A multi-tube engine system suitable for a recoverable reusable rocket and a control method thereof

Abstract

This invention relates to a multi-tube engine system and its control method suitable for reusable rockets. The turbopump includes a coaxially rotating gas turbine, a fuel pump, and an oxidizer pump. A first gas generator is connected to the gas turbine to drive its rotation. It also includes multiple thrust chambers connected in parallel. The fuel outlet of the fuel pump is connected to a fuel path annular flow equalizer, and the oxidizer outlet of the oxidizer pump is connected to an oxidizer path annular flow equalizer. Each thrust chamber is connected to both the fuel path annular flow equalizer and the oxidizer path annular flow equalizer. A second gas generator is also connected to the gas turbine to drive its rotation. The multi-tube engine system of this application has a simple structure, reducing the overall propellant delivery pipeline structure mass. All components can participate in test runs and life evaluation tests, fully meeting test coverage requirements, achieving consistency between ground and space, with a short process test cycle, fewer test runs, and low cost.

Description

TECHNICAL FIELD

[0001] The present application relates to the technical field of liquid rocket engines, and particularly relates to a multi-tube engine system suitable for a reusable rocket and a control method thereof. BACKGROUND

[0002] With the rise and rapid development of commercial spaceflight, the competition in the commercial launch market is becoming increasingly fierce. As the world's largest commercial space company in terms of market share, SpaceX has been constantly breaking its own record of creating new reuse records and successfully recycling records of continuous recycling of the "Falcon 9" rocket. It has been proven that the reuse of rockets greatly reduces launch costs and achieves cost advantages, winning the international market. The nine-parallel arrangement of the nine engines has strong fault tolerance, and one engine is allowed to work in failure among the nine engines, which improves the reliability of the multi-stage parallel engine. Other commercial space companies at home and abroad basically follow this strategy, with seven or nine parallel gas generator cycle engines. In general, this strategy has obvious advantages. First, the technology is mature, the development difficulty is moderate, and the performance of the single engine is improved through gradual iteration and upgrading. The development and testing of the single engine are improved, the cost of single engine testing is small, and the testing capacity of the test bed is small. Second, through the nine-parallel arrangement, the rocket carrying capacity is improved, and the engine system reliability is improved. Finally, through recycling and reuse, the launch cost and launch cycle are greatly reduced, achieving the lowest cost and highest performance in the current commercial space market. Other commercial space companies at home and abroad basically follow this strategy, with seven or nine parallel gas generator cycle engines. In general, this strategy has obvious advantages. First, the technology is mature, the development difficulty is moderate, and the performance of the single engine is improved through gradual iteration and upgrading. The development and testing of the single engine are improved, the cost of single engine testing is small, and the testing capacity of the test bed is small. Second, through the nine-parallel arrangement, the rocket carrying capacity is improved, and the engine system reliability is improved. Finally, through recycling and reuse, the launch cost and launch cycle are greatly reduced, achieving the lowest cost and highest performance in the current commercial space market.

[0003] However, the multi-engine parallel form still has many defects and deficiencies that cannot be ignored. Each engine in the multi-engine parallel engine needs to be subjected to process test, and the cycle required to complete the process test of the entire engine is long, and the total test cost is high. Due to the high cost of power system test, full power system test is rarely performed, and the test is insufficient and cannot fully cover the mission profile, which is a very serious problem for reusable engines. Each engine propellant inlet of the multi-engine parallel scheme needs to be individually connected to the overall tank through a pipeline, and multiple accumulators need to be set to suppress longitudinal low-frequency coupled vibration. The more the parallel engines, the more complex the supply pipeline, the greater the structural mass, and the greater the axial size occupied, which occupies the axial space of the overall system; each thrust chamber is provided with a turbine pump, the structural mass is large, and the outer size of the multi-engine parallel is large. There are as many starting and ignition systems and sub-systems as the number of engines, and there are more valves and assembled elements, and the structural mass is also large. SUMMARY

[0004] The purpose of the present application is to provide a multi-tube engine system suitable for a reusable rocket and a control method thereof, which solves the problems of system complexity and high test cost of the existing reusable rocket multi-engine parallel.

[0005] In order to solve the above problems, the present application provides a multi-tube engine system suitable for a reusable rocket, comprising a turbine pump, the turbine pump comprising a gas turbine, a fuel pump and an oxidizer pump rotating synchronously on the same shaft, a first gas generator (main gas generator) connected to the gas turbine to drive the rotation of the gas turbine, a fuel inlet of the fuel pump being in communication with a fuel tank, an oxidizer inlet of the oxidizer pump being in communication with an oxidizer tank, a plurality of parallel thrust chambers, a fuel outlet of the fuel pump being connected with a fuel ring-shaped flow equalizer, an oxidizer outlet of the oxidizer pump being connected with an oxidizer ring-shaped flow equalizer, a fuel inlet of each of the thrust chambers being in communication with the fuel ring-shaped flow equalizer through a first fuel pipe, and an oxidizer inlet of each of the thrust chambers being in communication with the oxidizer ring-shaped flow equalizer through a first oxidizer pipe.

[0006] The gas turbine is further connected with a second gas generator (auxiliary gas generator) to drive the rotation of the gas turbine, and the rated power of the second gas generator is 1-9 to 1 / 3 of the rated power of the first gas generator; during the rocket launch and acceleration ascent phase, the first gas generator drives the rotation of the gas turbine, and during the rocket recovery phase, the second gas generator drives the rotation of the gas turbine.

