Multi-tube engine system suitable for recyclable rocket and control method of multi-tube engine system

Abstract

The invention relates to a multi-tube engine system suitable for a recoverable and reusable rocket and a control method of the multi-tube engine system. A turbine pump comprises a gas turbine, a fuel pump and an oxidizing agent pump which coaxially rotate, and a first gas generator for driving the gas turbine to rotate is connected to the gas turbine; the system further comprises a plurality of thrust chambers connected in parallel, a fuel outlet in the fuel pump is connected with a fuel path annular flow equalizer, an oxidant outlet in the oxidant pump is connected with an oxidant path annular flow equalizer, all the thrust chambers communicate with the fuel path annular flow equalizer, and all the thrust chambers communicate with the oxidant path annular flow equalizer. The gas turbine is further connected with a second gas generator driving the gas turbine to rotate. According to the multi-pipe engine system, the structure is simple, and the structural mass of the overall propellant conveying pipeline is reduced; all components can participate in test run and service life assessment tests, the test coverage requirement is completely met, heaven and earth consistency is achieved, the process test run period is short, the number of times is small, and the cost is low.

Description

Technical Field

[0001] The present invention relates to liquid rocket engine technology, in particular to a multi-tube engine system applicable to reusable rockets and a control method therefor. Background Art

[0002] With the rise and rapid development of commercial spaceflight, the competition in the commercial launch market has become increasingly fierce. As the commercial space company with the largest market share in the world at present, the "Falcon 9" rocket designed and manufactured by SpaceX has continuously refreshed its own new reuse records and consecutive recovery success records. Facts have proved that the reuse of rockets has greatly reduced the launch cost, gained a cost advantage, and won the international market. Its nine-engine parallel layout has strong fault tolerance, allowing one engine to fail in operation among the nine engines, which has improved the reliability of the multi-stage parallel engines. Other commercial space companies at home and abroad basically follow this strategy, with seven or nine gas generator cycle engines in parallel. Generally speaking, this strategy has obvious advantages. First, the technology maturity is high, the development difficulty is moderate, and the single-engine performance is improved through gradual iterative upgrades. The development and test of a single engine to improve the thrust-to-weight ratio have a small single-engine test run cost and require a small test bench test run capacity. Second, by having nine engines in parallel, while improving the rocket's carrying capacity, the reliability of the engine system is enhanced. Finally, through recovery 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 gas generator cycle engines in parallel. Generally speaking, this strategy has obvious advantages. First, the technology maturity is high, the development difficulty is moderate, and the single-engine performance is improved through gradual iterative upgrades. The development and test of a single engine to improve the thrust-to-weight ratio have a small single-engine test run cost and require a small test bench test run capacity. Second, by having nine engines in parallel, while improving the rocket's carrying capacity, the reliability of the engine system is enhanced. Finally, through recovery 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, there are still many drawbacks and deficiencies that cannot be ignored in the form of multi-engine parallel connection. In a multi-engine parallel connection engine, each engine needs to undergo a process test run. It takes a relatively long cycle to complete the process test run of the entire engine, and the total test run cost is high. Due to the high cost of the power system test run, full power system test runs are rarely conducted, resulting in insufficient assessment and inability to fully cover the mission profile, which is a very serious problem for reusable engines. For each engine in the multi-engine parallel connection scheme, the propellant inlet needs to be separately connected to the overall storage tank through pipelines, and multiple accumulators need to be set up to suppress longitudinal low-frequency coupled vibration. The more engines are connected in parallel, the more complex the supply pipelines are, the greater the structural mass is, the larger the occupied axial dimension is, and the overall axial space is encroached upon; each thrust chamber is equipped with a turbopump, with a large structural mass and a large overall outline dimension for multi-engine parallel connection. There are as many start ignition systems and subsystems as there are engines, with more valves and assembly components and a relatively large structural mass. Summary of the Invention

[0004] The purpose of the present invention is to provide a multi-tube engine system applicable to reusable rockets and its control method, which is used to solve the problems of complex systems, high test run costs, etc. existing in the parallel connection of multiple engines of existing reusable rockets.

[0005] To solve the above problems, the present invention provides a multi-tube engine system applicable to reusable rockets, including a turbopump. The turbopump includes a gas turbine, a fuel pump, and an oxidizer pump that rotate coaxially and synchronously. A first gas generator (main gas generator) for driving the gas turbine to rotate is connected to the gas turbine. The fuel inlet on the fuel pump is communicated with a fuel tank, and the oxidizer inlet on the oxidizer pump is communicated with an oxidizer tank. The system also includes multiple parallel thrust chambers. A fuel path annular flow equalizer is connected to the fuel outlet on the fuel pump, and an oxidizer path annular flow equalizer is connected to the oxidizer outlet on the oxidizer pump. The fuel inlets of each thrust chamber are communicated with the fuel path annular flow equalizer through a first fuel pipe, and the oxidizer inlets of each thrust chamber are communicated with the oxidizer path annular flow equalizer through a first oxidizer pipe; A second gas generator (auxiliary gas generator) for driving the gas turbine to rotate is also connected to the gas turbine. 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 acceleration and ascent stage, the first gas generator drives the gas turbine to rotate, and during the rocket recovery stage, the second gas generator drives the gas turbine to rotate.

[0006] The multi-tube engine system applicable to reusable rockets provided by the present invention also has the following technical features: Further, the gas turbine is also connected with a starting gas cylinder through a pipeline, and a starting control valve is further provided on the connecting pipeline between the starting gas cylinder and the gas turbine; the fuel inlet of the first gas generator is communicated with the fuel outlet of the fuel pump through a pipeline, and a first gas generator fuel sub-valve (main gas generator fuel sub-valve) is provided on the connecting pipeline; the oxidant inlet of the first gas generator is communicated with the oxidant outlet of the oxidant pump through a pipeline, and a first gas generator oxidant sub-valve (main gas generator oxidant sub-valve) is provided on the connecting pipeline; the fuel inlet of the second gas generator is communicated with the fuel outlet of the fuel pump through a pipeline, and a second gas generator fuel sub-valve (auxiliary gas generator fuel sub-valve) is provided on the connecting pipeline; the oxidant inlet of the second gas generator is communicated with the oxidant outlet of the oxidant pump through a pipeline, and a second gas generator oxidant sub-valve (auxiliary gas generator oxidant sub-valve) is provided on the connecting pipeline.

