CN114542285B - Pressure-reducing aeroengine, control method thereof and aircraft - Google Patents
Pressure-reducing aeroengine, control method thereof and aircraft Download PDFInfo
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- CN114542285B CN114542285B CN202210217441.1A CN202210217441A CN114542285B CN 114542285 B CN114542285 B CN 114542285B CN 202210217441 A CN202210217441 A CN 202210217441A CN 114542285 B CN114542285 B CN 114542285B
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- 238000000034 method Methods 0.000 title claims abstract description 11
- 238000006243 chemical reaction Methods 0.000 claims abstract description 22
- 238000005192 partition Methods 0.000 claims description 3
- 230000007423 decrease Effects 0.000 claims description 2
- 125000004122 cyclic group Chemical group 0.000 abstract description 4
- 238000002485 combustion reaction Methods 0.000 description 7
- 230000000694 effects Effects 0.000 description 2
- 239000000463 material Substances 0.000 description 2
- 238000012986 modification Methods 0.000 description 2
- 230000004048 modification Effects 0.000 description 2
- 230000007704 transition Effects 0.000 description 2
- CNKHSLKYRMDDNQ-UHFFFAOYSA-N halofenozide Chemical compound C=1C=CC=CC=1C(=O)N(C(C)(C)C)NC(=O)C1=CC=C(Cl)C=C1 CNKHSLKYRMDDNQ-UHFFFAOYSA-N 0.000 description 1
- 238000002347 injection Methods 0.000 description 1
- 239000007924 injection Substances 0.000 description 1
Classifications
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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
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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
- F02C3/14—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
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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
- 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
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- 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 channel and is used for reducing the pressure of the gas in the cold end flow channel and then introducing the gas into the hot end flow channel; the mode switching device is arranged in the cold end flow channel and can adjust the opening of the first flow channel and the opening of the second flow channel. According to the method, 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 switching device can control the air flow to pass through the second flow channel, so that the pressure is not increased or reduced; when the aircraft flies at a high speed, the mode conversion device can control the air flow to pass through the first turbine of the first flow channel, and the air flow is depressurized and then is introduced into the hot end flow channel. The engine can conveniently realize the large-scale adjustment of the cyclic pressure ratio, greatly expand the flying speed limit of the engine and give consideration to the performance of the engine in a high-low speed state.
Description
Technical Field
The disclosure belongs to the technical field of aeroengines, and particularly relates to a depressurization aeroengine, a control method thereof and an aircraft.
Background
The flying speed range of the aero-engine is generally Ma 0-2, and the main purpose of the technical approach of the variable cycle engine is 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 external air flow.
However, the idea of variable cycle of the aero-engine can be used for increasing the speed and extending the flying speed range to Ma3 to 4. The air compressor is required to provide higher pressure ratio (such as 25-30) when the aeroengine flies at low speed so as to ensure higher cycle efficiency and lower oil consumption. And the supercharging effect of the air inlet channel is gradually obvious in the high-speed flight above Ma3, so that the compressor is required to provide a lower supercharging ratio (less than 10).
Disclosure of Invention
In order to solve the above technical problems, an object of the present disclosure is to provide an engine capable of generating a depressurization effect on an air flow in a high-speed state.
In order to achieve the purpose of the disclosure, the technical scheme adopted by the disclosure is as follows:
the utility model provides a step-down aeroengine, includes the engine body, have the cross-section in the engine body and be annular sprue, sprue is including cold junction runner and the hot junction runner of intercommunication 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 extremely the hot junction runner still includes:
the first turbine is arranged in the first flow channel and is used for reducing the pressure of gas in the cold end flow channel and then introducing the gas into the hot end flow channel;
and the mode conversion device is arranged in the cold end flow channel and can adjust the opening of the first flow channel and the opening of the second flow channel.
Optionally, a first air compressor is installed in the second flow channel, and the first air compressor is used for boosting the air in the cold end flow channel and then introducing the boosted air into the hot end flow channel.
