CN114542285A - Engine, control method thereof and aircraft - Google Patents

Engine, control method thereof and aircraft Download PDF

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Publication number
CN114542285A
CN114542285A CN202210217441.1A CN202210217441A CN114542285A CN 114542285 A CN114542285 A CN 114542285A CN 202210217441 A CN202210217441 A CN 202210217441A CN 114542285 A CN114542285 A CN 114542285A
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Prior art keywords
flow channel
engine
speed
valve plate
aircraft
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CN202210217441.1A
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CN114542285B (en
Inventor
周琨
苗辉
李亚忠
马薏文
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China Aero Engine Research Institute
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China Aero Engine Research Institute
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    • 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
    • F02C3/00—Gas-turbine plants characterised by the use of combustion products as the working fluid
    • F02C3/04—Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor
    • 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
    • F02C3/00—Gas-turbine plants characterised by the use of combustion products as the working fluid
    • F02C3/14—Gas-turbine plants characterised by the use of combustion products as the working fluid characterised by the arrangement of the combustion chamber in the plant
    • 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
    • F02C9/00—Controlling gas-turbine plants; Controlling fuel supply in air- breathing jet-propulsion plants
    • F02C9/16—Control of working fluid flow
    • F02C9/18—Control of working fluid flow by bleeding, bypassing or acting on variable working fluid interconnections between turbines or compressors or their stages
    • 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
    • F02C9/00—Controlling gas-turbine plants; Controlling fuel supply in air- breathing jet-propulsion plants
    • F02C9/16—Control of working fluid flow
    • F02C9/20—Control of working fluid flow by throttling; by adjusting vanes
    • 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

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  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Physics & Mathematics (AREA)
  • Fluid Mechanics (AREA)
  • Control Of Turbines (AREA)

Abstract

An engine, a control method thereof and an aircraft, wherein the engine comprises a first turbine and a mode conversion device; the first turbine is arranged in the first flow passage and is used for depressurizing the gas in the cold end flow passage and then introducing the gas into the hot end flow passage; the mode conversion device is arranged in the cold-end flow channel and can adjust the opening degree of the first flow channel and the opening degree of the second flow channel. According to the aircraft, the first flow channel and the second flow channel are arranged in the cold-end flow channel, the first turbine is arranged in the first flow channel, and when the aircraft flies at a low speed, the mode conversion device can control airflow to pass through the second flow channel, so that pressurization and depressurization are not carried out; when the aircraft flies at a high speed, the mode conversion device controls the air flow to pass through the first turbine of the first flow channel, and the air flow is depressurized and then introduced into the hot end flow channel. The large-range adjustment of the circulating pressure ratio can be conveniently realized, the flight speed limit of the engine can be greatly expanded, and the performance of the engine in a high-speed state and a low-speed state can be considered.