[0007] The multi-tube engine system suitable for a reusable rocket provided by the present application further has the following technical features:

[0008] Further, the gas turbine is also connected with a starting gas cylinder through a pipeline, and a starting control valve is arranged on the communication pipeline of the starting gas cylinder and the gas turbine; the fuel inlet of the first fuel generator is communicated with the fuel outlet of the fuel pump through a pipeline, and a first fuel generator fuel auxiliary valve (main fuel generator fuel auxiliary valve) is arranged on the communication pipeline; the oxidant inlet of the first fuel generator is communicated with the oxidant outlet of the oxidant pump through a pipeline, and a first fuel generator oxidant auxiliary valve (main fuel generator oxidant auxiliary valve) is arranged on the communication pipeline; the fuel inlet of the second fuel generator is communicated with the fuel outlet of the fuel pump through a pipeline, and a second fuel generator fuel auxiliary valve (auxiliary fuel generator fuel auxiliary valve) is arranged on the communication pipeline; the oxidant inlet of the second fuel generator is communicated with the oxidant outlet of the oxidant pump through a pipeline, and a second fuel generator oxidant auxiliary valve (auxiliary fuel generator oxidant auxiliary valve) is arranged on the communication pipeline.

[0009] Further, the first fuel pipe is sequentially provided with a thrust chamber fuel path electric butterfly valve and a thrust chamber fuel main valve from the fuel path annular flow distributor to the thrust chamber.

[0010] Further, the first fuel pipe is sequentially provided with a thrust chamber fuel path electric butterfly valve and a thrust chamber fuel main valve from the fuel path annular flow distributor to the thrust chamber.

[0011] Further, the plurality of parallel thrust chambers include a central thrust chamber and a plurality of outer ring thrust chambers, and the outer ring thrust chambers are arranged uniformly in a circumferential direction with the central thrust chamber as the center.

[0012] Further, the number of the outer ring thrust chambers is six or eight.

[0013] Further, the thrust chambers can swing bidirectionally, the swing circle of the central thrust chamber is 5°, and the swing circle of the outer ring thrust chamber is 2.5°.

[0014] The application also provides a control method suitable for the multi-tube engine of the reusable rocket, which is based on the multi-tube engine system suitable for the reusable rocket and includes the following steps.

[0015] S10, the gas turbine of the turbine pump is driven by the rocket engine starting module to work, the fuel pump and the oxidant pump are started with the gas turbine, the pressure in the pipeline after the pump is increased, the first gas generator oxidant sub valve (main gas generator oxidant sub valve) and the first gas generator fuel sub valve (main gas generator fuel sub valve) of the first gas generator are opened, the propellant enters the first gas generator (main gas generator), the first gas generator is ignited, the high-temperature rich gas generated after the first gas generator is ignited enters the gas turbine to drive the turbine pump to work continuously, and the rocket engine starting module is closed; the pressure after the pump is increased to the required pressure, the fuel and the oxidant are provided for each thrust chamber according to the time sequence determined by the test, each thrust chamber is ignited and started, after successful ignition, with the gradual increase of the pressure after the turbine pump, the power of the turbine pump is increased, the system enters the main stage working state, and the ground starting process is ended.

[0016] S20, after the rocket propels the load to the predetermined height, the first-stage rocket engine is shut down, the first-stage rocket is separated from the rocket body and enters the recovery procedure; after the system issues a shutdown instruction, the first gas generator stops working, the power of the turbine pump is attenuated, the fuel and the oxidant pipelines of each thrust chamber are closed, the engine thrust is attenuated to zero, and the shutdown process is ended.

[0017] S30, after the first-stage rocket is separated from the rocket body, the first-stage rocket enters the recovery stage; after the system issues a starting instruction, the gas turbine of the turbine pump is driven by the rocket engine starting module to work, the fuel pump and the oxidant pump are started with the gas turbine, the pressure in the pipeline after the pump is increased, the second gas generator oxidant sub valve (sub gas generator oxidant sub valve) and the second gas generator fuel sub valve (sub gas generator fuel sub valve) of the second gas generator are opened, the propellant enters the second gas generator (sub gas generator), the second gas generator is ignited, the high-temperature rich gas generated after the second gas generator is ignited enters the gas turbine to drive the turbine pump to work continuously, and the rocket engine starting module is closed; the pressure after the pump is increased to the required pressure, the fuel and the oxidant are provided for the thrust chamber according to the time sequence determined by the test, the thrust chamber is ignited and started, after successful ignition, with the gradual increase of the pressure after the turbine pump, the power of the turbine pump is increased, the system enters the main stage working state, and the in-flight starting process is ended.

[0018] Further, the following steps are further included:

[0019] S11, in the rocket ascending stage, when the thrust needs to be adjusted, the oxidant flow of the first gas generator is adjusted through the control valve, the propellant mixing ratio and flow of the first gas generator are changed, the first gas generator chamber pressure and the gas temperature are changed, the energy of the working gas input into the gas turbine is changed, the output power of the gas turbine is changed, and the output power of the oxidant pump and the fuel pump is changed accordingly;

[0020] S31, in the rocket recovery phase, when thrust adjustment is needed, the oxidant flow of the second gas generator is adjusted through the control valve, the propellant mixing ratio and flow of the second gas generator are changed, the chamber pressure and gas temperature of the second gas generator are changed, the energy of the working gas input into the gas turbine is changed, the output power of the gas turbine is changed, and the output power of the oxidant pump and the fuel pump is changed accordingly.

[0021] Further, in the rocket recovery phase, multiple shutdowns and in-flight starts can be performed as needed.

[0022] The application has the following beneficial effects: the multi-tube engine system and control method suitable for reusable rockets according to the application, multiple parallel rocket engine thrust chambers form a multi-tube engine, multiple thrust chambers are pumped by a same turbine pump with fuel and oxidant, the structure is simple, and the overall propellant delivery pipeline structure mass is reduced; the turbine pump is relatively far from the thrust chamber, and the pipeline between the turbine pump and the thrust chamber has a flexible connection structure, which can reduce the coupled vibration between the two vibration sources of the thrust chamber and the turbine pump, improve the mechanical environment of each component, improve the structural reliability of the engine, and lay a solid foundation for rocket reuse; all components of the multi-tube engine system can participate in the test and life evaluation test, fully meet the test coverage requirements, and realize consistency between the earth and the sky; the whole system test can be performed, or the test can be performed when part of the thrust units of the rocket work, the coordination of the multi-tube engine system can be fully evaluated, and the possibility of flight failure can be greatly reduced; the one-stage rocket using the multi-tube engine system only needs to be subjected to a process test, the process test cycle is short, the number of times is small, and the cost is low. BRIEF DESCRIPTION OF DRAWINGS

[0023] Figure 1 The figure is a system schematic diagram of the multi-tube engine system suitable for reusable rockets according to the embodiment of the application;

[0024] Figure 2 The figure is a thrust chamber arrangement mode schematic diagram in the embodiment of the application;

[0025] Figure 3 The figure is a specific structure schematic diagram of the multi-tube engine system suitable for reusable rockets according to the embodiment of the application. DETAILED DESCRIPTION

[0026] Hereinafter, the application will be described in detail with reference to the accompanying drawings and in combination with the embodiments. It should be noted that the embodiments in the application and the features in the embodiments can be combined with each other without conflict.