[0007] Further, a thrust chamber fuel path electric butterfly valve and a thrust chamber fuel main valve are sequentially provided on the first fuel pipe from the fuel path annular flow equalizer to the thrust chamber; a thrust chamber oxidant path electric butterfly valve and a thrust chamber oxidant main valve are sequentially provided on the first oxidant pipe from the oxidant path annular flow equalizer to the thrust chamber.

[0008] Further, a thrust chamber fuel path swing hose is further provided between the thrust chamber fuel main valve on the first fuel pipe and the thrust chamber; a thrust chamber oxidant path swing hose is further provided between the thrust chamber oxidant main valve on the first oxidant pipe and the thrust chamber.

[0009] Further, the multiple parallel thrust chambers include a central thrust chamber and multiple outer ring thrust chambers, and the outer ring thrust chambers are circumferentially and evenly arranged around the central thrust chamber.

[0010] Further, the number of the outer ring thrust chambers is six or eight.

[0011] Further, each of 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 chambers is 2.5°.

[0012] The present invention also provides a control method for a multi-tube engine applicable to a reusable rocket, based on the above multi-tube engine system applicable to a reusable rocket, including the following steps: S10. The gas turbine of the turbopump is driven by the rocket engine start module to rotate and do work. The fuel pump and the oxidizer pump rotate with the gas turbine. The pressure in the pipeline behind the pumps rises. The first gas generator oxidizer sub-valve (main gas generator oxidizer sub-valve) and the first gas generator fuel sub-valve (main gas generator fuel sub-valve) of the first gas generator open. The propellant enters the first gas generator (main gas generator). The first gas generator is ignited. The high-temperature fuel-rich gas generated after the ignition of the first gas generator enters the gas turbine to drive the turbopump to work continuously. The rocket engine start module is turned off. The pressure behind the pumps rises to the required pressure. Fuel and oxidizer are provided to each thrust chamber according to the timing determined by the test. Each thrust chamber is ignited and started. After successful ignition, as the pressure behind the turbopump gradually increases, the power of the turbopump increases, and the system enters the main stage working state, and the ground start process ends. S20. After the rocket propels the payload to the predetermined height, the first-stage engine shuts down. The first-stage rocket separates from the rocket body and enters the recovery procedure. After the system issues the shutdown command, the first gas generator stops working, the power of the turbopump decays, the fuel and oxidizer pipelines of each thrust chamber are closed, and the thrust of the engine decays to zero, and the shutdown process ends. S30. After the first-stage rocket separates from the rocket body, the first-stage rocket enters the recovery stage. After the system issues the start command, the gas turbine of the turbopump is driven by the rocket engine start module to rotate and do work. The fuel pump and the oxidizer pump rotate with the gas turbine. The pressure in the pipeline behind the pumps rises. The second gas generator oxidizer sub-valve (auxiliary gas generator oxidizer sub-valve) and the second gas generator fuel sub-valve (auxiliary gas generator fuel sub-valve) of the second gas generator open. The propellant enters the second gas generator (auxiliary gas generator). The second gas generator is ignited. The high-temperature fuel-rich gas generated after the ignition of the second gas generator enters the gas turbine to drive the turbopump to work continuously. The rocket engine start module is turned off. The pressure behind the pumps rises to the required pressure. Fuel and oxidizer are provided to the thrust chamber according to the timing determined by the test. The thrust chamber is ignited and started. After successful ignition, as the pressure behind the turbopump gradually increases, the power of the turbopump increases, and the system enters the main stage working state, and the in-air start process ends.

[0013] Further, the following steps are also included: S11. During the ascending stage of the rocket, when thrust adjustment is required, the oxidizer flow rate of the first gas generator is adjusted through the control valve, the propellant mixing ratio and flow rate of the first gas generator are changed, the chamber pressure and gas temperature of the first gas generator are changed, the energy of the working gas input to the gas turbine is changed, the output power of the gas turbine is changed, and the output powers of the oxidizer pump and the fuel pump are changed accordingly. S31. During the rocket recovery phase, when thrust adjustment is required, the oxidizer flow rate of the second gas generator is adjusted through a control valve to change the propellant mixing ratio and flow rate of the second gas generator, thereby changing the chamber pressure and gas temperature of the second gas generator, changing the energy of the working gas input to the gas turbine, changing the output power of the gas turbine, and consequently changing the output powers of the oxidizer pump and the fuel pump.

[0014] Furthermore, during the rocket recovery phase, shutdown and in-air startup can be performed multiple times as needed.

[0015] The present invention has the following beneficial effects: The multi-tube engine system and control method applicable to reusable rockets of the present application. Multiple juxtaposed rocket engine thrust chambers form a multi-tube engine. The multiple thrust chambers are pumped with fuel and oxidizer by the same turbopump, with a simple structure, reducing the overall mass of the propellant delivery pipeline structure. The turbopump is relatively far from the thrust chamber, and the pipelines between the two have flexible connection structures, which can reduce the coupled vibration between the two vibration sources of the thrust chamber and the turbopump, improve the mechanical environment of each component, and enhance the structural reliability of the engine, laying a solid foundation for rocket reuse. All components of this multi-tube engine system can participate in the test run and life assessment test, fully meeting the test coverage requirements and achieving space-ground consistency. It can conduct a full-system test run or a test run with some thrust units working during rocket return, which can fully assess the coordination of the operation of this multi-tube engine system and greatly reduce the possibility of flight failures. When the first-stage rocket adopts this multi-tube engine system, only one process test run is required, with a short process test run cycle, few times, and low cost. BRIEF DESCRIPTION OF THE DRAWINGS

[0016] Figure 1 is a system schematic diagram of the multi-tube engine system applicable to reusable rockets according to an embodiment of the present invention; Figure 2 is a schematic diagram of the thrust chamber arrangement method in an embodiment of the present invention; Figure 3 is a specific structural schematic diagram of the multi-tube engine system applicable to reusable rockets according to an embodiment of the present invention. DETAILED DESCRIPTION OF THE EMBODIMENTS

[0017] The present invention will be described in detail below with reference to the drawings and in conjunction with embodiments. It should be noted that, without conflict, the embodiments in the present invention and the features in the embodiments can be combined with each other.