Description

Engine, control method thereof and aircraft
Technical Field
The disclosure belongs to the technical field of aero-engines, and particularly relates to an engine, a control method of the engine and an aircraft.
Background
The flight speed range of the aircraft engine is generally Ma 0-2, and the technical approach of the variable-cycle engine aims to further increase the working capacity in a low-speed state, increase the thrust and reduce the oil consumption. The measure is mainly to increase the circulation capacity of the culvert airflow.
However, the variable cycle concept of the aircraft engine can be used for speeding up, and the flight speed range is expanded to Ma 3-4. When the aircraft engine flies at a low speed, the air compressor is required to provide a high pressure ratio (such as 25-30) so as to ensure high cycle efficiency and low oil consumption. And when the aircraft flies above Ma3 at a high speed, the supercharging effect of the air inlet channel is gradually obvious, and the compressor is required to provide a lower supercharging ratio (less than 10).
Disclosure of Invention
In order to solve the above technical problem, an object of the present disclosure is to provide an engine capable of generating a pressure reduction effect on an air flow at a high speed state.
In order to achieve the purpose of the disclosure, the technical scheme adopted by the disclosure is as follows:
an engine, includes the engine block, have the sprue for annular in the cross-section in the engine block, the sprue is including cold junction runner and hot junction runner that communicate in proper order, still be provided with first runner and second runner in the cold junction runner, air in the cold junction runner can pass through first runner and/or the second runner intercommunication to the hot junction runner, still includes:
the first turbine is arranged in the first flow passage and is used for depressurizing the gas in the cold end flow passage and then introducing the gas into the hot end flow passage;
and the mode conversion device is arranged in the cold-end flow channel and can adjust the opening degree of the first flow channel and the opening degree of the second flow channel.
Optionally, a first compressor is installed in the second flow channel, and the first compressor is used for boosting the gas in the cold end flow channel and then introducing the boosted gas into the hot end flow channel.
Optionally, the mode switching device includes a mode switching valve plate and a mode switching unit for controlling the mode switching valve plate to act;
when the mode conversion valve plate is at the first position, the first flow channel is opened, and the second flow channel is closed;
when the mode conversion valve plate is at the second position, the first flow channel is closed, and the second flow channel is opened;
when the mode conversion valve plate is located between the first position and the second position, the proportion of the first flow channel opening and the second flow channel opening can be adjusted.
Optionally, the cross sections of the first flow channel and the second flow channel are both annular, the first flow channel and the second flow channel are coaxially arranged in the cold end flow channel, and a partition plate is arranged between the first flow channel and the second flow channel;
the mode switching valve plate comprises a first annular valve plate and a second annular valve plate, the first annular valve plate and the second annular valve plate are connected with the mode switching device, the first annular valve plate is hinged to the position between the inlet of the first flow channel and the inlet of the second flow channel, and the second annular valve plate is hinged to the position between the outlet of the first flow channel and the outlet of the second flow channel.
Optionally, the annular first flow passage is arranged at the outer ring of the annular second flow passage.
Optionally, the second flow channel is an empty flow channel, and the annular second flow channel is arranged on the outer ring of the annular first flow channel.
Optionally, a second compressor is installed at an inlet of the cold-end runner, a third compressor is installed at an outlet of the cold-end runner, and the first turbine is arranged between the second compressor and the third compressor.
The present disclosure also provides an aircraft comprising the above engine
The present disclosure also provides a control method based on the above engine,
when the speed of the aircraft is less than or equal to a first preset speed, the mode conversion device controls the second flow channel to be opened, and the first flow channel is closed;
when the speed of the aircraft is between a first preset speed and a second preset speed and the speed of the aircraft is increased, the mode conversion device increases the opening degree of the first flow channel and reduces the opening degree of the second flow channel; when the speed of the aircraft is between a first preset speed and a second preset speed and the speed of the aircraft is reduced, the mode conversion device reduces the opening degree of the first flow channel and increases the opening degree of the second flow channel;
when the speed of the aircraft is greater than or equal to a second preset speed, the mode conversion device controls the first flow channel to be opened, and the second flow channel is closed;
the first preset speed is less than the second preset speed.
Optionally, the first preset speed is at mach 1.8-2.2, and the second preset speed is at mach 2.8-3.2.
In the disclosure, the first flow passage and the second flow passage are arranged in the cold-end flow passage, and the first turbine is arranged in the first flow passage, so that when the aircraft flies at a low speed, the mode conversion device can control airflow to pass through the second flow passage, and the air flow is not pressurized or depressurized; when the aircraft flies at a high speed, the mode conversion device controls the air flow to pass through the first turbine of the first flow channel, and the air flow is depressurized and then introduced into the hot end flow channel. The large-range adjustment of the circulating pressure ratio can be conveniently realized, the flight speed limit of the engine can be greatly expanded, and the performance of the engine in a high-speed state and a low-speed state can be considered.
Drawings
The accompanying drawings, which are included to provide a further understanding of the disclosure and are incorporated in and constitute a part of this specification, illustrate exemplary embodiments of the disclosure and together with the description serve to explain the principles of the disclosure.
FIG. 1 is a schematic block diagram of an engine according to the present disclosure;
fig. 2 is a schematic view of the structure of the engine of the present disclosure, fig. two.
Detailed Description
