WO2015146356A1 - 燃焼器、ジェットエンジン、飛しょう体及びジェットエンジンの動作方法 - Google Patents

燃焼器、ジェットエンジン、飛しょう体及びジェットエンジンの動作方法 Download PDF

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Publication number
WO2015146356A1
WO2015146356A1 PCT/JP2015/054025 JP2015054025W WO2015146356A1 WO 2015146356 A1 WO2015146356 A1 WO 2015146356A1 JP 2015054025 W JP2015054025 W JP 2015054025W WO 2015146356 A1 WO2015146356 A1 WO 2015146356A1
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WIPO (PCT)
Prior art keywords
igniter
jet engine
air
inlet
combustor
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Ceased
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PCT/JP2015/054025
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English (en)
French (fr)
Japanese (ja)
Inventor
祥彦 上野
鈴木　佑
正二郎 古谷
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Mitsubishi Heavy Industries Ltd
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Mitsubishi Heavy Industries Ltd
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by Mitsubishi Heavy Industries Ltd filed Critical Mitsubishi Heavy Industries Ltd
Priority to EP15767934.1A priority Critical patent/EP3098427B1/en
Priority to US15/120,859 priority patent/US11067036B2/en
Publication of WO2015146356A1 publication Critical patent/WO2015146356A1/ja
Anticipated expiration legal-status Critical
Ceased legal-status Critical Current

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Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02K—JET-PROPULSION PLANTS
    • F02K7/00—Plants in which the working fluid is used in a jet only, i.e. the plants not having a turbine or other engine driving a compressor or a ducted fan; Control thereof
    • F02K7/10—Plants in which the working fluid is used in a jet only, i.e. the plants not having a turbine or other engine driving a compressor or a ducted fan; Control thereof characterised by having ram-action compression, i.e. aero-thermo-dynamic-ducts or ram-jet engines
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02K—JET-PROPULSION PLANTS
    • F02K7/00—Plants in which the working fluid is used in a jet only, i.e. the plants not having a turbine or other engine driving a compressor or a ducted fan; Control thereof
    • F02K7/10—Plants in which the working fluid is used in a jet only, i.e. the plants not having a turbine or other engine driving a compressor or a ducted fan; Control thereof characterised by having ram-action compression, i.e. aero-thermo-dynamic-ducts or ram-jet engines
    • F02K7/105—Plants in which the working fluid is used in a jet only, i.e. the plants not having a turbine or other engine driving a compressor or a ducted fan; Control thereof characterised by having ram-action compression, i.e. aero-thermo-dynamic-ducts or ram-jet engines using a solid fuel
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
    • F02C7/26—Starting; Ignition
    • F02C7/264—Ignition
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2220/00—Application
    • F05D2220/10—Application in ram-jet engines or ram-jet driven vehicles
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2260/00—Function
    • F05D2260/99—Ignition, e.g. ignition by warming up of fuel or oxidizer in a resonant acoustic cavity
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T50/00—Aeronautics or air transport
    • Y02T50/60—Efficient propulsion technologies, e.g. for aircraft

Definitions

  • the present invention relates to a combustor, a jet engine, a flying object, and a method for operating the jet engine.
  • Turbojet engines including turbofan engines, ramjet engines, and scramjet engines are known as jet engines that fly faster than the speed of sound. These are jet engines that operate by taking in air. In particular, in the ramjet engine and the scramjet engine, the speed of the taken-in air strongly depends on the flying speed.
  • FIG. 1 is a schematic cross-sectional view schematically showing a configuration of a conventional jet engine.
  • the jet engine includes a fuselage 110 and a cowl 140 provided to form a space 150 through which gas can flow under the fuselage 110.
  • the lower part in front of the airframe 110 and the front part of the cowl 140 constitute an inlet 111 for introducing air into the space 150.
  • an inlet cover is detachably provided in front of the inlet 111.
  • the inlet cover is attached in front of the inlet 111 until the engine is started.
  • the inlet cover is used to reduce the aerodynamic resistance of the fuselage and prevent foreign matter from entering the engine until the engine is started.
