WO2018051566A1 - 電動アシスト液体燃料ロケット推進システム - Google Patents
電動アシスト液体燃料ロケット推進システム Download PDFInfo
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- WO2018051566A1 WO2018051566A1 PCT/JP2017/015074 JP2017015074W WO2018051566A1 WO 2018051566 A1 WO2018051566 A1 WO 2018051566A1 JP 2017015074 W JP2017015074 W JP 2017015074W WO 2018051566 A1 WO2018051566 A1 WO 2018051566A1
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- fuel
- combustor
- oxidant
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02K—JET-PROPULSION PLANTS
- F02K9/00—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof
- F02K9/42—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof using liquid or gaseous propellants
- F02K9/44—Feeding propellants
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02K—JET-PROPULSION PLANTS
- F02K9/00—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof
- F02K9/42—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof using liquid or gaseous propellants
- F02K9/44—Feeding propellants
- F02K9/46—Feeding propellants using pumps
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02K—JET-PROPULSION PLANTS
- F02K9/00—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof
- F02K9/42—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof using liquid or gaseous propellants
- F02K9/44—Feeding propellants
- F02K9/46—Feeding propellants using pumps
- F02K9/48—Feeding propellants using pumps driven by a gas turbine fed by propellant combustion gases or fed by vaporized propellants or other gases
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02K—JET-PROPULSION PLANTS
- F02K9/00—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof
- F02K9/42—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof using liquid or gaseous propellants
- F02K9/44—Feeding propellants
- F02K9/50—Feeding propellants using pressurised fluid to pressurise the propellants
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02K—JET-PROPULSION PLANTS
- F02K9/00—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof
- F02K9/42—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof using liquid or gaseous propellants
- F02K9/44—Feeding propellants
- F02K9/56—Control
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02K—JET-PROPULSION PLANTS
- F02K9/00—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof
- F02K9/42—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof using liquid or gaseous propellants
- F02K9/44—Feeding propellants
- F02K9/56—Control
- F02K9/563—Control of propellant feed pumps
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02K—JET-PROPULSION PLANTS
- F02K9/00—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof
- F02K9/42—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof using liquid or gaseous propellants
- F02K9/44—Feeding propellants
- F02K9/56—Control
- F02K9/58—Propellant feed valves
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02K—JET-PROPULSION PLANTS
- F02K9/00—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof
- F02K9/95—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof characterised by starting or ignition means or arrangements
Definitions
- This disclosure relates to an electrically assisted liquid fuel rocket propulsion system that uses liquid fuel and liquid oxidant as propellants.
- liquid fuel rockets are more complex than solid fuel rockets, they are known to be superior in thrust control.
- a general liquid fuel rocket uses liquid hydrogen as a fuel and liquid oxygen as an oxidant.
- Development of rockets using liquefied methane as fuel is also underway (see Patent Document 1). These so-called propellants are introduced into the rocket propulsion system as a cryogenic liquid.
- Patent Document 2 discloses a configuration for suppressing this mass consumption.
- the cooling unit through which the propellant flows is provided on the outer periphery of the feed line from the storage tank to the pump. Thereby, cooling of the feed line is promoted and consumption of the propellant is suppressed.
- the rocket propulsion system must be sufficiently cooled in advance.
- the propellant flows through the propulsion system during pre-cooling and is then discharged out of the propulsion system. That is, the propellant at this time is not used for combustion of the engine but is discarded as it is.
- the large consumption of propellant during pre-cooling results in an increase in propulsion system mass, i.e., a decrease in payload mass injected into the orbit, contributing to increased launch costs.
- the conventional propulsion system does not have the power to drive the turbo pump for boosting the propellant before the ignition of the engine. Therefore, the propellant supply depends on the heat capacity of the combustor or is supplemented with a gas bottle dedicated to starting. However, the heat capacity of the combustor is difficult to determine due to a design change or the like, and varies depending on the surrounding environment. Gas bottles increase the propulsion system mass excessively.
- an object of the present disclosure is to provide an electrically assisted liquid fuel rocket propulsion system that can suppress the consumption of propellant during pre-cooling.
- the present disclosure aims to stabilize the turbo pump drive at the time of starting and reduce the mass of the apparatus therefor.