[0018] As Figures 1 to 2In the embodiment of the multi-tube engine system and its control method applicable to recyclable rockets shown, the multi-tube engine system applicable to recyclable rockets includes a turbo pump, and the turbo pump includes a gas turbine 3, a fuel pump 5, and an oxidizer pump 4 that rotate synchronously. A first gas generator 7 (main gas generator) for driving the gas turbine 3 to rotate is connected to the gas turbine 3. The fuel inlet on the fuel pump 5 is communicated with the fuel tank, and the oxidizer inlet on the oxidizer pump 4 is communicated with the oxidizer tank. It is characterized in that it further includes a plurality of parallel thrust chambers 100. A fuel path annular flow equalizer 8 is connected to the fuel outlet on the fuel pump 5, and an oxidizer path annular flow equalizer 9 is connected to the oxidizer outlet on the oxidizer pump 4. The fuel inlet of each thrust chamber 100 is communicated with the fuel path annular flow equalizer 8 through a first fuel pipe, and the oxidizer inlet of each thrust chamber 100 is communicated with the oxidizer path annular flow equalizer 9 through a first oxidizer pipe. A second gas generator 6 (auxiliary gas generator) for driving the gas turbine 3 to rotate is also connected to 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. Preferably, the rated power of the second gas generator 6 is 1 / 3 of the rated power of the first gas generator 7. During the rocket launch acceleration and ascent stage, the gas turbine 3 is driven to rotate by the first gas generator 7. During the rocket recovery stage, the gas turbine 3 is driven to rotate by the second gas generator 6.

[0019] The multi-tube engine system applicable to reusable rockets in this application. Multiple juxtaposed rocket engine thrust chambers form a multi-tube engine. The multiple thrust chambers are pumped with fuel and oxidizer by the same turbopump. Regardless of the number of thrust chambers, only two main propellant pipelines are required for the storage tank and the engine. Only one accumulator is needed at the oxidizer inlet. This not only simplifies the structure but also reduces the structural mass of the overall propellant delivery pipeline; the mechanical environment of each component is good. The turbopump and the thrust chamber are relatively far apart, and the pipelines between them have flexible connection structures, which can reduce the coupled vibration between the two vibration sources of the thrust chamber and the turbopump, improve the mechanical environment of each component, and enhance the structural reliability of the engine, laying a solid foundation for rocket reuse; the degree of integration is high. Only one set of starting, control, and purging systems is required, only one high-power turbopump, one set of subsystems, and one set of propellant supply pipelines, making the engine have a small structural mass and a high thrust-to-weight ratio; the system is simple and has strong fault tolerance. It allows one thrust unit to fail and isolates the failed thrust unit, and the overall reliability is higher than that of a multi-engine parallel engine; like a pump-back-swing engine, only the thrust chamber swings. Compared with a pump-front-swing, the swing torque required for attitude control is small, reducing the requirement for the load capacity of the servo mechanism; using a multi-tube engine with multiple thrust chambers in parallel makes the layout more compact, and the surface thrust of its rocket is relatively high, especially the surface thrust of a seven-nozzle engine is greater; like a multi-engine parallel engine, each thrust unit can be started and stopped independently to meet the requirements of a reusable rocket, and it can also shut down in groups under variable conditions to reduce the water hammer pressure during shutdown; all components of this multi-tube engine system can participate in the commissioning test and life assessment test, fully meeting the test coverage requirements and achieving space-ground consistency; it can conduct a full-system commissioning test or a commissioning test with some thrust units working during rocket return, which can fully assess the coordination of the operation of this multi-tube engine system and greatly reduce the possibility of flight failures; when the first-stage rocket uses this multi-tube engine system, only one process commissioning test is required, and the process commissioning test has a short cycle, few times, and low cost.

[0020] The multi-tube engine system applicable to reusable rockets in this application is provided with two gas generators for driving the gas turbine 3 to rotate. The first gas generator 7 serves as the main gas generator to drive the gas turbine 3 to rotate during the rocket launch acceleration and ascent stage, and the second gas generator 6 serves as the auxiliary gas generator to drive the gas turbine 3 to rotate during the rocket recovery stage, which can avoid the combustion deterioration of the gas generator during the rocket recovery stage caused by the large changes in operating conditions during the rocket launch acceleration and ascent stage and the rocket recovery stage. Specifically, taking the parallel connection of 9 thrust chambers as an example, when launching a payload, all 9 thrust chambers are working. It is allowed that in the case of one thrust unit failing and shutting down, it does not affect the overall normal operation of the engine and does not affect the completion of the payload mission; during the rocket launch acceleration and ascent stage, all 9 thrust chambers are working, and the turbopump adjusts its operating conditions as needed. The operating conditions in this stage do not change much, and the variable operating condition requirements can be met by changing the oxidizer flow rate of the gas generator; during the rocket recovery stage, the required engine thrust is very small, only about 1 / 3 to 1 / 9 of the full operating condition. Only 1 to 3 thrust chambers need to work during the return process; the operating condition of the turbopump is only 11% to 33% of the rated operating condition, seriously deviating from the design operating condition, resulting in a decrease in the efficiency of the turbopump, but this is also acceptable in engineering; if there is only one gas generator, there is a risk of combustion deterioration when the gas generator operates at 11% of the operating condition during the rocket recovery stage; in this application, by providing two gas generators for driving the gas turbine to rotate, during the rocket recovery stage, even if only one thrust chamber is working, the lowest operating condition of the second gas generator is 33% of the rated operating condition, greatly improving the working conditions of the gas generator; the variable operating condition of the thrust chamber can adjust the oxygen flow rate of the second gas generator as needed to change the output power of the gas turbine.