  • the lower part in the middle of the fuselage 110 and the middle part of the cowl 140 constitute a combustor 112 that mixes and burns fuel and air.
  • the lower part behind the body 110 and the rear part of the cowl 140 constitute a nozzle 113 that expands and discharges combustion gas.
  • the combustor 112 includes a fuel injector 120, an igniter 121, an igniter actuator 122, and an igniter controller 123.
  • the fuel injector 120 is provided in a portion corresponding to the combustor 112 in the lower portion of the fuselage 110.
  • the fuel injector 120 injects fuel toward the space 150.
  • the igniter 121 radiates a flame for igniting a mixture of fuel and air toward the space 150.
  • the igniter actuator 122 starts the igniter 121 with electric energy or thermal energy.
  • the igniter controller 123 generates a signal and the like for starting the igniter actuator 122.
  • the operation procedure of the jet engine in FIG. 1 is as follows. First, after the flying body reaches the desired speed, the inlet cover is separated from the inlet 111. Second, the igniter controller 123 generates a signal or the like for starting the igniter actuator 122, and the signal or the like is transmitted to the igniter actuator 122. Third, based on the transmitted signal or the like, the igniter actuator 122 is started to generate electric energy or thermal energy or the like. The igniter 121 is started by the electric energy or thermal energy. The started igniter 121 radiates a flame toward the space 150. The fuel injector 120 injects fuel toward the space 150 at the timing before and after the emission of the flame. The injected fuel is ignited and burned by the flame. The combustion is continued by continuously injecting fuel from the fuel injector 120. Combustion gas generated by the combustion is discharged from the nozzle 113. The fuselage 110 flies and gains thrust by the released combustion gas.
  • Japanese Patent Publication No. 6-60597 discloses a scramjet combustor ignition and flame holding method.
  • the publication describes an igniter, but the specific operation procedure of the igniter is unknown.
  • An object of the present invention is to provide a combustor, a jet engine, a flying body, and a method for operating a jet engine that can reduce the mass of the airframe and simplify the design of the airframe.
  • the combustor is a combustor that burns fuel using air taken in from an inlet.
  • the combustor includes an igniter for igniting the air-fuel mixture.
  • the igniter is configured to automatically ignite and operate by heat and pressure generated by compression of the air taken in from the inlet.
  • An operation method of a jet engine includes an inlet that takes in air, a combustor that burns fuel using the air to generate combustion gas, and delivers the combustion gas from behind the jet engine.
  • a method of operating a jet engine comprising a nozzle.
  • the combustor includes an igniter for igniting an air-fuel mixture of the air and the fuel.
  • the operation method of the jet engine includes a step of taking air from the inlet, automatically igniting and operating the igniter by heat and pressure generated by compression of the taken-in air, and igniting the igniter. Radiating the resulting flame toward the mixture of the air and the fuel, and combusting the mixture with the flame to start the jet engine.
  • the present invention provides a combustor, a jet engine, a flying body, and a jet engine operating method capable of reducing the mass of the airframe and simplifying the design of the airframe.
  • FIG. 1 is a schematic cross-sectional view schematically showing a configuration of a conventional jet engine.
  • FIG. 2 is a perspective view showing an example of the configuration of the flying object according to the embodiment.
  • FIG. 3 is a schematic cross-sectional view schematically showing the configuration of the jet engine according to the first embodiment, and shows a state before the engine and the igniter are started.
  • FIG. 4 is a schematic cross-sectional view schematically showing the configuration of the jet engine according to the first embodiment, and is a view showing a state when the engine and the igniter are started.
  • FIG. 1 is a schematic cross-sectional view schematically showing a configuration of a conventional jet engine.
  • FIG. 2 is a perspective view showing an example of the configuration of the flying object according to the embodiment.
  • FIG. 3 is a schematic cross-sectional view schematically showing the configuration of the jet engine according to the first embodiment, and shows a state before the engine and the igniter are started.
  • FIG. 4 is a schematic cross-sectional view schematically
  • FIG. 5A is a table showing the relationship among Mach number, altitude and total temperature.
  • FIG. 5B is a table showing the relationship among Mach number, altitude and total pressure.