- One aspect of the present disclosure is an electrically assisted liquid fuel rocket propulsion system, which is a mixed gas of a subcombustor that generates fuel and oxidant combustion gas, and the fuel and the combustion gas discharged from the subcombustor.
- a main combustor that combusts the gas, a turbine that is rotated by the flow of the combustion gas, and a pump that is driven by the rotation of the turbine.
- a turbo pump that supplies the oxidant from the agent tank to the sub-combustor and the main combustor, an electric motor that rotates the turbine, and a clutch that connects and disconnects the electric motor and the turbine.
- the gist is to provide.
- the electrically assisted liquid fuel rocket propulsion system may further include a return path for returning the fuel and the oxidant discharged from the turbo pump to the fuel tank and the oxidant tank, respectively.
- the electrically assisted liquid fuel rocket propulsion system may further include a cooling device that cools the fuel tank and the oxidant tank.
- the electrically assisted liquid fuel rocket propulsion system includes a first valve that adjusts a supply amount of the fuel to the sub-combustor, a second valve that adjusts a supply amount of the oxidant to the main combustor, and the sub-combustor. And a third valve that adjusts the amount of the oxidant supplied to the combustor, and the first valve, the second valve, and the third valve may all be variable flow valves. .
- the electrically assisted liquid fuel rocket propulsion system may further include a laser ignition device that ignites the sub-combustor and the main combustor using laser light.
- an electrically assisted liquid fuel rocket propulsion system that can suppress the consumption of propellant during pre-cooling.
- FIG. 1 is a schematic configuration diagram of an electrically assisted liquid fuel rocket propulsion system according to an embodiment of the present disclosure.
- FIG. 2 is a diagram showing an operation during pre-cooling of the electrically assisted liquid fuel rocket propulsion system shown in FIG.
- FIG. 3 is a schematic configuration diagram illustrating a modified example of the electrically assisted liquid fuel rocket propulsion system according to the embodiment of the present disclosure.
- FIG. 4 is a diagram showing an operation during pre-cooling of the electrically assisted liquid fuel rocket propulsion system shown in FIG.
- FIG. 1 is a schematic configuration diagram of a liquid fuel rocket propulsion system according to the present embodiment.
- the liquid fuel rocket propulsion system includes a main combustor 10, a subcombustor 11, and a turbo pump 20.
- the liquid fuel rocket propulsion system is simply referred to as “propulsion system”.
- “at the time of pre-cooling” means a period in which the propulsion system needs to be cooled in a state where the main combustor 10 and the sub-combustor 11 are not combusting. Therefore, “pre-cooling” is not limited to cooling before the rocket is launched from the ground, but also includes cooling before the main combustor 10 and the sub-combustor 11 are reburned.
- the propulsion system of the present embodiment uses liquid hydrogen or liquefied hydrocarbon (for example, liquefied methane) as fuel and liquid oxygen as an oxidant.
- the operating cycle of the propulsion system is a two-stage combustion cycle. That is, the combustion gas generated by the auxiliary combustor (preburner) 11 drives the turbo pump 20 (specifically, the turbine 21), and is then further combusted by the main combustor 10.
- Fuel is sucked from the fuel tank 30 by the turbo pump 20 and supplied to the sub-combustor 11 through the heat exchanger 12 provided in the main combustor 10. In the heat exchanger 12, the fuel cools the combustion chamber of the main combustor 10.
- Oxidant is also sucked from the oxidant tank 40 by the turbo pump 20. Most of the sucked oxidant is supplied to the main combustor 10, and the remainder is supplied to the sub-combustor 11.
- the auxiliary combustor 11 generates combustion gas of fuel and oxidant. The generated combustion gas flows through the turbine of the turbo pump 20 as a driving gas for the turbine 21 of the turbo pump 20 and is then supplied to the main combustor 10.
- the main combustor 10 burns a mixed gas of the oxidant discharged from the turbo pump 20 and the combustion gas discharged from the sub-combustor 11.
- the combustion gas is discharged from the nozzle through the throat.
- the rocket thrust is obtained by the acceleration accompanying the expansion of the combustion gas at this time.
- this embodiment can also be applied to a propulsion system for a gas generator cycle.
- the combustion gas of the subcombustor (gas generator) 11 that has passed through the turbine 21 of the turbo pump 20 is discharged to the outside.
