric bypass supply system is activated as an independent emergency power source when the main propulsion system fails. The fuel drain valve and the oxidizer drain valve are closed, and the electric fuel pump and the electric oxidizer pump are activated. Propellant enters the electric fuel pump and the electric oxidizer pump through the fuel inlet and the oxidizer inlet. After being pressurized, it is injected into the main thrust chamber through the bypass dedicated injector to generate micro-thrust.

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