  • FIG. 5C is a graph showing the relationship between Mach number, altitude and total temperature.
  • FIG. 5D is a graph showing the relationship between Mach number, altitude, and total pressure.
  • FIG. 6 is a schematic cross-sectional view schematically showing a configuration of a jet engine according to a modification of the first embodiment, and is a diagram showing a state when the engine and the igniter are started.
  • FIG. 7 is a schematic cross-sectional view schematically showing the configuration of the combustor 7 of the jet engine of the second embodiment, and is a diagram showing a state before the engine and the igniter are started.
  • FIG. 6 is a schematic cross-sectional view schematically showing a configuration of a jet engine according to a modification of the first embodiment, and is a diagram showing a state when the engine and the igniter are started.
  • FIG. 8 is a schematic cross-sectional view schematically showing the configuration of the combustor 7 of the jet engine according to the second embodiment, and is a diagram showing a state when the engine and the igniter are started.
  • FIG. 9 is a schematic cross-sectional view schematically showing the configuration of the combustor 7 of the jet engine according to the third embodiment, and shows a state before the engine and the igniter are started.
  • FIG. 10 is a schematic cross-sectional view schematically illustrating the configuration of the combustor 7 of the jet engine according to the third embodiment, and is a diagram illustrating a state when the engine and the igniter are started.
  • FIG. 11 is a schematic cross-sectional view schematically showing the configuration of the combustor 7 of the jet engine of the fourth embodiment, and is a diagram showing a state before the engine and the igniter are started.
  • FIG. 2 is a perspective view showing an example of the configuration of the flying object 1 according to the embodiment.
  • the flying body 1 includes a jet engine 2 and a propulsion device 5.
  • the propulsion device 5 accelerates the flying object 1 from a speed at the start of flying to a desired speed when flying the flying object 1 from the launching device.
  • the speed at the start of flying is zero when the flying object 1 is launched from a stationary launching device, and the flying object is moving (or in flight) (or ,
  • the flying speed of the moving object (or flying object) is the moving speed (or flying speed) of the moving object (or flying object).
  • a specific example of the propulsion device 5 is a rocket motor.
  • the propulsion device 5 may be any device as long as it can accelerate the flying object to a desired speed. For example, when the flying object 1 is loaded on another second flying object and accelerated to a desired speed, the second flying object becomes the propulsion device 5.
  • the jet engine 2 After the propulsion device 5 is separated from the flying object 1, the jet engine 2 further accelerates the flying object 1 to fly toward the target.
  • the jet engine 2 includes a body 3 and a cowl 4.
  • the airframe 3 and the cowl 4 constitute an inlet, a combustor, and a nozzle of the jet engine 2 as described later.
  • the jet engine 2 takes in air from the front at the inlet, mixes the air and fuel with the combustor and burns them, expands the combustion gas with the nozzles, and sends them back. Thereby, the jet engine 2 obtains a propulsive force.
  • the jet engine 2 is composed of the lower part of the body 3 and the cowl 4, but the jet engine 2 may be composed of a cylindrical body installed in the lower part or inside of the body 3. In this case, the front part of the cylindrical body constitutes an inlet, the central part constitutes a combustor, and the rear part constitutes a nozzle.
  • FIGS. 3 and 4 are schematic cross-sectional views schematically showing an example of the configuration of the jet engine according to the first embodiment.
  • FIG. 3 is a diagram showing a state before the engine and the igniter are started
  • FIG. 4 is a diagram showing a state when the engine and the igniter are started.
  • the jet engine 2 includes an airframe 3 and a cowl 4 provided below the airframe 3 so as to form a space 50 through which gas can flow.
  • the lower part in front of the airframe 3 and the front part of the cowl 4 constitute an inlet 6 for introducing air into the space 50.
  • An inlet cover 9 is provided in front of the inlet 6 so as to be separable.
  • the lower part in the middle of the fuselage 3 and the middle part of the cowl 4 constitute a combustor 7 that mixes and burns fuel and air.
  • the lower part behind the airframe 3 and the rear part of the cowl 4 constitute a nozzle 8 that expands and discharges combustion gas.