[0021] In an embodiment of this application, preferably, the gas turbine 3 is also connected with a starting gas cylinder 1 through a pipeline, and a starting control valve 2 is also provided on the connecting pipeline between the starting gas cylinder 1 and the gas turbine 3; the fuel inlet of the first gas generator 7 is connected with the fuel outlet of the fuel pump 5 through a pipeline, and a first gas generator fuel sub-valve 72 (main gas generator fuel sub-valve) is provided on the connecting pipeline; the oxidizer inlet of the first gas generator 7 is connected with the oxidizer outlet of the oxidizer pump 4 through a pipeline, and a first gas generator oxidizer sub-valve 71 (main gas generator oxidizer sub-valve) is provided on the connecting pipeline; the fuel inlet of the second gas generator 6 is connected with the fuel outlet of the fuel pump 5 through a pipeline, and a second gas generator fuel sub-valve 61 (auxiliary gas generator fuel sub-valve) is provided on the connecting pipeline; the oxidizer inlet of the second gas generator 6 is connected with the oxidizer outlet of the oxidizer pump 5 through a pipeline, and a second gas generator oxidizer sub-valve 62 (auxiliary gas generator oxidizer sub-valve) is provided on the connecting pipeline.

[0022] The multi-tube engine system applicable to reusable rockets in this application can select three starting schemes, namely gas cylinder starting, pyrotechnic starter, and self-starting, during the full-thrust start on the ground. As described in this application Figure 1 The embodiment shown is gas cylinder starting. Specifically, during the ground starting preparation stage, the high-pressure helium gas in the starting gas cylinder 1 is charged. After completing other starting preparation work, the engine control system issues a starting command, enters the starting timing sequence, the starting control valve 2 opens, and the high-pressure gas in the starting gas cylinder 1 enters the gas turbine 3. After expanding through the turbine nozzle, it drives the gas 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 3 drives the oxidizer pump 4 and the fuel pump 5 to start rotating, the pressure in the pipeline behind the pumps increases. The first gas generator oxidizer sub-valve 71 (main gas generator oxidizer sub-valve) and the first gas generator fuel sub-valve 72 (main gas generator fuel sub-valve) of the first gas generator 7 open, and the 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 ignition enters the gas turbine 3 to drive the turbopump to work continuously. The starting control valve 2 is closed, the pressure behind the pumps rises to the required pressure, and fuel and oxidizer are provided to each thrust chamber according to the timing sequence determined by the test. Each thrust chamber is ignited and started. After successful ignition, as the pressure behind the turbopump gradually increases, the power of the turbopump increases, and the system enters the main stage working state, and the ground starting process ends.

[0023] The above-mentioned multi-tube engine system applicable to reusable rockets in this application adopts an open gas generator cycle. The gas generator works under fuel-rich conditions. Most of the propellant enters the thrust chamber through the fuel pump and the oxidizer pump, burns to generate high-temperature and high-pressure gas, which is accelerated by the nozzle and ejected at high speed to generate thrust; a small amount of propellant enters the gas generator to burn to generate high-temperature and high-pressure fuel-rich gas, and the fuel-rich gas enters the turbopump to do work, replacing the cold gas to drive the turbine. The fuel-rich gas after doing work enters the heat exchanger and is discharged through the exhaust pipe; there is also a small amount of liquid oxygen that enters the heat exchanger and turns into oxygen to pressurize the oxygen storage tank.

[0024] In an embodiment of this application, preferably, an electric butterfly valve 102 for the fuel path of the thrust chamber and a main valve 104 for the fuel of the thrust chamber are sequentially provided on each of the first fuel pipes from the fuel path annular flow distributor 8 to the thrust chamber 100; an electric butterfly valve 101 for the oxidizer path of the thrust chamber and a main valve 103 for the oxidizer of the thrust chamber are sequentially provided on each of the first oxidizer pipes from the oxidizer path annular flow distributor 9 to the thrust chamber 100. In the multi-tube engine system applicable to reusable rockets in this embodiment, each thrust chamber 100 can be individually controlled by the electric butterfly valve 101 for the oxidizer path of the thrust chamber, the main valve 103 for the oxidizer of the thrust chamber, the electric butterfly valve 102 for the fuel path of the thrust chamber, and the main valve 104 for the fuel of the thrust chamber, improving the reliability of the system.

[0025] Specifically, during the full-thrust start-up phase of ground launch, after the gas turbine 3 starts and drives the turbopump to continuously operate, the pump-out pressures of the fuel pump 5 and the oxidizer pump 4 rise to the required pressures. According to the timing determined by the test, the electric butterfly valves 101 of the oxidizer circuit of each thrust chamber 100 and the electric butterfly valves 102 of the fuel circuit of the thrust chamber are opened. After the pressure is boosted to the specified pressure, the main oxidizer valves 103 of each thrust chamber 100 and the main fuel valves 104 of the thrust chamber automatically open and are in a self-sustaining state. After the propellant enters each thrust chamber, the thrust chamber ignites. After successful ignition, as the pump-out pressure gradually increases, the power of the turbopump increases, and the engine enters the main stage working state, and the ground start-up process ends.