- the turbo pump 20 supplies propellant (that is, fuel and oxidant) to the main combustor 10 and the subcombustor 11.
- the turbo pump 20 supplies fuel from the fuel tank 30 to the sub-combustor 11 and supplies oxidant from the oxidant tank 40 to the sub-combustor 11 and the main combustor 10.
- the turbo pump 20 includes a turbine 21, and a first pump (fuel pump) 22 and a second pump (oxidant pump) 23 driven by the turbine 21.
- the rotating shaft of the turbo pump 20 is connected (connected) to the rotating shaft of the electric motor 25 via the clutch 24.
- the turbine 21 is rotated by the flow of combustion gas from the sub-combustor 11 to the main combustor 10, and transmits this rotational force to the first pump 22 and the second pump 23. That is, the first pump 22 and the second pump 23 are driven by the turbine 21.
- the first pump 22 sucks the fuel in the fuel tank 30 and boosts and discharges the fuel.
- the second pump 23 sucks the oxidant in the oxidant tank 40 and boosts and discharges the oxidant.
- Each pump 22, 23 is a centrifugal pump having an impeller (wheel), for example.
- the propulsion system of this embodiment includes an electric motor 25 and a clutch 24.
- the electric motor 25 is connected to the turbine 21 of the turbo pump 20 via the clutch 24.
- the clutch 24 connects and releases the rotating shaft of the electric motor 25 and the rotating shaft of the turbine 21.
- a well-known structure can be employed for the clutch 24.
- the electric motor 25 is provided for at least the following two purposes.
- the first purpose is to maintain the cooling state of the propulsion system by recirculating a small amount of propellant with the turbo pump 20 having a low speed.
- the second purpose is to create a flow rate and pressure environment necessary for ignition in the main combustor 10 and the subcombustor 11 at the start of the propulsion system with the turbo pump 20 having a medium speed.
- the rotation speed of the turbo pump 20 increases to a region beyond the operation region of the electric motor 25.
- the clutch 24 physically (mechanically) separates the electric motor 25 from the rotating shaft of the turbo pump 20 in order to avoid the destruction of the electric power system due to the excessive electromotive force of the electric motor 25.
- the power source of the electric motor 25 can be changed (switched) according to the environment outside the rocket. For example, before the rocket launches, the electric power of the electric motor 25 is supplied from a storage battery (not shown) or an external power source (not shown) mounted on the rocket. Supplied from a storage battery (not shown) or a solar battery (not shown).
- the turbine 21, the first pump 22, and the second pump 23 may be attached to one rotating shaft shown in FIG. 1, or may be attached to each rotating shaft via a transmission mechanism such as a gear. . In any case, the rotational force from the turbine 21 is transmitted to the first pump 22 and the second pump 23.
- the turbo pump 20 may have a turbine 21 individually, and may be configured by a first pump 22 and a second pump 23 that suck and supply fuel and an oxidant, respectively.
- the electric motor 25 and the clutch 24 are individually provided in the first pump 22 and the second pump 23, and the combustion gas discharged from the sub-combustor 11 is sent to the main combustor 10 via each turbine. Supplied.
- each rotation speed of the 1st pump 22 and the 2nd pump 23 is set individually according to the mixing ratio assumed with each physical property of a fuel and an oxidizing agent.
- the flow path of the fuel discharged from the turbo pump 20 (first pump 22) is divided into a main fuel flow path 32 reaching the heat exchanger 12 and a sub fuel flow path 33 through which fuel flows during precooling at a branch point 31. Branch.
- a fuel supply valve (first valve) 34 is provided in the main fuel flow path 32.
- the fuel supply valve 34 adjusts the amount of fuel supplied to the sub-combustor 11.
- a fuel discharge valve 35 is provided in the auxiliary fuel flow path 33, and its downstream side is open to the outside (atmosphere, outer space).
- the fuel supply valve 34 and the fuel discharge valve 35 are controlled by the controller 13. During precooling, for example, the fuel supply valve 34 is closed and the fuel discharge valve 35 is open. During combustion, the fuel supply valve 34 is open and the fuel discharge valve 35 is closed.
- the flow path of the oxidant discharged from the turbo pump 20 includes a main oxidant flow path 42 that reaches the main combustor 10 at a branch point 41, and a sub-oxidant flow through which the oxidant flows during precooling. Branch to Road 43.