[0027] As Figures 1 to 2In the shown embodiment of the application, the multi-tube engine system suitable for reusable rocket and its control method comprises a turbine pump, the turbine pump comprises a synchronous rotating gas turbine 3, a fuel pump 5, and an oxidant pump 4, the gas turbine 3 is connected with a first gas generator 7 (main gas generator) for driving the rotation of the gas turbine 3, the fuel inlet of the fuel pump 5 is communicated with a fuel tank, the oxidant inlet of the oxidant pump 4 is communicated with an oxidant tank, and the multi-tube engine system further comprises a plurality of parallel thrust chambers 100, the fuel outlet of the fuel pump 5 is connected with a fuel loop annular flow distributor 8, the oxidant outlet of the oxidant pump 4 is connected with an oxidant loop annular flow distributor 9, the fuel inlet of each thrust chamber 100 is communicated with the fuel loop annular flow distributor 8 through a first fuel pipe, and the oxidant inlet of each thrust chamber 100 is communicated with the oxidant loop annular flow distributor 9 through a first oxidant pipe; the gas turbine 3 is further connected with a second gas generator 6 (auxiliary gas generator) for driving the rotation of the gas turbine 3, the rated power of the second gas generator 6 is 1 / 9 to 1 / 3 of the rated power of the first gas generator 7, and preferably, the rated power of the second gas generator 6 is 1 / 3 of the rated power of the first gas generator 7; in the rocket launch and ascending phase, the rotation of the gas turbine 3 is driven by the first gas generator 7, and in the rocket recovery phase, the rotation of the gas turbine 3 is driven by the second gas generator 6.

[0028] The multi-tube engine system of the application is suitable for a recoverable reusable rocket. A plurality of parallel rocket engine thrust chambers form a multi-tube engine. A plurality of thrust chambers are pumped by a same turbine pump with fuel and oxidant. No matter how many thrust chambers are, only two paths of propellant main pipelines of a tank and an engine are needed. Only one pressure accumulator is needed at an oxidant inlet. The structure is not only simple, but also reduces the overall propellant delivery pipeline structure mass. The mechanical environment of each component is good. The turbine pump is relatively far from the thrust chamber. The pipeline between the turbine pump and the thrust chamber is flexibly connected. The coupling vibration between the two vibration sources of the thrust chamber and the turbine pump can be reduced. The mechanical environment of each component is improved. The structural reliability of the engine is improved. A solid foundation is laid for rocket reuse. The degree of integration is high. Only one set of starting, control and blowing system and one high-power turbine pump are needed. One set of auxiliary system and one set of propellant supply pipeline make the engine structure mass small and the thrust-to-weight ratio high. The system is simple and has strong fault tolerance. One thrust unit is allowed to fail. The failed thrust unit is isolated. The overall reliability is higher than that of a multi-engine parallel engine. Like a pump-after-swing engine, only the thrust chamber swings. Compared with a pump-before-swing, the swing torque required for attitude control is small. The demand for load capacity of the servo mechanism is reduced. The multi-tube engine with multiple parallel thrust chambers makes the layout more compact. The surface thrust of the rocket is relatively high. Especially, the surface thrust of the seven-jet engine is larger. Like a multi-engine parallel engine, each thrust unit can be started and stopped individually to meet the requirements of the recoverable rocket. The working condition can be changed to group shutdown to reduce the shutdown water hammer pressure. All components of the multi-tube engine system can participate in the test and life test. The test coverage requirements are fully met. The consistency of heaven and earth is realized. The whole system test can be performed. The test of the work of part of the thrust units when the rocket returns can be performed. The coordination of the multi-tube engine system can be fully evaluated. The possibility of flight failure is greatly reduced. A one-stage rocket using the multi-tube engine system only needs to perform a process test. The process test cycle is short. The number of times is small. The cost is low.

[0029] The multi-tube engine system suitable for the recyclable reusable rocket of the application can avoid the combustion deterioration of the gas generator in the rocket recovery stage caused by the large change of working conditions in the rocket launch acceleration ascending stage and the rocket recovery stage. Specifically, taking nine parallel thrust chambers as an example, when launching a load, nine thrust chambers work, allowing the engine to work normally and the load task to be completed without being affected by the failure of one thrust unit to close; in the rocket launch acceleration ascending stage, nine thrust chambers work, and the turbine pump adjusts the working condition as needed. The working condition changes little in this stage, and the oxidant flow of the gas generator can be changed to meet the variable working condition requirement; in the rocket recovery stage, the required engine thrust is very small, only about 1 / 3~1 / 9 of the full working condition, and only 1~3 thrust chambers work in the return process; the working condition of the turbine pump is only 11%~33% of the rated working condition, which seriously deviates from the design working condition, resulting in a decrease in the efficiency of the turbine pump, but this is acceptable in engineering; if there is only one gas generator, the gas generator will work at 11% of the working condition in the rocket recovery stage, which will have the risk of combustion deterioration; the application sets two gas generators for driving the gas turbine to rotate, and even if only one thrust chamber works in the rocket recovery stage, the minimum working condition of the second gas generator is 33% of the rated working condition, which greatly improves the working condition of the gas generator; the oxygen flow of the second gas generator can be adjusted according to the variable working condition of the thrust chamber to change the output power of the gas turbine.