[0026] In an embodiment of the present application, preferably, a swing hose 106 for the fuel circuit of the thrust chamber is further provided between the main fuel valve 104 of the thrust chamber on the first fuel pipe and the thrust chamber 100, and a swing hose 105 for the oxidizer circuit of the thrust chamber is further provided between the main oxidizer valve 103 of the thrust chamber on the first oxidizer pipe and the thrust chamber 100; so that the thrust chamber can swing within a certain range as required; like a pump-back swing engine, only the thrust chamber swings. Compared with a pump-front swing, the swing torque required for attitude control is small, reducing the requirement for the load capacity of the servo mechanism.

[0027] In an embodiment of the present application, preferably, as Figure 2 shown, multiple parallel thrust chambers 100 include a central thrust chamber 110 and multiple outer-ring thrust chambers 120. The outer-ring thrust chambers 120 are circumferentially and evenly arranged around the central thrust chamber 110. Preferably, the number of the outer-ring thrust chambers 120 is six or eight. Taking the number of the outer-ring thrust chambers 120 being eight as an example, the eight outer-ring thrust chambers 120 are sequentially numbered I, II, III, IV, V, VI, VII, VIII, and the central thrust chamber 110 is numbered IX. Using such a multi-tube engine with multiple parallel thrust chambers makes the layout more compact, and the surface thrust of the rocket is relatively high.

[0028] In an embodiment of the present application, preferably, each thrust chamber 100 can swing bidirectionally. The swing circle of the central thrust chamber 110 is 5°, and the swing circle of the outer-ring thrust chambers 120 is 2.5°. Specifically, there are a total of nine or seven thrust units, one in the center and eight or six in the outer circle circumferentially distributed; each thrust unit has a bidirectional swing function. The swing circle of the outer-ring engine is 2.5°, and the swing circle of the central engine is 5°. The roll control is the same tangential swing, and the swing directions of the pitch, yaw or composite control are the same; the swing angles of each thrust chamber can be flexibly controlled as required, and the attitude of the rocket can be flexibly controlled and adjusted.

[0029] The present application also provides a control method for a multi-tube engine applicable to a reusable rocket, including the above multi-tube engine system applicable to a reusable rocket, and further including the following steps: S10. The gas turbine of the turbopump is rotated by the rocket engine start module to drive the gas turbine blades to do work. The fuel pump and the oxidizer pump rotate with the gas turbine. The pressure in the pipeline behind the pumps rises. The first gas generator oxidizer sub-valve 71 (main gas generator oxidizer 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-burn gas generated after the ignition of the first gas generator enters the gas turbine 3 to drive the turbopump to work continuously. The rocket engine start module is closed. The pressure behind the pumps rises to the required pressure. Fuel and oxidizer are provided for each thrust chamber according to the timing determined by the test. Each thrust chamber is ignited and started. After successful ignition, as the pressure behind the turbopump gradually increases, the power of the turbopump increases, and the system enters the main stage working state. The ground start process ends. S20. After the rocket propels the payload to the predetermined height, 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 power of the turbopump decays, the fuel and oxidizer pipelines of each thrust chamber are closed, and the thrust of the engine decays to zero. The shutdown process ends. S30. After the first-stage rocket separates from the rocket body, the first-stage rocket enters the recovery stage. After the system issues a start command, the gas turbine of the turbopump is rotated by the rocket engine start module to drive the gas turbine blades to do work. The fuel pump and the oxidizer pump rotate with the gas turbine. The pressure in the pipeline behind the pumps rises. The second gas generator oxidizer sub-valve 61 (auxiliary gas generator oxidizer sub-valve) and the second gas generator fuel sub-valve 62 (auxiliary gas generator fuel sub-valve) of the second gas generator 6 are opened. The propellant enters the second gas generator 6 (auxiliary gas generator). The second gas generator 6 is ignited. The high-temperature rich-burn gas generated after the ignition of the second gas generator enters the gas turbine 3 to drive the turbopump to work continuously. The rocket engine start module is closed. The pressure behind the pumps rises to the required pressure. Fuel and oxidizer are provided for the thrust chamber according to the timing determined by the test. The thrust chamber is ignited and started. After successful ignition, as the pressure behind the turbopump gradually increases, the power of the turbopump increases, and the system enters the main stage working state. The in-air start process ends.

[0030] In an embodiment of the present application, preferably, it further includes the step of: S11. During the ascending stage of the rocket, when thrust adjustment is required, the oxidizer flow rate of the first gas generator is adjusted through a control valve to change the propellant mixing ratio and flow rate of the first gas generator, so as to change the chamber pressure and gas temperature of the first gas generator, change the energy of the working gas input to the gas turbine, change the output power of the gas turbine, and the output powers of the oxidizer pump and the fuel pump are changed accordingly. S31. During the rocket recovery phase, when thrust adjustment is required, the oxidizer flow rate of the second gas generator is adjusted through a control valve to change the propellant mixing ratio and flow rate of the second gas generator, thereby changing the chamber pressure and gas temperature of the second gas generator, changing the energy of the working gas input to the gas turbine, changing the output power of the gas turbine, and consequently changing the output powers of the oxidizer pump and the fuel pump.

[0031] In an embodiment of the present application, preferably, during the rocket recovery phase, shutdown and in-air startup can be performed multiple times as needed.

[0032] The multi-tube engine system applicable to reusable rockets and its control method of the present application will be described below by taking nine thrust chambers as an example.

[0033] As Figure 1 shown, the cross-sections where the three thrust chambers of VII-IX-III are located are schematically given. The layout of the nine thrust chambers is as Figure 2 shown. Each thrust chamber is connected to the fuel path annular flow equalizer 8 and the oxidizer path annular flow equalizer 9 through pipelines.