  • the combustor 7 includes an igniter 61 and a fuel injector 62.
  • the fuel injector 62 is provided in a portion corresponding to the combustor 7 in the lower portion of the body 3.
  • the fuel injector 62 injects the fuel stored in the airframe 3 toward the space 50.
  • the igniter 61 radiates a flame toward the space 50 by burning solid fuel, for example.
  • the igniter 61 is automatically ignited and operated by heat and pressure generated by compression of air taken in from the inlet 6 as described later. In the present embodiment, since the igniter automatically ignites, the igniter controller and the igniter actuator are not provided.
  • a jet engine in addition to an igniter, additional devices such as an igniter controller and an igniter actuator may be provided.
  • additional devices such as an igniter controller and an igniter actuator may be provided.
  • the igniter controller and the igniter actuator are not provided. As a result, the mass of the aircraft can be reduced, and the design of the aircraft can be simplified.
  • the cable portion connected from the igniter controller to the igniter actuator can be a fragile part in the engine. For this reason, the circumference
  • the igniter 61 is embedded in a groove 63 provided on the wall surface of the combustor 7.
  • the igniter 61 is, for example, a solid rocket motor (solid RM).
  • a solid rocket motor is defined as a device that burns solid fuel and emits a flame.
  • the solid rocket motor used in the present embodiment is preferably a solid rocket motor having a low ignition point that ignites at approximately 0.8 times or less of the temperature corresponding to the flying Mach number and altitude shown in FIGS. 5A and 5C.
  • Examples of the solid rocket motor having a low ignition point include KClO 3 (70% by weight), lactose (25% by weight), and aliphatic polycarbonate (5% by weight), which are materials that spontaneously ignite at a temperature of 195 degrees Celsius. it can. The materials are described in detail in JP-T-2001-503350.
  • the temperature and pressure of the air near the wall surface of the combustor 7 is the temperature 626. This is a value obtained by dividing the temperature drop due to heat transfer to the wall surface and mainstream air from 8K (353.7 degrees), and a value obtained by dividing the pressure loss during compression from the pressure of 0.97 MPa.
  • a portion 64 facing the space 50 hereinafter referred to as “igniter surface portion 64” also has the same temperature and temperature as the wall surface of the combustor 7. Pressure increases.
  • the igniter 61 automatically ignites and radiates a flame toward the space 50.
  • the higher the pressure the lower the spontaneous ignition temperature. Therefore, when KClO 3 (70 wt%), lactose (25 wt%), and aliphatic polycarbonate (5 wt%), which are materials that spontaneously ignite at a temperature of 195 degrees under atmospheric pressure, are used as solid rocket motors. Under conditions higher than atmospheric pressure, the solid rocket motor will spontaneously ignite at a temperature lower than 195 degrees.
  • the inlet cover 9 is separated from the inlet 6. Then, high-speed air flows into the space 50, and as a result, the temperature and pressure of the igniter surface portion 64 rise as described above. Subsequently, when the temperature and pressure of the igniter surface portion 64 rise above the ignition condition of the solid rocket motor of the igniter 61, the igniter 61 automatically ignites and radiates a flame toward the space 50.
  • the fuel injector 62 is directed toward the space 50 at a timing before and after the flame emission (that is, a timing immediately before the flame emission, a timing at the same time as the flame emission, or a timing immediately after the flame emission). Inject fuel. The injected fuel is ignited and burned by the flame. The combustion is continued by continuously injecting fuel from the fuel injector 62. The combustion gas generated by the combustion is discharged from the nozzle 8.
  • the flying object 1 flies with the thrust of the released combustion gas.
  • the start of the igniter 61 can be controlled by the timing of separation of the inlet cover 9.
  • the time required from separation of the inlet cover 9 to ignition is about several milliseconds to several seconds. Separation of the inlet cover 9 may be performed automatically or manually by remote control.
  • the time elapsed from the reference time is measured with a timer, etc., when the flying object 1 takes off, etc., and the inlet cover 9 is automatically separated when the elapsed time exceeds a preset time. You may be made to do.