[0030] In an embodiment of the application, preferably, the gas turbine 3 is further connected with a starting gas cylinder 1 through a pipeline, and a starting control valve 2 is further arranged on the communication pipeline between the starting gas cylinder 1 and the gas turbine 3; the fuel inlet of the first gas generator 7 is communicated with the fuel outlet of the fuel pump 5 through a pipeline, and a first gas generator fuel auxiliary valve 72 (main gas generator fuel auxiliary valve) is arranged on the communication pipeline; the oxidant inlet of the first gas generator 7 is communicated with the oxidant outlet of the oxidant pump 4 through a pipeline, and a first gas generator oxidant auxiliary valve 71 (main gas generator oxidant auxiliary valve) is arranged on the communication pipeline; the fuel inlet of the second gas generator 6 is communicated with the fuel outlet of the fuel pump 5 through a pipeline, and a second gas generator fuel auxiliary valve 61 (auxiliary gas generator fuel auxiliary valve) is arranged on the communication pipeline; the oxidant inlet of the second gas generator 6 is communicated with the oxidant outlet of the oxidant pump 5 through a pipeline, and a second gas generator oxidant auxiliary valve 62 (auxiliary gas generator oxidant auxiliary valve) is arranged on the communication pipeline.

[0031] This application describes a multi-tube engine system applicable to reusable rockets. During ground-launch full-thrust start-up, it offers three starting options: cylinder start-up, propellant starter start-up, and self-start-up. This application... Figure 1 The illustrated embodiment describes cylinder starting. Specifically, during the ground starting preparation phase, the high-pressure helium in starting cylinder 1 is filled. After completing other starting preparations, the engine control system issues a starting command, initiating the starting sequence. The starting control valve 2 opens, and the high-pressure gas in starting cylinder 1 enters the gas turbine 3. After expanding through the turbine nozzle, it drives the gas turbine blades to perform work. The gas turbine 3 is coaxially connected to the oxidizer pump 4 and fuel pump 5. After the gas turbine 3 drives the oxidizer pump 4 and fuel pump 5 to start, the pressure in the downstream pipeline increases. The first gas generator oxidizer auxiliary valve 71 (main gas generator oxidizer auxiliary valve 71) of the first gas generator 7... The oxidizer auxiliary valve of the generator and the fuel auxiliary valve 72 of the first gas generator (fuel auxiliary valve of the main gas generator) are opened, and propellant enters the first gas generator 7 (main gas generator). The first gas generator 7 is ignited, and the high-temperature fuel-rich gas generated by the ignition enters the gas turbine 3, driving the turbopump to work continuously. The start-up control valve 2 is closed, and the pressure after the pump rises to the required pressure. According to the timing determined by the test, fuel and oxidizer are provided to each thrust chamber. Each thrust chamber is ignited and started. After successful ignition, as the pressure after the turbopump gradually increases, the turbopump power increases, the system enters the main stage working state, and the ground start-up process ends.

[0032] The multi-tube engine system described above, applicable to reusable rockets, employs an open-loop gas generator cycle. The gas generator operates under fuel-rich conditions. Most of the propellant enters the thrust chamber via fuel and oxidizer pumps, where it burns to produce high-temperature, high-pressure gas. This gas is accelerated through nozzles and ejected at high speed to generate thrust. A small amount of propellant enters the gas generator and burns to produce high-temperature, high-pressure, fuel-rich gas. This fuel-rich gas then enters the turbopump to perform work, replacing the cooling gas to drive the turbine. After performing work, the fuel-rich gas enters the heat exchanger and is discharged through the exhaust pipe. A small amount of liquid oxygen enters the heat exchanger and is converted into oxygen, pressurizing the oxygen storage tank.

[0033] In one embodiment of this application, preferably, each of the first fuel pipes is sequentially equipped with a thrust chamber fuel line electric butterfly valve 102 and a thrust chamber fuel main valve 104 from the fuel line annular flow equalizer 8 to the thrust chamber 100; the first oxidizer pipe is sequentially equipped with a thrust chamber oxidizer line electric butterfly valve 101 and a thrust chamber oxidizer main valve 103 from the oxidizer line annular flow equalizer 9 to the thrust chamber 100. This embodiment, applicable to the multi-tube engine system of reusable rockets, allows each thrust chamber 100 to be individually controlled via the thrust chamber oxidizer line electric butterfly valve 101, the thrust chamber oxidizer main valve 103, the thrust chamber fuel line electric butterfly valve 102, and the thrust chamber fuel main valve 104, thereby improving system reliability.

[0034] Specifically, in the ground launch full-thrust starting phase, after the gas turbine 3 starts to drive the turbine pump to work continuously, the post-pump pressure of the fuel pump 5 and the oxidant pump 4 is increased to the required pressure, and the oxidant road electric butterfly valve 101 and the fuel road electric butterfly valve 102 of each thrust chamber 100 are opened according to the time sequence determined by the test, and the pressure is increased to the specified pressure, and then the thrust chamber oxidant main valve 103 and the thrust chamber fuel main valve 104 of each thrust chamber 100 are automatically opened and are in a self-sustaining state. After the propellant enters each thrust chamber, the thrust chamber is ignited, and after successful ignition, as the post-pump pressure gradually increases, the turbine pump power increases, and the engine enters the main stage working state, and the ground starting process ends.

[0035] In an embodiment of the present application, preferably, a thrust chamber fuel road swing hose 106 is further arranged between the thrust chamber fuel main valve 104 on the first fuel pipe and the thrust chamber 100, and a thrust chamber oxidant road swing hose 105 is further arranged between the thrust chamber oxidant main valve 103 on the first oxidant pipe and the thrust chamber 100; so that the thrust chamber can swing within a certain range as needed; and like the post-pump swing engine, only the thrust chamber swings, and compared with the pre-pump swing, the swing torque required for attitude control is small, reducing the demand for load capacity of the servo mechanism.