The present disclosure will be described in further detail with reference to the drawings and embodiments. It is to be understood that the specific embodiments described herein are for purposes of illustration only and are not to be construed as limitations of the present disclosure. It should be further noted that, for the convenience of description, only the portions relevant to the present disclosure are shown in the drawings.
It should be noted that the embodiments and features of the embodiments in the present disclosure may be combined with each other without conflict. The present disclosure will be described in detail below with reference to the accompanying drawings in conjunction with embodiments.
Referring to fig. 1, an engine according to the embodiment of the present disclosure includes an engine body 1, the engine body 1 may include a casing 11, a high-pressure shaft 12 and a low-pressure shaft 13 coaxially disposed are inserted into the casing 11, the high-pressure shaft 12 and the low-pressure shaft 13 are provided with a turbine and a compressor, a main flow passage having an annular cross section is provided between the casing 11 and the shaft in the engine body 1, the main flow passage includes a cold-end flow passage 14 and a hot-end flow passage 15 sequentially communicated with each other, the cold-end flow passage 14 is located from an inlet of the engine body 1 to an inlet of a combustion chamber in the engine body, the hot-end flow passage 15 is located from the inlet of the combustion chamber to an end position of a nozzle, the inlet of the cold-end flow passage 14 is provided with a second compressor 5, and an outlet of the cold-end flow passage 14 is provided with a third compressor 6; the hot end runner 15 is sequentially provided with a main combustion chamber 7, a second turbine 8 and a third turbine 9 in a gas flowing mode. External air enters the main combustion in the hot end flow passage 15 through the cold end flow passage 14, pushes the turbine to do work after the combustion, and then is sprayed out from the afterburner, and an oil spray rod can be further installed in the afterburner.
The cold end flow channel 14 is further provided with a first flow channel 141 and a second flow channel 142, air in the cold end flow channel 14 can be communicated to the hot end flow channel 15 through the first flow channel 141 and/or the second flow channel 142, and external air can enter the hot end flow channel 15 through one or two of the first flow channel 141 and the second flow channel 142. The first flow passage 141 and the second flow passage 142 may be two separate pipes; the first flow channel 141 and the second flow channel 142 may be annular flow channels, the first flow channel 141 and the second flow channel 142 are coaxially disposed in the cold-end flow channel 14, the second flow channel 142 may be a hollow flow channel, no other component is disposed inside, the annular second flow channel 142 is disposed on an outer ring of the annular first flow channel, and the first flow channel 141 on the inner ring may be used to mount the first turbine 2.
The engine also comprises a first turbine 2 arranged in the first flow passage 141, wherein the first turbine 2 is used for depressurizing the gas in the cold end flow passage 14 and then introducing the gas into the hot end flow passage 15; the first turbine 2 may be provided with a ring along the low pressure shaft 13 in the cold end flow path 14. The first turbine 2 is arranged between the second compressor 5 and the third compressor 6.
The engine further includes a mode conversion device 3 installed in the cold-side flow passage 14, and the mode conversion device 3 can adjust the opening degree of the first flow passage 141 and adjust the opening degree of the second flow passage 142. I.e. the mode switching device 3 can adjust the open state of the flow passage.
For example, in a low-speed flight state of the aircraft, the mode switching device 3 may control the second flow passage 142 to be open, the first flow passage 141 to be closed, and the airflow to pass through the empty flow passage without pressurization or depressurization.
When the aircraft is in a high-speed flight state, the mode switching device 3 can control the first flow passage 141 to be opened, the second flow passage 142 to be closed, and the airflow enters the main combustion chamber of the hot-end flow passage 15 after being depressurized through the first turbine 2.
Further, when the aircraft is in a low-speed flight state and is in a high-speed flight state, the mode switching device 3 may further control the first flow passage 141 to be gradually opened and the second flow passage 142 to be gradually closed. When the speed of the aircraft is traveling at a speed between the low-speed flight state and the high-speed flight state, the mode switching device 3 may control and maintain the opening degrees of the first flow path 141 and the second flow path 142 in accordance with the speed ratio. The opening degrees of the first and second flow paths 141 and 142 may also be adjusted in proportion to the opening degrees of the first and second flow paths 141 and 142 empirically based on the speed of the aircraft.
Therefore, the engine of the present disclosure can adjust the opening state or the opening degree of the first flow channel 141 and the opening state or the opening degree of the second flow channel 142 according to the flight speed of the aircraft, can conveniently realize the large-range adjustment of the circulation pressure ratio, can greatly expand the flight speed limit of the engine, and can give consideration to the performance of the engine in the high and low speed states.
In some implementations, referring to fig. 2, a first compressor 4 is installed in the second flow passage 142, and the first compressor 4 is configured to boost the pressure of the gas in the cold-side flow passage 14 and then introduce the gas into the hot-side flow passage 15. By arranging the first compressor 4, in a low-speed flight state of the aircraft, the mode conversion device 3 controls the second flow passage 142 to be opened, the first flow passage 141 to be closed, and the airflow enters the main combustion chamber of the hot end flow passage 15 after being pressurized by the first compressor 4. A wide range of adjustment of the circulation pressure ratio can be further achieved.
In other embodiments, the mode switching device 3 includes a mode switching valve plate 31 and a mode switching unit (not shown in the drawings) for controlling the operation of the mode switching valve plate 31;