  • an altimeter, speedometer, or Mach meter is installed on the flying object 1, and the inlet cover 9 is automatically separated when the altitude, speed, or Mach number exceeds a preset value. Also good. For example, in the case of a ramjet, when the Mach number exceeds a preset value in the range of 1 to 5, it is preferable that the ramjet is automatically separated.
  • the jet engine may include a controller (not shown) that controls the separation of the inlet cover.
  • the controller is, for example, a computer including a hardware processor, a memory, and a communication interface.
  • an operation signal is input to the controller from a remote operation device (not shown).
  • the controller transmits a separation signal to an inlet cover separation mechanism (not shown).
  • the inlet cover separation mechanism separates the inlet cover 9 from the jet engine in response to receiving the separation signal.
  • the controller includes a timer (which may be a timer realized by a hardware processor executing a program).
  • the timer transmits an expiration signal to the controller when the elapsed time from the reference time exceeds a preset value.
  • the controller transmits a separation signal to an inlet cover separation mechanism (not shown).
  • the inlet cover separation mechanism separates the inlet cover 9 from the jet engine in response to receiving the separation signal.
  • the jet engine or flying object is equipped with a sensor such as an altimeter, speedometer or Mach meter.
  • the controller receives a signal corresponding to the measurement value from the sensor.
  • the controller transmits a separation signal to an inlet cover separation mechanism (not shown).
  • the inlet cover separation mechanism separates the inlet cover 9 from the jet engine in response to receiving the separation signal.
  • an igniter controller, an igniter actuator, and an igniter controller and an igniter actuator are used for starting the igniter 61.
  • the mass of the flying body 1, the jet engine 2, and the combustor 7 can be reduced, and the structure and design can be simplified.
  • the manufacturing cost of the flying object 1, the jet engine 2, and the combustor 7 can be reduced.
  • the performance of the flying object 1 and the jet engine 2 is improved as the mass is reduced.
  • by controlling the separation timing of the inlet cover 9 based on the value measured by the timer, altimeter, speedometer, or Mach meter, the timing of starting the igniter 61 despite the simple structure. Can be controlled accurately.
  • FIG. 6 is a schematic cross-sectional view schematically showing a configuration of a jet engine according to a modification of the first embodiment, and is a diagram showing a state when the engine and the igniter are started.
  • the start timing of the igniter 61 is controlled by separating the inlet cover 9.
  • the nozzle cover 10 is provided behind the nozzle 8, and the start timing of the igniter 61 can be controlled by separating the nozzle cover 10. That is, when the nozzle cover 10 is attached to the nozzle 8, the inflow of air into the space 50 is suppressed. For this reason, the temperature and pressure of the igniter surface portion 64 do not rise beyond the ignition conditions of the solid rocket motor of the igniter 61.
  • the nozzle cover 10 when the nozzle cover 10 is separated from the nozzle 8, high-speed air flows into the space 50. For this reason, the temperature and pressure of the igniter surface portion 64 rise beyond the ignition conditions of the solid rocket motor of the igniter 61, and as a result, the igniter 61 automatically ignites.
  • the nozzle cover 10 may be anything as long as it substantially covers the nozzle flow path, and the propulsion device 5 may also serve as the nozzle cover 10.
  • the temperature and pressure of the surface portion 64 of the igniter can be controlled from a state not exceeding the ignition condition of the solid rocket motor of the igniter 61 to a state exceeding the ignition condition, the inlet cover or the nozzle cover is separated.
  • the modified example of the first embodiment has the same effect as the first embodiment.
  • a nozzle cover separation mechanism (not shown) may be provided.
  • the operation procedure of the nozzle cover separation mechanism is the same as the operation procedure of the inlet cover in the first embodiment (for example, the nozzle cover in response to an operation signal, an expiration signal, or a signal received from the sensor by the controller).
  • a separation signal is transmitted to the separation mechanism, and the nozzle cover separation mechanism separates the nozzle cover from the jet engine in response to receiving the separation signal.
  • FIGS. 7 and 8 are schematic cross-sectional views schematically showing an example of the configuration of the combustor 7 of the jet engine according to the second embodiment.