[0036] In an embodiment of the present application, preferably, as shown in Figure 2 The plurality of parallel thrust chambers 100 include a central thrust chamber 110 and a plurality of outer circle thrust chambers 120, and the outer circle thrust chambers 120 are arranged uniformly in a circumferential direction with the central thrust chamber 110 as the center. Preferably, the number of outer circle thrust chambers 120 is six or eight. Taking the number of outer circle thrust chambers 120 as eight as an example, the eight outer circle thrust chambers 120 are sequentially numbered I, II, III, IV, V, VI, VII, and VIII, and the central thrust chamber 110 is numbered IX. The multi-tube engine with multiple parallel thrust chambers makes the layout more compact, and the surface thrust of the rocket is relatively high.

[0037] In an embodiment of the present application, preferably, the thrust chambers 100 can swing bidirectionally, the swing circle of the central thrust chamber 110 is 5°, and the swing circle of the outer circle thrust chamber 120 is 2.5°. Specifically, there are nine or seven thrust units in total, one in the center and eight or six in the outer circle; each thrust unit has a bidirectional swing function, the swing circle of the outer circle engine is 2.5°, the swing circle of the central engine is 5°, the roll control is the same tangential swing, the pitch and yaw or compound control swing directions are the same; the swing angle of each thrust chamber can be flexibly controlled as needed, and the attitude of the rocket can be flexibly controlled and adjusted.

[0038] The present application also provides a control method suitable for the multi-tube engine of the reusable rocket, which comprises the multi-tube engine system suitable for the reusable rocket and further comprises the following steps:

[0039] S10, the gas turbine of the turbine pump is driven by the rocket engine starting module to rotate and drive the gas turbine blade to work, the fuel pump and the oxidant pump rotate with the gas turbine, the pressure in the pipeline after the pump rises, the first gas generator oxidant sub valve 71 (main gas generator oxidant sub valve) and the first gas generator fuel sub valve 72 (main gas generator fuel sub valve) of the first gas generator 7 are opened, the propellant enters the first gas generator 7 (main gas generator), the first gas generator 7 is ignited, the high-temperature rich gas generated after the first gas generator is ignited enters the gas turbine 3 to drive the turbine pump to work continuously, and the rocket engine starting module is closed; the pressure after the pump rises to the required pressure, the fuel and the oxidant are provided for each thrust chamber according to the time sequence determined by the test, each thrust chamber is ignited and started, after successful ignition, with the gradual increase of the pressure after the turbine pump, the power of the turbine pump increases, the system enters the main stage working state, and the ground starting process ends.

[0040] S20, after the rocket propels the load to a predetermined height, the first-stage rocket engine is shut down, the first-stage rocket is separated from the rocket body and enters the recovery program; after the system issues a shutdown instruction, the first gas generator stops working, the power of the turbine pump decays, the fuel and oxidant pipelines of each thrust chamber are closed, the engine thrust decays to zero, and the shutdown process ends.

[0041] S30, after the first-stage rocket is separated from the rocket body, the first-stage rocket enters the recovery stage; after the system issues a starting instruction, the gas turbine of the turbine pump is driven by the rocket engine starting module to rotate and drive the gas turbine blade to work, the fuel pump and the oxidant pump rotate with the gas turbine, the pressure in the pipeline after the pump rises, the second gas generator oxidant sub valve 61 (sub gas generator oxidant sub valve) and the second gas generator fuel sub valve 62 (sub gas generator fuel sub valve) of the second gas generator 6 are opened, the propellant enters the second gas generator 6 (sub gas generator), the second gas generator 6 is ignited, the high-temperature rich gas generated after the second gas generator is ignited enters the gas turbine 3 to drive the turbine pump to work continuously, and the rocket engine starting module is closed; the pressure after the pump rises to the required pressure, the fuel and the oxidant are provided for the thrust chamber according to the time sequence determined by the test, the thrust chamber is ignited and started, after successful ignition, with the gradual increase of the pressure after the turbine pump, the power of the turbine pump increases, the system enters the main stage working state, and the air starting process ends.

[0042] In an embodiment of the present application, preferably, the method further comprises the steps of:

[0043] S11. During the rocket's ascent phase, when thrust adjustment is required, the oxidizer flow rate of the first gas generator is adjusted by the control valve, thereby changing the propellant mixing ratio and flow rate of the first gas generator, which in turn changes the chamber pressure and gas temperature of the first gas generator, changes the energy of the working gas input to the gas turbine, changes the output power of the gas turbine, and consequently changes the output power of the oxidizer pump and fuel pump.

[0044] S31. During the rocket recovery phase, when thrust adjustment is required, the oxidizer flow rate of the second gas generator is adjusted by controlling the valve, thereby changing the propellant mixing ratio and flow rate of the second gas generator, which in turn changes the chamber pressure and gas temperature of the second gas generator, changes the energy of the working gas input to the gas turbine, changes the output power of the gas turbine, and consequently changes the output power of the oxidizer pump and fuel pump.

[0045] In one embodiment of this application, preferably, during the rocket recovery phase, the rocket can be shut down and restarted in mid-air multiple times as needed.

[0046] The following uses a 9-thrust-chamber example to illustrate the multi-tube engine system and its control method applicable to reusable rockets in this application.

[0047] like Figure 1 As shown, the cross-sections of thrust chambers VII-IX-III are schematically given, and the arrangement of the nine thrust chambers is as follows: Figure 2 As shown, each thrust chamber is connected to the fuel path annular flow equalizer 8 and the oxidant path annular flow equalizer 9 via pipelines.

[0048] Working principle:

[0049] The multi-tube engine uses an open-loop gas generator cycle. The gas generator operates under fuel-rich conditions. Most of the propellant enters the thrust chamber through the fuel and oxidizer pumps and is burned to produce high-temperature, high-pressure gas, which is accelerated through the nozzle and ejected at high speed to generate thrust. A small amount of propellant enters the gas generator and is burned to produce high-temperature, high-pressure, fuel-rich gas. This fuel-rich gas enters the turbopump to do work, replacing the cool gas to drive the turbine. After doing work, the fuel-rich gas enters the heat exchanger and is discharged through the exhaust pipe. A small amount of liquid oxygen enters the heat exchanger and is converted into oxygen to pressurize the oxygen tank. Rocket attitude control only requires oscillating the thrust chamber. The oscillation torque is very small, and the servo mechanism can meet the requirements by using an electric method.