for example, the mode switching valve plate 31 may be a single valve plate disposed in the first flow passage 141 and the second flow passage 142, and the mode switching unit controls the valve plate in the first flow passage 141 to be opened and closed or controls the valve plate in the second flow passage 142 to be opened and closed, respectively; for another example, the control mode switching valve plate 31 may also be a single valve plate, the valve plate is switched between the first flow channel 141 and the second flow channel 142 by the mode switching unit, and when the valve plate enters the first flow channel 141, the first flow channel 141 is closed, and the second flow channel 142 is opened; when the valve plate enters the second flow channel 142, the second flow channel 142 is closed, and the first flow channel 141 is opened;
the mode switching valve sheet 31 may be disposed at the inlet positions of the first and second flow passages 141 and 142; the mode switching valve sheet 31 may also be disposed at the outlet positions of the first and second flow passages 141 and 142; and the mode switching valve sheet 31 may be further provided at both the inlet and outlet positions of the first and second flow passages 141 and 142.
In a specific embodiment, the cross sections of the first flow channel 141 and the second flow channel 142 are both annular, the first flow channel 141 and the second flow channel 142 are coaxially disposed in the cold-end flow channel 14, the annular first flow channel 141 may be disposed on the outer ring of the annular second flow channel 142, the second flow channel 142 may also be disposed on the outer ring of the annular first flow channel 141, and a partition 16 may be disposed between the first flow channel 141 and the second flow channel 142; the mode switching valve plate 31 comprises a first annular valve plate 311 and a second annular valve plate 312, the first annular valve plate 311 and the second annular valve plate 312 are connected with the mode switching device, the first annular valve plate 311 is hinged between the inlet of the first flow passage 141 and the inlet of the second flow passage 142, and the second annular valve plate 312 is hinged between the outlet of the first flow passage 141 and the outlet of the second flow passage 142. The first annular valve plate 311 and the second annular valve plate 312 can be flipped about the hinge position to open and close the first flow passage 141 or the second flow passage 142.
Referring to fig. 2, the adjustment of the opening degree of the first flow passage 141 and the adjustment of the opening degree of the second flow passage 142 are achieved by controlling the mode switching valve sheet 31, including,
when the aircraft is in a high-speed flight state, and the mode conversion valve plate 31 is in the first position, the first flow channel 141 is opened, and the second flow channel 142 is closed; that is, the mode switching valve sheet 31 is flipped down, the inlet and outlet of the first flow passage 141 are opened, and the inlet and outlet of the second flow passage 142 are closed.
When the aircraft is in a low-speed flight state, and the mode conversion valve plate 31 is in the second position, the first flow channel 141 is closed, and the second flow channel 142 is opened; that is, the mode switching valve sheet 31 is flipped up, the inlet and outlet of the first flow passage 141 are closed, and the inlet and outlet of the second flow passage 142 are opened.
When the aircraft is in a low-speed flight state and a high-speed flight state, the mode conversion valve plate 31 is located between the first position and the second position, and the proportion of the opening of the first flow channel 141 and the opening of the second flow channel 142 can be adjusted.
The disclosure also provides a control method based on the engine, which includes that a first preset speed and a second preset speed are preset, wherein the first preset speed is smaller than the second preset speed, the first preset speed range can be Mach 1.8-2.2, and the second preset speed range is Mach 2.8-3.2. In the present embodiment, the first preset speed is set to mach 2, and the second preset speed is set to mach 3.
The control method comprises the following steps:
when the speed of the aircraft is less than or equal to a first preset speed (Ma 0-2), the mode conversion device 3 controls the second flow passage 142 to be opened, and the first flow passage 141 to be closed;
when the speed of the aircraft is between a first preset speed and a second preset speed (Ma 2-3), and the speed of the aircraft is increased, increasing the opening degree of the first flow channel 141 and reducing the opening degree of the second flow channel 142; when the speed of the aircraft is between a first preset speed and a second preset speed (Ma 2-3), and the speed of the aircraft is reduced, reducing the opening degree of the first flow channel 141 and increasing the opening degree of the second flow channel 142; the opening degree of the first flow path 141 and the opening degree of the second flow path 142 may be adjusted based on the ratio of the aircraft speed between the first preset speed and the second preset speed;
when the speed of the aircraft is greater than or equal to a second preset speed (more than Ma 3), the mode switching device controls the first flow passage to be opened, and the second flow passage to be closed.
The present disclosure also provides an aircraft comprising the above-described engine. The aircraft may be an aircraft, spacecraft, rocket, missile, guided munition, or the like.
In the description herein, reference to the description of the terms "one embodiment/mode," "some embodiments/modes," "example," "specific example," or "some examples," etc., means that a particular feature, structure, material, or characteristic described in connection with the embodiment/mode or example is included in at least one embodiment/mode or example of the application. In this specification, the schematic representations of the terms used above are not necessarily intended to be the same embodiment/mode or example. Furthermore, the particular features, structures, materials, or characteristics described may be combined in any suitable manner in any one or more embodiments/modes or examples. Furthermore, the various embodiments/aspects or examples and features of the various embodiments/aspects or examples described in this specification can be combined and combined by one skilled in the art without conflicting therewith.
Furthermore, the terms "first", "second" and "first" are used for descriptive purposes only and are not to be construed as indicating or implying relative importance or implicitly indicating the number of technical features indicated. Thus, a feature defined as "first" or "second" may explicitly or implicitly include at least one such feature. In the description of the present application, "plurality" means at least two, e.g., two, three, etc., unless explicitly specified otherwise.
It will be understood by those skilled in the art that the foregoing embodiments are merely for clarity of illustration of the disclosure and are not intended to limit the scope of the disclosure. Other variations or modifications may occur to those skilled in the art, based on the foregoing disclosure, and are still within the scope of the present disclosure.