  • FIG. 7 is a diagram showing a state before the engine and the igniter are started
  • FIG. 8 is a diagram showing a state when the engine and the igniter are started.
  • the same reference numerals are used for the same components as those in the first embodiment.
  • the second embodiment differs from the first embodiment in that it includes a stagnation point former 65 and in that the ignition temperature of an igniter 61, for example, a solid rocket motor, is relatively high.
  • the stagnation point former 65 is provided in the vicinity of the igniter surface portion 64.
  • the wall surface of the combustor 7 is provided behind the igniter surface portion 64.
  • the stagnation point former 65 is preferably provided behind the igniter surface portion 64 in contact with the igniter surface portion 64 or at a small distance from the igniter surface portion 64.
  • the terms “front” and “rear” mean the upstream side and the downstream side with respect to the flow of the mainstream air.
  • the operation principle of the igniter 61 in this embodiment will be described.
  • the mainstream air flows into the space 50 triggered by the separation of the inlet cover 9 and the nozzle cover 10.
  • Part of the mainstream air is dammed by the stagnation point former 65, and the temperature and pressure of the damped air rises significantly.
  • the igniter 61 automatically ignites and radiates a flame toward the space 50.
  • the heat transfer between the air blocked by the stagnation point former 65 and the air flowing in the vicinity of the wall surface and the stagnation point former 65 is an air whose velocity is zero at the igniter surface portion 64 in the first embodiment.
  • the solid rocket motor used in the present embodiment can be one having a higher ignition temperature than the solid rocket motor used in the first embodiment.
  • a solid rocket motor having a high ignition point that ignites at a temperature corresponding to the flying Mach number and altitude shown in FIGS. 5A and 5C is used.
  • the solid rocket motor having a high ignition point include AP / C 2 H 4 O / Al that ignites at a temperature of 400 degrees Celsius.
  • the stagnation point forming device 65 becomes a fluid resistance against the mainstream air, and thus becomes a factor that degrades the performance of the engine 2.
  • the stagnation point former can be composed of an ablation material such as silica or phenol, a metal such as aluminum, or an explosive.
  • the stagnation point former 65 may be formed behind the igniter 61 over the same span (length in the width direction) as that of the igniter 61, or over a shorter span (length in the width direction) than the igniter 61. It may be formed. In the latter case, a plurality of stagnation point formers 65 may be formed along the span of the igniter 61.
  • This embodiment has the following effects in addition to the same effects as the first embodiment.
  • the material of the solid rocket motor in this embodiment is a material having a high ignition temperature compared to the first embodiment, there is a risk that the solid rocket motor will spontaneously ignite unexpectedly. Can be reduced.
  • solid rocket motors with a high ignition point have higher energy generation per unit area / mass than solid rocket motors with a low ignition point. Can be miniaturized.
  • the stagnation point former 65 by providing the stagnation point former 65 and significantly increasing the temperature and pressure in the area in front of the stagnation point former 65, the ignition of the igniter 61 becomes more reliable.
  • the stagnation point former 65 by configuring the stagnation point former 65 with a member that automatically disappears due to heat and pressure after the igniter 61 is started, the performance of the jet engine 2 is not deteriorated.
  • FIGS. 9 and 10 are schematic cross-sectional views schematically showing an example of the configuration of the combustor 7 of the jet engine according to the third embodiment.
  • FIG. 9 is a diagram showing a state before the engine and the igniter are started
  • FIG. 10 is a diagram showing a state when the engine and the igniter are started.
  • the same reference numerals are used for the same components as those in the first embodiment.
  • the third embodiment is different from the first embodiment in that the igniter 61 includes an ignition explosive 66, and the ignition temperature of the igniter 61, for example, a solid rocket motor, is relatively high. Different.
  • the ignition explosive 66 is provided in the vicinity of the igniter surface portion 64 or the igniter surface portion 64. Typically, it is provided on the igniter surface portion 64, for example, as shown in FIGS.
  • the operation principle of the igniter 61 in this embodiment will be described.
  • the mainstream air flows into the space 50 triggered by the separation of the inlet cover 9 and the nozzle cover 10. Similar to the first embodiment, the mainstream air is compressed. And the temperature and pressure of the igniter surface part 64 rise with the temperature and pressure rise of compressed air.