[0034] Working principle: The multi-tube engine adopts an open-cycle gas generator. The gas generator operates under rich-burn conditions. Most of the propellants enter the thrust chamber through the fuel and oxidizer pumps, burn to generate high-temperature and high-pressure gas, which is accelerated by the nozzle and ejected at high speed to generate thrust; a small amount of propellants enter the gas generator to burn to generate high-temperature and high-pressure rich-burn gas. The rich-burn gas enters the turbopump to do work, replacing the cold gas to drive the turbine. After doing work, the rich-burn gas enters the heat exchanger and is then discharged through the exhaust pipe; a small amount of liquid oxygen enters the heat exchanger to become oxygen and pressurizes the oxygen storage tank; for rocket attitude control, only the thrust chamber needs to be swung, and the swing torque is very small. An electric servo mechanism can meet the requirements.

[0035] Thrust unit layout: There are a total of nine or seven thrust units, one in the center and eight or six evenly distributed in a circle on the outer ring; each thrust unit has a two-way swing function; the outer ring engines swing in a circle by 2.5°, and the central engine swings in a circle by 5°. The roll control is the same tangential swing, and the swing directions for pitch, yaw, or compound control are the same.

[0036] There are two recovery modes, off-site recovery and on-site recovery; during off-site recovery, the maximum number of engine operations is 3 times. The central thrust unit can operate 3 times, and the remaining thrust units have the ability to operate twice; during on-site recovery, the number of engine operations is 4 times. The central thrust unit can operate 4 times, and the remaining thrust units have the ability to operate 2 times.

[0037] When launching the payload, all 9 thrust chambers are operating. It is allowed that in the case of one thrust unit failing and shutting down, it will not affect the overall normal operation of the engine and the completion of the payload mission. During the rocket launch and acceleration phase, all the thrust chambers in parallel of the first-stage power system are operating, and the turbopump adjusts its operating conditions as needed. The operating conditions in this phase do not change much, and the requirements for variable operating conditions can be met by changing the oxidizer flow rate of the gas generator.

[0038] During the recovery phase, the required engine thrust is very small, only about 1 / 3 to 1 / 9 of the full operating conditions. Only 1 to 3 thrust chambers need to operate during the return process. The operating conditions of the turbopump are only 11% to 33% of the rated operating conditions, seriously deviating from the design operating conditions, resulting in a reduction in the efficiency of the turbopump, but this is also acceptable in engineering. There will be certain risks when the gas generator operates at 11% of the operating conditions - combustion deterioration. For this reason, two gas generators are set up, one is the main gas generator and the other is the auxiliary gas generator. The main gas generator operates during the payload launch phase, and the auxiliary gas generator operates during the recovery phase. During the recovery process, even if only one thrust unit is operating, the minimum operating condition of the generator is 33% of the rated operating condition, improving the operating conditions of the generator; the engine variable operating conditions can adjust the oxygen flow rate of the auxiliary generator as needed to change the turbine output power.

[0039] The system that supplies the propellant to the thrust chamber starting from the engine propellant inlet under specified operating conditions is called the main path system, including the turbopump, the electric butterfly valve for the oxidizer path, the fuel electric butterfly valve, the oxidizer main valve, the fuel main valve, the metal bellows, etc. The metal bellows are set to meet the requirements of the double pendulum (swing circle) of the thrust chamber. The oxidizer main valve and the fuel main valve have the function of fluid on-off. These two main valves are normally closed valves and open and maintain the open state under pressure. The oxidizer main valve also has a liquid oxygen precooling reflux channel, and the fuel main valve has a discharge channel. When a certain thrust chamber has a fault that affects the operation of the power system or when a certain thrust unit does not need to operate according to requirements, the electric butterfly valve is closed to forcibly cut off the supply of the propellant corresponding to the thrust unit, and the remaining thrust chambers can operate normally, improving the fault tolerance of the power system and expanding the engine thrust adjustment range. In addition, the set electric butterfly valve can also change the mixture ratio of the thrust chamber and has the function of the propellant utilization system.

[0040] The start-up can adopt solid starter start-up, gas starter start-up, liquid starter and self-starting, without changing the main components and control methods of the system. The ignition methods of the thrust chamber and the gas generator can select methods such as spark plug ignition, pyrotechnic ignition and chemical ignition. The differences in ignition methods do not affect the main components and control methods of the system. The thrust unit can be nine or seven thrust units, with one in the center and eight or six evenly distributed in a circle on the outer ring. Or further, more than nine thrust units can be selected according to needs. The differences in the number and layout of the thrust chambers do not affect the main components and control methods of the system.

[0041] Working process: 1) Full-thrust start for ground launch: The engine system can select three start-up schemes: gas cylinder start-up, pyrotechnic starter, and self-start. The system diagram given in this application takes gas cylinder start-up as an example. As Figure 1 shown, specifically, during the ground start-up preparation stage, the high-pressure helium gas in the start-up gas cylinder 1 is charged. After completing other start-up preparation work, the engine control system issues a start command and enters the start timing sequence. The start 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 gas 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 rotating, the pressure in the pipeline behind the propellant pump increases. The oxidizer sub-valve 71 and the fuel sub-valve 72 of the first gas generator 7 (main gas generator) open, and the propellant enters the first gas generator 7 (main gas generator). The first gas generator 7 ignites, and the high-temperature rich-burn gas generated by ignition enters the gas turbine, continuing to drive the turbopump to work continuously. The start control valve 2 is closed, and the pressure behind the pump rises to the required pressure. According to the timing sequence determined by the test, the oxidizer line motorized butterfly valves 101 and the fuel line motorized butterfly valves 102 of each thrust chamber from I to IX are opened. After the pressure rises to the specified pressure, the oxidizer main valves 103 and the fuel main valves 104 of each thrust chamber automatically open and are in a self-sustaining state. After the propellant enters each thrust chamber, the thrust chamber ignites. After successful ignition, as the pressure behind the pump gradually increases, the turbopump power increases, and the engine enters the main stage working state, and the ground start-up process ends.