[0050] Thrust unit layout: There are nine or seven thrust units in total, one in the center and eight or six in the outer ring. Each thrust unit has a bidirectional oscillation function. The outer ring engines oscillate at 2.5° and the center engine oscillates at 5°. Roll control is tangential oscillation, and pitch, yaw, or combined control oscillate in the same direction.

[0051] There are two modes of recovery, remote recovery and local recovery. In remote recovery, the engine works for 3 times at most, the central thrust unit works for 3 times, and the rest of the thrust units have the ability to work for 2 times. In local recovery, the engine works for 4 times, the central thrust unit works for 4 times, and the rest of the thrust units have the ability to work for 2 times.

[0052] When launching the load, all 9 thrust chambers work, allowing the engine to work normally as a whole and the load task to be completed without being affected by the failure of one thrust unit. During the rocket launch and ascent phase, all the parallel thrust chambers of the substage power system work, and the turbine pump adjusts the working condition as needed. The working condition changes little in this stage, and the requirement of variable working condition can be met by changing the oxidant flow of the generator.

[0053] During the recovery phase, the required engine thrust is very small, only about 1 / 3 to 1 / 9 of the full working condition, and only 1 to 3 thrust chambers need to work during the return process. The working condition of the turbine pump is only 11% to 33% of the rated working condition, which deviates from the design working condition seriously, resulting in a decrease in the efficiency of the turbine pump. However, this is acceptable in engineering. The gas generator works at 11% of the working condition, which has certain risks of combustion deterioration. Therefore, the gas generator is set to two, one is the main gas generator, and the other is the auxiliary gas generator. The main gas generator works during the launch and load phase, and the auxiliary gas generator works during the recovery phase. During the recovery process, even if only one thrust unit is working, the minimum working condition of the generator is 33% of the rated working condition, which improves the working condition of the generator. The engine variable working condition can adjust the oxygen flow of the auxiliary generator as needed to change the turbine output power.

[0054] The system that supplies propellant to the thrust chamber from the engine propellant inlet under the specified working condition is called the main system, which includes the turbine pump, the oxidizer electric butterfly valve, the fuel electric butterfly valve, the oxidizer main valve, the fuel main valve, and the metal bellows. The metal bellows is set to meet the thrust chamber double swing (swing circle) requirement. The oxidizer main valve and the fuel main valve have fluid on-off function. Both valves are normally closed and open under pressure and maintain the open state. The oxidizer main valve also has a liquid oxygen pre-cooling backflow channel, and the fuel main valve has a discharge channel. When a thrust chamber fails or is not needed to work according to the need, the electric butterfly valve is closed to forcibly cut off the supply of propellant to the corresponding thrust unit, and the rest of the thrust chambers can work normally, improving the fault tolerance of the power system and expanding the thrust adjustment range of the engine. In addition, the electric butterfly valve can also change the mixture ratio of the thrust chamber, which has the function of propellant utilization system.

[0055] The starting can adopt solid starter starting, gas starter starting, liquid starter and self starting, without changing the main components and control mode of the system. The ignition mode of the thrust chamber and the gas generator can select spark plug ignition, pyrotechnic ignition and chemical ignition, etc. The different ignition modes do not affect the main components and control mode of the system. The thrust unit can be nine or seven thrust units, one in the center and eight or six distributed in the outer circle. Or further, the thrust unit can be selected to be more than nine according to the needs. The different number and arrangement of the thrust chamber do not affect the main components and control mode of the system.

[0056] Working process:

[0057] 1) Full-thrust starting of ground launch: The engine system can select three starting schemes of gas cylinder starting, powder starter and self starting. The system diagram given in the application takes the gas cylinder starting as an example, as shown in Figure 1 Specifically, the ground starting preparation stage completes the inflation of the high-pressure helium gas in the starting cylinder 1, and after completing other starting preparation work, the engine control system issues a starting instruction, enters the starting timing, and the starting control valve 2 is opened. The high-pressure gas in the starting cylinder enters the gas turbine 3, expands through the turbine nozzle, and pushes the gas turbine blade to work. The gas turbine 3 is coaxially connected with the oxidant pump 4 and the fuel pump 5. After the gas turbine drives the connected pump to rotate, the pressure in the propellant pump after-pipeline rises, the oxidant auxiliary valve 71 and the fuel auxiliary valve 72 of the first gas generator 7 (main gas generator) are opened, the propellant enters the first gas generator 7 (main gas generator), the first gas generator 7 is ignited, the high-temperature rich gas generated by ignition enters the gas turbine, continues to drive the turbine pump to work continuously, the starting control valve 2 is closed, the pressure after the pump rises to the required pressure, the oxidant road electric butterfly valve 101 and the fuel road electric butterfly valve 102 of each thrust chamber I~IX are opened according to the timing determined by the test, the oxidant main valve 103 and the fuel main valve 104 of each thrust chamber are automatically opened and are in a self-sustaining state after the pressure rises to the specified pressure, the propellant enters each thrust chamber, the thrust chamber is ignited, after the ignition is successful, with the gradual increase of the pressure after the pump, the turbine pump power increases, the engine enters the main stage working state, and the ground starting process is completed.

[0058] 2) Thrust regulation of the main stage working condition in the ascending stage of ground launch: The thrust regulation of the main stage working condition in the ascending stage of ground launch is realized by changing the oxidant flow of the first gas generator 7 (main gas generator), as shown in Figure 1 Specifically, by changing the oxidant flow of the first gas generator 7, the propellant mixing ratio of the main gas generator is changed, the chamber pressure and the gas temperature of the gas generator are changed, the energy of the working gas input into the gas turbine 3 is changed, correspondingly, the power of the oxidant pump and the fuel pump is also changed, the propellant flow into the thrust chamber is changed, and finally the working condition adjustment of the whole engine main stage working is realized.