- the main oxidant flow path 42 is provided with an oxidant supply valve (second valve) 44.
- the oxidant supply valve 44 adjusts the supply amount of the oxidant to the main combustor 10.
- an oxidant discharge valve 45 is provided in the sub-oxidant channel 43. The downstream side of the oxidant discharge valve 45 is open to the outside (atmosphere, outer space).
- the oxidant supply valve 44 and the oxidant discharge valve 45 are controlled by the controller 13. At the time of pre-cooling, for example, the oxidant supply valve 44 is closed and the oxidant discharge valve 45 is open. During combustion, the oxidant supply valve 44 is open and the oxidant discharge valve 45 is closed.
- an oxidant adjusting valve third valve 36 is provided in the oxidant flow path 37 from the turbo pump 20 to the sub-combustor 11.
- the oxidant adjusting valve 36 is controlled by the controller 13 and adjusts the amount of oxidant supplied to the sub-combustor 11.
- the fuel supply valve 34 is closed and the fuel discharge valve 35 is open.
- the oxidant supply valve 44 is closed and the oxidant discharge valve 45 is open. Further, the oxidant adjustment valve 36 is closed.
- the electric motor 25 and the turbine 21 are connected via a clutch 24. While this connection is maintained, the electric motor 25 rotates the turbine 21 to drive the turbo pump 20 (the first pump 22 and the second pump 23). The rotational speed of the turbine 21 by the electric motor 25 is much lower than the rotational speed of the turbine 21 when the propulsion system is operating autonomously.
- FIG. 2 shows a flow path cooled by the above-described operation by a bold line.
- the electric motor 25 rotates (drives) the turbine 21 in a state where the electric motor 25 and the turbine 21 are connected to each other via the clutch 24.
- the first pump 22 and the second pump 23 boost the fuel and oxidant and supply the fuel and oxidant into the sub-combustor 11 at a rotational speed that can be handled by the driving force of the electric motor 25. Can do.
- the ignition device 62 ignites (ignites) the sub-combustor 11.
- the auxiliary combustor 11 generates fuel and oxidant combustion gas, and this combustion gas flows through the turbine 21 of the turbo pump 20.
- the turbine 21 is rotated by the flow of the combustion gas, and the first pump 22 and the second pump 23 are driven.
- the first pump 22 boosts the fuel and supplies it to the sub-combustor 11.
- the second pump 23 boosts the oxidizer and supplies it to the main combustor 10 and the subcombustor 11.
- the series of operations increases the pressure in the main combustor 10, and after the mixing ratio of the oxidant and the combustion gas from the subcombustor 11 reaches a predetermined value, the ignition device 61 ignites the main combustor 10. (Ignition) and shift to steady combustion.
- the clutch 24 releases the connection between the electric motor 25 and the turbine 21 under the control of the controller 13. That is, the clutch 24 blocks transmission of rotational force from the electric motor 25 to the turbine 21.
- the clutch 24 releases the connection between the electric motor 25 and the turbine 21. The release of the connection prevents over-rotation of the electric motor 25 and generation of back electromotive force due to high rotation of the turbo pump 20 during combustion.
- the electric motor 25 is used to drive the turbo pump 20 that sucks the propellant from the propellant tank (that is, the fuel tank 30 and the oxidant tank 40) and supplies the propellant to the propulsion system. Is used. Further, the electric motor 25 functions as an auxiliary device for driving the turbo pump 20 during a period when the turbine 21 is not sufficiently driven by the combustion gas from the auxiliary combustor 11 (for example, at the start). Therefore, at the time of pre-cooling, there is no need to pressurize the propellant tank with a pressurized tank as in the prior art, and the pressurized tank can be reduced in size or omitted, and efficient cooling becomes possible. That is, wasteful consumption of the propellant can be suppressed.
- the maximum horsepower (maximum output) required for the electric motor 25 may be a value that can obtain a pressure at which the main combustor 10 is stably ignited. Such a value is about one-hundred compared with the horsepower (tens of thousands of horsepower) of the turbo pump 20 that must discharge high-pressure propellant during steady combustion. That is, it is sufficient that the electric motor 25 is relatively small, and this also contributes to the weight reduction of the propulsion system.