[0042] 2) Thrust regulation under the main stage working condition during the ground launch ascending stage: The thrust regulation under the main stage working condition during the ground launch ascending stage is achieved by changing the oxidizer flow rate of the first gas generator 7 (main gas generator). As Figure 1 shown, specifically, by changing the oxidizer flow rate of the first gas generator 7, the propellant mixing ratio of the main gas generator is changed, thereby changing the chamber pressure and gas temperature of the gas generator. In this way, 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, and the propellant flow rate flowing into the thrust chamber also changes accordingly, ultimately realizing the working condition adjustment of the entire engine main stage.

[0043] 3) Shutdown process during the ground launch ascending stage: The multi-tube engine system applicable to the reusable rocket in this application is generally used for the first stage of the reusable rocket. After the multi-tube engine system propels the payload to the predetermined height, the first stage engine shuts down, and the first stage rocket separates from the rocket body and enters the recovery program. The entire shutdown process of the engine should be able to quickly and accurately stop the engine from working, with a small aftereffect impulse deviation. As Figure 1As shown in the figure, the specific shutdown process is as follows: after the shutdown instruction is issued, the oxidizer sub-valve 71 of the first gas generator closes, followed by the fuel sub-valve 72 closing. The power of the first gas generator and the turbopump decays. After the power decays to a certain extent, the motorized butterfly valves 101 of the oxidizer circuits of the thrust chambers I - IX close. After the motorized butterfly valves of the oxidizer circuits close, the chamber pressure of the thrust chambers rapidly decreases, and the pressure behind the valves also rapidly decreases. After the pressure decreases to the designed closing pressure of the oxidizer main valve, the oxidizer main valves 103 of the thrust chambers I - IX close. Then, in sequence, the motorized butterfly valves 102 of the fuel circuits of the thrust chambers I - IX are closed. After the motorized butterfly valves of the fuel circuits close, the pressure behind the valves also rapidly decreases. After the pressure decreases to the designed closing pressure of the fuel main valve, the fuel main valves 104 of the thrust chambers I - IX close. The thrust of the engine decays to zero, and the shutdown process ends.

[0044] 4) Second / third / fourth in-air starts during the recovery phase: After the separation of the first and second stages of the rocket, the first-stage engine enters the recovery phase. During the recovery phase, the engine system only requires the second gas generator 6 to drive the turbopump to do work, which can meet the requirements for 1 - 3 thrust chambers to operate. Taking the case where only the central thrust chamber operates during return as an example, the engine control system issues a start instruction and enters the start timing sequence. The start control valve 2 opens, and the high-pressure gas in the start gas cylinder enters the gas turbine 3. After expanding through the turbine nozzles, 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 rotating, the pressure in the pipeline behind the propellant pumps increases. The oxidizer sub-valve 61 and the fuel sub-valve 62 of the second gas generator 6 (auxiliary gas generator) open, and the propellant enters the second gas generator 6. The second gas generator 6 ignites, and the high-temperature rich-burn gas generated by ignition enters the gas turbine, continuing to drive the turbopump to operate continuously. The start control valve 2 is closed, and the pressure behind the pumps increases to the required pressure. According to the timing sequence determined by the test, the motorized butterfly valves 101 of the oxidizer circuit and the motorized butterfly valves 102 of the fuel circuit of the IX thrust chamber are opened. After the pressure rises to the specified pressure, the oxidizer main valve 103 and the fuel main valve 104 of the engine 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 behind the pumps gradually increases, the power of the turbopump increases, and the engine enters the main stage operating state. The in-air start process ends, and the working processes of the second / third / fourth starts during the entire recovery phase are the same.

[0045] 5) Thrust regulation during the recovery phase: The thrust regulation in the main stage condition during the recovery phase is achieved by changing the oxidizer flow rate of the second gas generator 6, such as Figure 1As shown, specifically, by changing the oxidizer flow rate of the second gas generator 6 and changing the propellant mixing ratio of the second gas generator 6, the chamber pressure and gas temperature of the gas generator are changed. In this way, 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, and the propellant flow rate flowing into the thrust chamber also changes accordingly. Finally, the working conditions of the main stage of the entire engine are adjusted, and the main stage thrust adjustment methods during the entire recovery stage are the same.

[0046] 6) Secondary / tertiary / quaternary shutdown during the recovery stage: After the shutdown command is issued, the oxidizer sub-valve 61 of the second gas generator 6 closes, and then the fuel sub-valve 62 closes. The power of the second gas generator and the turbopump decays. After the power decays to a certain extent, the electric butterfly valve 101 of the oxidizer line of the IX thrust chamber closes. After the electric butterfly valve 101 of the oxidizer line closes, the chamber pressure of the IX thrust chamber rapidly decreases, and the pressure behind the valve also rapidly decreases. After the pressure decreases to the designed closing pressure of the oxidizer main valve, the oxidizer main valve 103 of the IX thrust chamber closes. Then, in sequence, the electric butterfly valve 102 of the fuel line of the IX thrust chamber is closed. After the electric butterfly valve of the fuel line closes, the pressure behind the valve also rapidly decreases. After the pressure decreases to the designed closing pressure of the fuel main valve, the fuel main valve 104 of the IX thrust chamber closes. The engine thrust decays to zero, and the shutdown process ends. The working processes of the secondary / tertiary / quaternary shutdown during the entire recovery stage are the same.

[0047] Finally, it should be noted that the above embodiments are only used to illustrate the technical solutions of the present invention and are not intended to limit them. Although the present invention has been described in detail with reference to the foregoing embodiments, those of ordinary skill in the art should understand that they can still modify the technical solutions described in the foregoing embodiments or perform equivalent replacements for some of the technical features. However, these modifications or replacements do not cause the essence of the corresponding technical solutions to deviate from the spirit and scope of the technical solutions of the embodiments of the present invention.