Claims (10)

1. The utility model provides an engine, its characterized in that includes the engine body, it is annular sprue to have the cross-section in the engine body, the sprue is including cold junction runner and hot junction runner that communicate in proper order, still be provided with first runner and second runner in the cold junction runner, air in the cold junction runner can pass through first runner and/or the second runner communicates to the hot junction runner, still includes:
the first turbine is arranged in the first flow passage and is used for depressurizing the gas in the cold end flow passage and then introducing the gas into the hot end flow passage;
and the mode conversion device is arranged in the cold-end flow channel and can adjust the opening degree of the first flow channel and the opening degree of the second flow channel.
2. The engine of claim 1, wherein: and a first gas compressor is arranged in the second flow channel and is used for boosting the gas in the cold end flow channel and then introducing the boosted gas into the hot end flow channel.
3. The engine of claim 2, wherein: the mode conversion device comprises a mode conversion valve plate and a mode conversion unit for controlling the mode conversion valve plate to act;
when the mode conversion valve plate is at the first position, the first flow channel is opened, and the second flow channel is closed;
when the mode conversion valve plate is at the second position, the first flow channel is closed, and the second flow channel is opened;
when the mode conversion valve plate is located between the first position and the second position, the proportion of the first flow channel opening and the second flow channel opening can be adjusted.
4. An engine as set forth in claim 3 wherein: the cross sections of the first flow channel and the second flow channel are both annular, the first flow channel and the second flow channel are coaxially arranged in the cold end flow channel, and a partition plate is arranged between the first flow channel and the second flow channel;
the mode switching valve plate comprises a first annular valve plate and a second annular valve plate, the first annular valve plate and the second annular valve plate are connected with the mode switching device, the first annular valve plate is hinged to the position between the inlet of the first flow channel and the inlet of the second flow channel, and the second annular valve plate is hinged to the position between the outlet of the first flow channel and the outlet of the second flow channel.
5. The engine of claim 4, wherein: the annular first flow passage is arranged on the outer ring of the annular second flow passage.
6. The engine of claim 1, wherein: the second flow passage is an empty flow passage, and the annular second flow passage is arranged on the outer ring of the annular first flow passage.
7. The engine of claim 1, wherein: the inlet position of the cold end runner is provided with a second gas compressor, the outlet of the cold end runner is provided with a third gas compressor, and the first turbine is arranged between the second gas compressor and the third gas compressor.
8. A control method of an engine according to any one of claims 1 to 7, characterized in that:
when the speed of the aircraft is less than or equal to a first preset speed, the mode conversion device controls the second flow channel to be opened, and the first flow channel is closed;
when the speed of the aircraft is between a first preset speed and a second preset speed and the speed of the aircraft is increased, the mode conversion device increases the opening degree of the first flow channel and reduces the opening degree of the second flow channel; when the speed of the aircraft is between a first preset speed and a second preset speed and the speed of the aircraft is reduced, the mode conversion device reduces the opening degree of the first flow channel and increases the opening degree of the second flow channel;
when the speed of the aircraft is greater than or equal to a second preset speed, the mode conversion device controls the first flow channel to be opened, and the second flow channel is closed;
the first preset speed is less than the second preset speed.
9. The control method of an engine according to claim 8, characterized in that: the first preset speed is at Mach 1.8-2.2, and the second preset speed is at Mach 2.8-3.2.
10. An aircraft, characterized in that: including an engine as claimed in any one of claims 1 to 7.
CN202210217441.1A 2022-03-07 2022-03-07 Pressure-reducing aeroengine, control method thereof and aircraft Active CN114542285B (en)