- the propulsion system of the present embodiment may further include a reflux path 53 for returning the fuel and oxidant discharged from the turbo pump 20 to the fuel tank 30 and the oxidant tank 40, respectively.
- the reflux path 53 is constituted by the sub fuel channel 33 and the sub oxidant channel 43, the downstream side of the sub fuel channel 33 is connected to the fuel tank 30, and the downstream side of the sub oxidant channel 43 is Connect to oxidant tank 40.
- the auxiliary fuel channel 33 communicates between the main fuel channel 32 and the fuel tank 30, and the auxiliary oxidant channel 43 communicates between the main fuel channel and the oxidant tank 40.
- the propellant can return to the fuel tank 30 and the oxidant tank 40 because the sub fuel flow path 33 and the sub oxidant flow path 43 constitute the reflux path 53.
- the fuel supply valve 34 is closed, the fuel discharge valve 35 is opened, the oxidant supply valve 44 is closed, and the oxidant discharge valve 45 is open, the fuel is in a liquid state and the first pump 22 is in a liquid state.
- the oxidant is in a liquid state, passes through the second pump 23, and returns to the oxidant tank 40 via the sub-oxidant channel 43. That is, the propellant is not discarded from the propulsion system and circulates within the propulsion system.
- FIG. 4 shows the flow path cooled by the above-described operation with a bold line.
- the above-mentioned propellant circulation is achieved by the electric motor 25 driving the turbo pump 20. That is, the driving of the turbo pump 20 by the electric motor 25 enables the propulsion system to be cooled and the propellant to be reused for combustion in each combustor, and the propellant consumption can be greatly suppressed. This suppression of consumption is particularly effective in orbit (space) where propellant cannot be replenished from the outside.
- the propellant in each tank is agitated. That is, the electric motor 25, the turbo pump 20, and the reflux path 53 constitute a propellant stirring device in the fuel tank 30 and the oxidant tank 40. Since the temperature diffusion of the propellant by convection cannot be used on the orbit, the temperature distribution of the propellant tends to be biased. However, in this embodiment, a propellant is stirred by the refilling of the propellant mentioned above. As a result, the uneven temperature distribution of the propellant in the tank is suppressed.
- a cooling device 50 may be provided according to demands such as cooling efficiency.
- the cooling device 50 is a cryogenic refrigerator operated by electric power. At least the fuel tank 30 and the oxidant tank 40 are cooled.
- the cooling device 50 includes a first cooler 51 that cools the fuel tank 30 and a second cooler 52 that cools the oxidant tank 40. However, these may be integrated into one cooler.
- the cooling device 50 cools the propellant in each tank, and contributes to long-term storage of the propellant, suppression of evaporation, and liquefaction.
- the fuel supply valve (first valve) 34 is provided in the main fuel flow path 32, and the oxidant supply valve (second valve) 44 is provided in the main oxidant flow path 42.
- An oxidant adjusting valve (third valve) 36 is provided in the oxidant flow path 37 from the turbo pump 20 to the sub-combustor 11.
- these valves 34, 44, 36 so-called on / off (ON / OFF) valves that hold either fully open or fully closed positions are generally used.
- variable flow valve is used to increase the stability of the operation of the propulsion system, for example, by making the mixture ratio of the combustion gas constant at the start by the assist of the electric motor 25 described above. At the same time, it functions as a throttle that increases or decreases the thrust while maintaining the mixing ratio by appropriately adjusting the opening during autonomous operation.
- the mixing ratio of the fuel and the oxidant flowing into the sub-combustor 11 is set to a fuel-rich or oxidant-rich value, thereby suppressing an excessive increase in the temperature of the combustion gas. Is preventing.
- the fuel-rich or oxidant-rich mixed gas has a tendency that the combustion temperature changes rapidly with respect to the change in the mixing ratio.
- the electric motor 25 defines the amount of fuel and oxidant supplied into the propulsion system. Therefore, when the oxidant adjustment valve 36 is a variable flow valve, the amount of oxidant can be accurately adjusted with respect to the amount of fuel flowing into the sub-combustor 11.
- the mixing ratio can be set to a value at which high-temperature and high-pressure combustion gas is obtained to such an extent that the supply amount is not wasted and the auxiliary combustor 11 and the turbine 21 are not melted.
- the ignition timing of the sub-combustor 11 can be optimized.