  • the ignition explosive 66 explodes. As the ignition explosive 66 explodes, the igniter 61 automatically ignites and radiates a flame toward the space 50.
  • the solid rocket motor that is the igniter 61 is ignited by using the ignition explosive 66, so that the solid rocket motor used in the present embodiment is the solid rocket used in the first embodiment.
  • the solid rocket motor used in the present embodiment is the solid rocket used in the first embodiment.
  • one having a high ignition temperature can be used.
  • a solid rocket motor having a high ignition point that ignites at a temperature of 400 degrees Celsius is used.
  • the solid rocket motor having a high ignition point include AP / C 2 H 4 O / Al that ignites at a temperature of 400 degrees.
  • This embodiment has the following effects in addition to the same effects as the first embodiment. 1stly, in this embodiment, since the solid rocket motor which is the igniter 61 is ignited using the ignition explosive 66, ignition of the igniter 61 becomes more reliable. In addition, solid rocket motors with a high ignition point have higher energy generation per unit area / mass than solid rocket motors with a low ignition point. Can be miniaturized.
  • FIG. 11 is a schematic cross-sectional view schematically showing an example of the configuration of the combustor 7 of the jet engine according to the fourth embodiment.
  • FIG. 11 is a diagram showing a state before the engine and the igniter are started.
  • the same reference numerals are used for the same components as those in the first embodiment.
  • the fourth embodiment differs from the first embodiment in that the igniters 61 are provided at a plurality of locations (two locations in FIG. 11) along the longitudinal direction of the engine.
  • the interval 67 between the plurality of igniters 61 is set to a distance at which the flame emission when one igniter 61 is activated does not propagate to the other igniter 61. For this reason, even if one igniter 61 spontaneously ignites at an unexpected time, the other igniter 61 is maintained in a non-operating state. For this reason, even when one of the igniters 61 becomes unusable due to unexpected ignition or the like, the timing of the jet engine start can be controlled by using the other igniter 61.
  • This embodiment has the following effects in addition to the same effects as the first embodiment. That is, even when one igniter 61 becomes unusable, the use of the other igniter 61 makes it possible to control the timing of starting the jet engine, thereby improving the reliability of jet engine operation. .
  • the flying object includes an aircraft or a rocket.
  • FIGS. 3 to 4 and FIGS. 6 to 11 do not show a flame holder.
  • a flame holder may be provided in the vicinity of the igniter 61, and a flame for continuously burning a mixture of fuel and air may be held by the flame holder.

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  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Toys (AREA)
  • Ignition Installations For Internal Combustion Engines (AREA)
PCT/JP2015/054025 2014-03-26 2015-02-13 燃焼器、ジェットエンジン、飛しょう体及びジェットエンジンの動作方法 Ceased WO2015146356A1 (ja)

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WO2017158856A1 (ja) * 2016-03-16 2017-09-21 三菱重工業株式会社 ジェットエンジン、および飛しょう体
RU186094U1 (ru) * 2018-01-11 2018-12-29 Акционерное Общество "Государственное Машиностроительное Конструкторское Бюро "Радуга" Имени А.Я. Березняка" Сверхзвуковой прямоточный воздушно-реактивный двигатель (варианты)

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CN114320609B (zh) * 2022-03-03 2022-06-07 中国空气动力研究与发展中心计算空气动力研究所 一种高超声速超燃发动机的燃料喷射装置

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RU186094U1 (ru) * 2018-01-11 2018-12-29 Акционерное Общество "Государственное Машиностроительное Конструкторское Бюро "Радуга" Имени А.Я. Березняка" Сверхзвуковой прямоточный воздушно-реактивный двигатель (варианты)

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US20170058835A1 (en) 2017-03-02
EP3098427A4 (en) 2017-03-15
US11067036B2 (en) 2021-07-20
EP3098427B1 (en) 2018-12-26
JP2015183683A (ja) 2015-10-22
EP3098427A1 (en) 2016-11-30
JP6310293B2 (ja) 2018-04-11

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