[0059] 3) Shutdown process during the ascent phase of ground launch: The multi-tube engine system applicable to reusable rockets described in this application is generally used in the first stage of reusable rockets. After the payload is propelled to a predetermined altitude by this multi-tube engine system, the first-stage engine shuts down, the first-stage rocket separates from the rocket body, and enters the recovery procedure. The entire engine shutdown process should be able to quickly and accurately stop the engine from operating, with a small aftereffect impulse deviation. For example... Figure 1 As shown, the specific shutdown process is as follows: After the shutdown command is issued, the oxidizer auxiliary valve 71 of the first gas generator closes, followed by the fuel auxiliary valve 72. The power of the first gas generator and the turbopump decreases. After the power decreases to a certain level, the electric butterfly valves 101 of the oxidizer circuit in thrust chambers I to IX close. After the electric butterfly valves of the oxidizer circuit close, the chamber pressure in the thrust chamber drops rapidly, and the pressure after the valve also drops rapidly. After the pressure drops to the design closing pressure of the oxidizer main valve, the oxidizer main valves 103 of each thrust chamber I to IX close. Then, the electric butterfly valves 102 of the fuel circuit in each thrust chamber I to IX close in sequence. After the electric butterfly valves of the fuel circuit close, the pressure after the valve also drops rapidly. After the pressure drops to the design closing pressure of the fuel main valve, the fuel main valves 104 of each thrust chamber I to IX close. The engine thrust decreases to zero, and the shutdown process ends.

[0060] 4) Second / Third / Fourth In-Flight Start-Ups During Recovery: After the first and second stages of the rocket separate, the first-stage engine enters the recovery phase. During recovery, the engine system only needs the second gas generator 6 to drive the turbopump to do work, which can meet the requirements of 1-3 thrust chambers. Taking the case where only the central thrust chamber is working during reentry as an example, the engine control system issues a start-up command and enters the start-up sequence. The start-up control valve 2 opens, and the high-pressure gas in the start-up gas cylinder enters the gas turbine 3. After expanding through the turbine nozzle, it drives the turbine blades to do work. The gas turbine 3 is coaxially connected to the oxidizer pump 4 and the fuel pump 5. After the gas turbine drives the pumps connected to it to start, the pressure in the propellant pump downstream pipeline increases, and the oxidizer auxiliary valve 61 and fuel auxiliary valve 61 of the second gas generator 6 (auxiliary gas generator) start working. Valve 62 opens, propellant enters the second gas generator 6, the second gas generator 6 ignites, and the high-temperature fuel-rich gas generated by ignition enters the gas turbine, continuing to drive the turbopump to work continuously. Start-up control valve 2 closes, and the pressure after the pump rises to the required pressure. According to the timing determined by the test, the electric butterfly valve 101 of the oxidizer line and the electric butterfly valve 102 of the fuel line in the IX thrust chamber are opened. After the pressure rises to the specified pressure, the engine's oxidizer main valve 103 and fuel main valve 104 automatically open and are in a self-sustaining state. After the propellant enters the thrust chamber, the thrust chamber ignites. After successful ignition, as the pressure after the pump gradually increases, the turbopump power increases, and the engine enters the main stage working state. The in-flight start-up process ends. The working process of the second / third / fourth start-up in the entire recovery phase is consistent.

[0061] 5) Thrust regulation in the recovery phase: the thrust regulation in the main stage working condition in the recovery phase is realized by changing the oxidizer flow rate of the second gas generator 6, as shown in Figure 1 specifically, by changing the oxidizer flow rate of the second gas generator 6, the propellant mixture ratio of the second gas generator 6 is changed, the gas generator chamber pressure and the gas temperature are changed, so that the energy of the working gas input into the gas turbine 3 can be changed, correspondingly, the power of the oxidizer pump and the fuel pump is also changed, the propellant flow rate flowing into the thrust chamber is changed, and finally the working condition adjustment of the main stage working of the whole engine is realized, and the main stage thrust regulation mode in the whole recovery phase is consistent.

[0062] 6) Secondary / tertiary / quaternary shutdown in the recovery phase: after the shutdown instruction is issued, the oxidizer auxiliary valve 61 of the second gas generator 6 is closed, then the fuel auxiliary valve 62 is closed, the power of the second gas generator and the turbine pump is attenuated, after the power is attenuated to a certain degree, the oxidizer road electric butterfly valve 101 of the Ⅸ thrust chamber is closed, after the oxidizer road electric butterfly valve 101 is closed, the chamber pressure of the Ⅸ thrust chamber is rapidly reduced, and the pressure after the valve is also rapidly reduced, after the pressure is reduced to the design closing pressure of the oxidizer main valve, the oxidizer main valve 103 of the Ⅸ thrust chamber is closed, then the fuel road electric butterfly valve 102 of the Ⅸ thrust chamber is closed according to the time sequence, after the fuel road electric butterfly valve is closed, the pressure after the valve is also rapidly reduced, after the pressure is reduced to the design closing pressure of the fuel main valve, the fuel main valve 104 of the Ⅸ thrust chamber is closed, the engine thrust is attenuated to zero, the shutdown process is ended, and the working process of the secondary / tertiary / quaternary shutdown in the whole recovery phase is consistent.

[0063] Finally, it should be noted that the above embodiments are only used to illustrate the technical solutions of the present application, and not to limit them; although the present application has been described in detail with reference to the foregoing embodiments, those skilled in the art should understand that they can still modify the technical solutions recorded in the foregoing embodiments, or make equivalent replacement for part of the technical features; and these modifications or replacements do not make the essence of the corresponding technical solutions deviate from the spirit and scope of the technical solutions of the embodiments of the present application.