- variable flow rate valves as the fuel supply valve 34 and the oxidant supply valve 44, the thrust adjustment at the rated combustion and the change of the mixture ratio for adjusting the propellant consumption can be continuously performed. (In other words, finely).
- the ignition timing of the main combustor 10 can be easily controlled.
- the propulsion system includes the ignition device 61 of the main combustor 10 and the ignition device 62 of the sub-combustor 11.
- These ignition devices 61 and 62 are constituted by, for example, an ignition agent, a spark plug, and a control device thereof.
- the ignition devices 61 and 62 may be configured by a laser ignition device using laser light.
- the laser ignition device has a light source that generates high-power laser light in a pulse form, and the laser light emitted from the light source heats or evaporates a target (not shown) in the main combustor 10 or the sub-combustor 11 ( So-called ablation) to ignite the surrounding gas mixture.
- the temporal control of the laser beam is very easy.
- ignition by laser light is less affected by the environment at the time of ignition than other ignition methods. Therefore, it is possible to operate in a wide operating range from ignition at low pressure to rated combustion at high pressure.
- the amount of propellant supplied to the propulsion system by the electric motor 25 is regulated, it is easy to control the mixture ratio of the propellant in the main combustor 10 and the subcombustor 11. Therefore, it becomes possible to set the optimal and reliable ignition timing, the ignition sequence is simplified, the development risk can be greatly reduced, and the development cost can be reduced.
- the laser ignition device has a simpler configuration than other ignition devices.
- a conventional ignition device using a spark plug or the like generally occupies nearly 10% of the total mass of the engine including its control device, and the weight can be reduced by replacing this with a laser ignition device. It becomes possible.
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Abstract
Description
Claims (5)
- 燃料と酸化剤の燃焼ガスを生成する副燃焼器と、
前記燃料と前記副燃焼器から排出された前記燃焼ガスとの混合ガスを燃焼する主燃焼器と、
前記燃焼ガスの流通によって回転するタービンと、前記タービンの回転によって駆動されるポンプとを有し、燃料タンクから前記副燃焼器に前記燃料を供給し、且つ、酸化剤タンクから前記副燃焼器及び前記主燃焼器に前記酸化剤を供給するターボポンプと、
前記タービンを回転させる電動モータと、
前記電動モータと前記タービンの連結及び前記連結の解除を行うクラッチと
を備える電動アシスト液体燃料ロケット推進システム。 - 前記ターボポンプから排出された前記燃料及び前記酸化剤をそれぞれ前記燃料タンク及び前記酸化剤タンクに戻す還流路
を更に備える請求項1に記載の電動アシスト液体燃料ロケット推進システム。 - 前記燃料タンク及び前記酸化剤タンクを冷却する冷却装置
を更に備える請求項2に記載の電動アシスト液体燃料ロケット推進システム。 - 前記副燃焼器への前記燃料の供給量を調整する第1バルブと、
前記主燃焼器への前記酸化剤の供給量を調整する第2バルブと、
前記副燃焼器への前記酸化剤の供給量を調整する第3バルブと、
を更に備え、
前記第1バルブ、前記第2バルブおよび前記第3バルブは、何れも可変流量バルブである請求項1~3のうちの何れか一項に記載の電動アシスト液体燃料ロケット推進システム。 - レーザー光を用いて前記副燃焼器及び前記主燃焼器を着火するレーザー着火装置を更に備える請求項1~4のうちの何れか一項に記載の電動アシスト液体燃料ロケット推進システム。
Priority Applications (4)
Application Number | Priority Date | Filing Date | Title |
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JP2018539510A JP6627983B2 (ja) | 2016-09-14 | 2017-04-13 | 電動アシスト液体燃料ロケット推進システム |
RU2019110414A RU2711887C1 (ru) | 2016-09-14 | 2017-04-13 | Жидкостная ракетная двигательная установка со вспомогательной элктрической мощностью |