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Citations (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR2728937A1 (en) * 1994-12-28 1996-07-05 Aisin Seiki VALVE VALVE STRUCTURE FOR TURBOCHARGER
CN104500269A (en) * 2014-12-11 2015-04-08 南京航空航天大学 Self-driven fan large-bypass-ratio turbofan engine with inner loop air turbine
EP3591193A1 (en) * 2018-07-05 2020-01-08 United Technologies Corporation Failure mitigation and failure detection of intercooled cooling air systems
CN110857662A (en) * 2018-08-23 2020-03-03 波音公司 Oil cooler for air-entraining supercharged engine
CN112780441A (en) * 2019-11-05 2021-05-11 北京动力机械研究所 Variable flow passage ramjet engine and design and adjustment method
CN214741487U (en) * 2020-12-21 2021-11-16 襄阳航力机电技术发展有限公司 Natural gas turbine decompression power generation system

Patent Citations (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR2728937A1 (en) * 1994-12-28 1996-07-05 Aisin Seiki VALVE VALVE STRUCTURE FOR TURBOCHARGER
CN104500269A (en) * 2014-12-11 2015-04-08 南京航空航天大学 Self-driven fan large-bypass-ratio turbofan engine with inner loop air turbine
EP3591193A1 (en) * 2018-07-05 2020-01-08 United Technologies Corporation Failure mitigation and failure detection of intercooled cooling air systems
CN110857662A (en) * 2018-08-23 2020-03-03 波音公司 Oil cooler for air-entraining supercharged engine
CN112780441A (en) * 2019-11-05 2021-05-11 北京动力机械研究所 Variable flow passage ramjet engine and design and adjustment method
CN214741487U (en) * 2020-12-21 2021-11-16 襄阳航力机电技术发展有限公司 Natural gas turbine decompression power generation system

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