Claims

1. A multi-tube engine system suitable for reusable rockets, characterized in that, The system includes a turbopump comprising a synchronously rotating gas turbine, a fuel pump, and an oxidizer pump. A first gas generator is connected to the gas turbine to drive its rotation. The fuel inlet of the fuel pump is connected to a fuel tank, and the oxidizer inlet of the oxidizer pump is connected to an oxidizer tank. The system is characterized by further comprising multiple parallel thrust chambers. The fuel outlet of the fuel pump is connected to a fuel path annular flow equalizer, and the oxidizer outlet of the oxidizer pump is connected to an oxidizer path annular flow equalizer. The fuel inlet of each thrust chamber is connected to the fuel path annular flow equalizer via a first fuel pipe, and the oxidizer inlet of each thrust chamber is connected to the oxidizer path annular flow equalizer via a first oxidizer pipe. A second gas generator is also connected to the gas turbine to drive its rotation. The rated power of the second gas generator is 1 / 9 to 1 / 3 of the rated power of the first gas generator. During the rocket's launch acceleration and ascent phase, the first gas generator drives the gas turbine to rotate; during the rocket recovery phase, the second gas generator drives the gas turbine to rotate. The gas turbine is also connected to a starting gas cylinder via a pipeline, and a starting control valve is provided on the pipeline connecting the starting gas cylinder and the gas turbine. The fuel inlet of the first gas generator is connected to the fuel outlet of the fuel pump via a pipeline, and a first gas generator fuel auxiliary valve is provided on the connecting pipeline. The oxidizer inlet of the first gas generator is connected to the oxidizer outlet of the oxidizer pump via a pipeline, and a first gas generator oxidizer auxiliary valve is provided on the connecting pipeline. The fuel inlet of the second gas generator is connected to the fuel outlet of the fuel pump via a pipeline, and a fuel auxiliary valve for the second gas generator is provided on the connecting pipeline. The oxidant inlet of the second gas generator is connected to the oxidant outlet of the oxidant pump via a pipeline, and a oxidant auxiliary valve for the second gas generator is provided on the connecting pipeline. From the fuel path annular flow equalizer to the thrust chamber, the first fuel pipe is sequentially equipped with a thrust chamber fuel path electric butterfly valve, a thrust chamber fuel main valve, and a thrust chamber fuel path swing hose. From the oxidant path annular flow equalizer to the thrust chamber, the first oxidant pipe is sequentially equipped with a thrust chamber oxidant path electric butterfly valve, a thrust chamber oxidant main valve, and a thrust chamber oxidant path swing hose.

2. The multi-tube engine system according to claim 1, characterized in that: A thrust chamber fuel path swing hose is also provided between the thrust chamber fuel main valve on the first fuel pipe and the thrust chamber, and a thrust chamber oxidant path swing hose is also provided between the thrust chamber oxidant main valve on the first oxidant pipe and the thrust chamber.

3. The multi-tube engine system according to claim 1, characterized in that: The multiple thrust chambers connected in parallel include a central thrust chamber and multiple outer ring thrust chambers, which are evenly arranged circumferentially around the central thrust chamber.

4. The multi-tube engine system according to claim 3, characterized in that: The number of outer thrust chambers is six or eight.

5. The multi-tube engine system according to claim 3, characterized in that: All thrust chambers can swing in both directions, with the central thrust chamber having a swing radius of 5° and the outer thrust chamber having a swing radius of 2.5°.

6. A control method for a multi-tube engine suitable for reusable rockets, comprising the multi-tube engine system for reusable rockets as described in any one of claims 1 to 5, characterized in that: It also includes the following steps: S10. The gas turbine of the turbopump, driven by the rocket engine start-up module, starts to rotate and perform work. The fuel pump and oxidizer pump start to rotate with the gas turbine. The pressure in the pipeline after the pump increases. The oxidizer auxiliary valve and fuel auxiliary valve of the first gas generator open, and the propellant enters the first gas generator. The first gas generator ignites. The high-temperature fuel-rich gas generated after the first gas generator ignites enters the gas turbine to drive the turbopump to work continuously. The rocket engine start-up module is shut down. The pressure after the pump rises to the required pressure. According to the timing determined by the test, fuel and oxidizer are provided to each thrust chamber. Each thrust chamber ignites and starts. After successful ignition, as the pressure after the turbopump gradually increases, the turbopump power increases. The system enters the main stage working state, and the ground start-up process ends. S20. After the rocket propels the payload to the predetermined altitude, the first-stage engine shuts down, the first-stage rocket separates from the rocket body and enters the recovery procedure; after the system issues a shutdown command, the first gas generator stops working, the turbopump power decreases, the fuel and oxidizer pipelines of each thrust chamber are shut down, the engine thrust decreases to zero, and the shutdown process ends. After the first-stage rocket separates from the rocket body, it enters the recovery phase. Upon receiving the start command, the rocket engine start module drives the gas turbine of the turbopump to start working. The fuel pump and oxidizer pump start with the gas turbine, increasing the pressure in the downstream pipeline. The oxidizer auxiliary valve and fuel auxiliary valve of the second gas generator open, allowing propellant to enter the second gas generator. The second gas generator ignites, and the high-temperature, fuel-rich gas generated after ignition enters the gas turbine, driving the turbopump to work continuously. The rocket engine start module is then shut down. Once the downstream pressure reaches the required level, fuel and oxidizer are supplied to the thrust chamber according to the experimentally determined sequence. The thrust chamber ignites and starts. After successful ignition, as the downstream pressure of the turbopump gradually increases, the turbopump power increases, and the system enters the main stage operating state, ending the in-flight start-up process.

7. The control method for a multi-tube engine according to claim 6, characterized in that: It also includes the following steps: S11. During the rocket's ascent phase, when thrust adjustment is required, the oxidizer flow rate of the first gas generator is adjusted by the control valve, thereby changing the propellant mixing ratio and flow rate of the first gas generator, which in turn changes the chamber pressure and gas temperature of the first gas generator, changes the energy of the working gas input to the gas turbine, changes the output power of the gas turbine, and consequently changes the output power of the oxidizer pump and fuel pump. S31. During the rocket recovery phase, when thrust adjustment is required, the oxidizer flow rate of the second gas generator is adjusted by controlling the valve, thereby changing the propellant mixing ratio and flow rate of the second gas generator, which in turn changes the chamber pressure and gas temperature of the second gas generator, changes the energy of the working gas input to the gas turbine, changes the output power of the gas turbine, and consequently changes the output power of the oxidizer pump and fuel pump.

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