Claims

1. A multi-tube engine system applicable to a recyclable and reusable rocket, characterized in that, It includes a turbopump, and the turbopump includes a gas turbine, a fuel pump, and an oxidizer pump that rotate synchronously. A first gas generator for driving the gas turbine to rotate is connected to the gas turbine. The fuel inlet on the fuel pump communicates with a fuel tank, and the oxidizer inlet on the oxidizer pump communicates with an oxidizer tank. It is characterized in that it further includes a plurality of parallel thrust chambers. A fuel path annular flow equalizer is connected to the fuel outlet on the fuel pump, and an oxidizer path annular flow equalizer is connected to the oxidizer outlet on the oxidizer pump. The fuel inlet of each thrust chamber communicates with the fuel path annular flow equalizer through a first fuel pipe, and the oxidizer inlet of each thrust chamber communicates with the oxidizer path annular flow equalizer through a first oxidizer pipe; A second gas generator for driving the gas turbine to rotate is also connected to 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 ascending stage, the first gas generator drives the gas turbine to rotate, and during the rocket recovery stage, the second gas generator drives the gas turbine to rotate.

2. The multi-tube engine system according to claim 1, wherein: The gas turbine is also connected with a starting gas cylinder through a pipeline, and a starting control valve is also provided on the connecting pipeline between the starting gas cylinder and the gas turbine; The fuel inlet of the first gas generator communicates with the fuel outlet of the fuel pump through a pipeline, and a first gas generator fuel sub-valve is provided on the connecting pipeline. The oxidizer inlet of the first gas generator communicates with the oxidizer outlet of the oxidizer pump through a pipeline, and a first gas generator oxidizer sub-valve is provided on the connecting pipeline; The fuel inlet of the second gas generator communicates with the fuel outlet of the fuel pump through a pipeline, and a second gas generator fuel sub-valve is provided on the connecting pipeline. The oxidizer inlet of the second gas generator communicates with the oxidizer outlet of the oxidizer pump through a pipeline, and a second gas generator oxidizer sub-valve is provided on the connecting pipeline.

3. The multi-tube engine system according to claim 1, characterized in that: On the first fuel pipe, a thrust chamber fuel path electric butterfly valve, a thrust chamber fuel main valve, and a thrust chamber fuel path swing hose are sequentially provided from the fuel path annular flow equalizer to the thrust chamber; On the first oxidizer pipe, a thrust chamber oxidizer path electric butterfly valve, a thrust chamber oxidizer main valve, and a thrust chamber oxidizer path swing hose are sequentially provided from the oxidizer path annular flow equalizer to the thrust chamber.

4. The multi-tube engine system according to claim 3, characterized in that: A thrust chamber fuel path swing hose is further provided between the thrust chamber fuel main valve on the first fuel pipe and the thrust chamber, and a thrust chamber oxidizer path swing hose is further provided between the thrust chamber oxidizer main valve on the first oxidizer pipe and the thrust chamber.

5. The multi-tube engine system according to claim 1, characterized in that: 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 circumferentially and evenly arranged with the central thrust chamber as the center.

6. The multi-tube engine system according to claim 5, characterized in that: The number of the outer ring thrust chambers is six or eight.

7. The multi-tube engine system according to claim 5, characterized in that: Each of 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 chambers is 2.5°.

8. A control method for a multi-tube engine applicable to a reusable rocket, comprising the multi-tube engine system applicable to a reusable rocket according to any one of claims 1 to 7, characterized in that: It also includes the following steps: S10. The gas turbine of the turbopump is driven by the rocket engine start module to rotate and do work. The fuel pump and oxidizer pump rotate with the gas turbine. The pressure in the pipeline behind the pumps increases. The first gas generator oxidizer sub-valve and the first gas generator fuel sub-valve of the first gas generator open. The propellant enters the first gas generator. The first gas generator is ignited. The high-temperature fuel-rich gas generated after the ignition of the first gas generator enters the gas turbine to drive the turbopump to work continuously. The rocket engine start module is shut down. When the pressure behind the pumps rises to the required pressure, fuel and oxidizer are provided to each thrust chamber according to the timing determined by the test. Each thrust chamber is ignited and started. After successful ignition, as the pressure behind the turbopump gradually increases, the power of the turbopump increases, and the system enters the main stage working state, and the ground start process ends. S20. After the rocket propels the payload to the predetermined height, 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 power of the turbopump decays, the fuel and oxidizer pipelines of each thrust chamber are closed, and the thrust of the engine decays to zero, and the shutdown process ends. S30. After the first-stage rocket separates from the rocket body, the first-stage rocket enters the recovery stage. After the system issues a start command, the gas turbine of the turbopump is driven by the rocket engine start module to rotate and do work. The fuel pump and oxidizer pump rotate with the gas turbine. The pressure in the pipeline behind the pumps increases. The second gas generator oxidizer sub-valve and the second gas generator fuel sub-valve of the second gas generator open. The propellant enters the second gas generator. The second gas generator is ignited. The high-temperature fuel-rich gas generated after the ignition of the second gas generator enters the gas turbine to drive the turbopump to work continuously. The rocket engine start module is shut down. When the pressure behind the pumps rises to the required pressure, fuel and oxidizer are provided to the thrust chamber according to the timing determined by the test. The thrust chamber is ignited and started. After successful ignition, as the pressure behind the turbopump gradually increases, the power of the turbopump increases, and the system enters the main stage working state, and the in-air start process ends.

9. The control method of the multi-tube engine according to claim 8, characterized in that: It also includes the following steps: S11. During the rocket ascent stage, when thrust adjustment is required, the oxidizer flow rate of the first gas generator is adjusted through a control valve, the propellant mixing ratio and flow rate of the first gas generator are changed, the chamber pressure and gas temperature of the first gas generator are changed, the energy of the working gas input to the gas turbine is changed, the output power of the gas turbine is changed, and the output powers of the oxidizer pump and fuel pump are changed accordingly. S31. During the rocket recovery stage, when thrust adjustment is required, the oxidizer flow rate of the second gas generator is adjusted through a control valve, the propellant mixing ratio and flow rate 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 to the gas turbine is changed, the output power of the gas turbine is changed, and the output powers of the oxidizer pump and fuel pump are changed accordingly.

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