EP17850473.4A EP3447274B1 (en) | 2016-09-14 | 2017-04-13 | Electric power-assisted liquid-propellant rocket propulsion system |
US16/148,426 US20190032605A1 (en) | 2016-09-14 | 2018-10-01 | Electric power-assisted liquid-propellant rocket propulsion system |
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JP2016179369 | 2016-09-14 | ||
JP2016-179369 | 2016-09-14 |
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US16/148,426 Continuation US20190032605A1 (en) | 2016-09-14 | 2018-10-01 | Electric power-assisted liquid-propellant rocket propulsion system |
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US (1) | US20190032605A1 (ja) |
EP (1) | EP3447274B1 (ja) |
JP (1) | JP6627983B2 (ja) |
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WO (1) | WO2018051566A1 (ja) |
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KR20180045030A (ko) * | 2015-09-14 | 2018-05-03 | 한국항공우주연구원 | 전기모터로 구동되는 부스터 펌프를 사용하는 액체로켓엔진 |
CN109736971A (zh) * | 2018-12-13 | 2019-05-10 | 西安航天动力研究所 | 一种电动泵压式液体火箭发动机 |
CN110761902A (zh) * | 2019-11-05 | 2020-02-07 | 西安中科宇航动力技术有限公司 | 一种电动泵自增压动力系统 |
JP2020133461A (ja) * | 2019-02-18 | 2020-08-31 | 三菱重工業株式会社 | ジェットエンジン |
CN111720238A (zh) * | 2019-07-03 | 2020-09-29 | 西安航天动力研究所 | 基于液氧膨胀循环的深度变推多次起动液体火箭发动机 |
KR20220083169A (ko) * | 2020-12-11 | 2022-06-20 | 한국항공우주연구원 | 하이브리드 터보펌프 시스템을 포함하는 액체 추진제 로켓 엔진 |
KR20230009253A (ko) * | 2021-07-08 | 2023-01-17 | 한양대학교 에리카산학협력단 | 하이브리드 유체 베어링 및 이를 구비하는 전기 펌프 |
WO2023171549A1 (ja) * | 2022-03-10 | 2023-09-14 | 株式会社荏原製作所 | ポンプシステムおよびエンジンシステム |
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KR102041568B1 (ko) | 2015-09-14 | 2019-11-06 | 한국항공우주연구원 | 전기모터로 구동되는 부스터 펌프를 사용하는 액체로켓엔진 |
KR20180045030A (ko) * | 2015-09-14 | 2018-05-03 | 한국항공우주연구원 | 전기모터로 구동되는 부스터 펌프를 사용하는 액체로켓엔진 |
CN109736971A (zh) * | 2018-12-13 | 2019-05-10 | 西安航天动力研究所 | 一种电动泵压式液体火箭发动机 |
CN109736971B (zh) * | 2018-12-13 | 2021-05-04 | 西安航天动力研究所 | 一种电动泵压式液体火箭发动机 |
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CN111720238A (zh) * | 2019-07-03 | 2020-09-29 | 西安航天动力研究所 | 基于液氧膨胀循环的深度变推多次起动液体火箭发动机 |
CN111720238B (zh) * | 2019-07-03 | 2021-08-10 | 西安航天动力研究所 | 基于液氧膨胀循环的深度变推多次起动液体火箭发动机 |
CN110761902A (zh) * | 2019-11-05 | 2020-02-07 | 西安中科宇航动力技术有限公司 | 一种电动泵自增压动力系统 |
KR102473169B1 (ko) * | 2020-12-11 | 2022-12-01 | 한국항공우주연구원 | 하이브리드 터보펌프 시스템을 포함하는 액체 추진제 로켓 엔진 |
KR20220083169A (ko) * | 2020-12-11 | 2022-06-20 | 한국항공우주연구원 | 하이브리드 터보펌프 시스템을 포함하는 액체 추진제 로켓 엔진 |
KR20230009253A (ko) * | 2021-07-08 | 2023-01-17 | 한양대학교 에리카산학협력단 | 하이브리드 유체 베어링 및 이를 구비하는 전기 펌프 |
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WO2023171549A1 (ja) * | 2022-03-10 | 2023-09-14 | 株式会社荏原製作所 | ポンプシステムおよびエンジンシステム |
Also Published As
Publication number | Publication date |
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RU2711887C1 (ru) | 2020-01-23 |
JP6627983B2 (ja) | 2020-01-08 |
US20190032605A1 (en) | 2019-01-31 |
EP3447274A4 (en) | 2019-12-11 |
EP3447274A1 (en) | 2019-02-27 |
JPWO2018051566A1 (ja) | 2019-02-07 |
EP3447274B1 (en) | 2021-06-16 |
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