Optionally, the mode conversion device comprises a mode conversion valve plate and a mode conversion unit for controlling the action of the mode conversion valve plate;
when the mode switching valve plate is at a first position, the first flow channel is opened, and the second flow channel is closed;
when the mode switching valve plate is at a second position, the first flow channel is closed, and the second flow channel is opened;
the mode switching valve plate is capable of adjusting the ratio of the first flow passage opening to the second flow passage opening when the mode switching valve plate is between a first position and a second position.
Optionally, the sections of the first flow channel and the second flow channel are 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 conversion 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 conversion device, the first annular valve plate is hinged at a position between the first flow channel inlet and the second flow channel inlet, and the second annular valve plate is hinged at a position between the first flow channel outlet and the second flow channel outlet.
Optionally, the annular first flow channel is disposed on an outer ring of the annular second flow channel.
Optionally, the second flow passage is an empty flow passage, and the annular second flow passage is disposed on an outer ring of the annular first flow passage.
Optionally, the inlet position of the cold end runner is provided with a second compressor, the outlet of the cold end runner is provided with a third compressor, and the first turbine is arranged between the second compressor and the third compressor.
The present disclosure also provides an aircraft comprising the engine
The present disclosure also provides a control method based on the above engine,
when the speed of the aircraft is smaller than or equal to a first preset speed, the mode switching device controls the second flow passage to be opened, and the first flow passage 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 of the first flow channel and decreases the opening 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 of the first flow channel and increases the opening of the second flow channel;
when the speed of the aircraft is greater than or equal to a second preset speed, the mode switching device controls the first flow passage to be opened, and the second flow passage to be closed;
the first preset speed is less than the second preset speed.
Optionally, the first preset speed is Mach 1.8-2.2, and the second preset speed is Mach 2.8-3.2.
In the method, 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 the mode conversion device can control the air flow to pass through the second flow channel without pressurization or depressurization when the aircraft flies at a low speed; when the aircraft flies at a high speed, the mode conversion device can control the air flow to pass through the first turbine of the first flow channel, and the air flow is depressurized and then is introduced into the hot end flow channel. The engine can conveniently realize the large-scale adjustment of the cyclic pressure ratio, greatly expand the flying speed limit of the engine and give consideration to the performance of the engine in a high-low speed state.
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 structural view of a step-down aircraft engine of the present disclosure;
fig. 2 is a schematic structural view of a step-down aircraft engine according to the present disclosure.
Detailed Description
The present disclosure is described in further detail below with reference to the drawings and the embodiments. It is to be understood that the specific embodiments described herein are merely illustrative of the relevant content and not limiting of the present disclosure. It should be further noted that, for convenience of description, only a portion relevant to the present disclosure is shown in the drawings.
In addition, embodiments of the present disclosure and features of the embodiments 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, a depressurization aeroengine according to an embodiment of the present disclosure includes an engine body 1, where the engine body 1 may include a housing 11, a high-pressure shaft 12 and a low-pressure shaft 13 coaxially disposed are penetrated in the housing 11, the high-pressure shaft 12 and the low-pressure shaft 13 are installed with a turbine and a compressor, a main runner with a circular section is provided between the housing 11 and the shaft in the engine body 1, the main runner includes a cold end runner 14 and a hot end runner 15 that are sequentially communicated, the cold end runner 14 is a position from an inlet position of the engine body 1 to an inlet position of a combustion chamber in the engine body, the hot end runner 15 is a position from an inlet position of the combustion chamber to an end position of a nozzle, the inlet position of the cold end runner 14 is installed with a second compressor 5, and an outlet of the cold end runner 14 is installed with a third compressor 6; the main combustion chamber 7, the second turbine 8 and the third turbine 9 are sequentially arranged in the hot end flow passage 15 along the gas flow 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 combustion, and is sprayed out of the afterburner, and an oil injection rod can be arranged 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 channel 141 and the second flow channel 142 may be two separate lines; the first flow passage 141 and the second flow passage 142 may be annular flow passages, the first flow passage 141 and the second flow passage 142 are coaxially arranged in the cold end flow passage 14, the second flow passage 142 may be an empty flow passage, no other parts are arranged inside the second flow passage 142, the annular second flow passage 142 is arranged on the annular outer ring of the first flow passage, and the first flow passage 141 of the inner ring may be used for installing the first turbine 2.
The engine further 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 depressurized gas into the hot end flow passage 15; the first turbine 2 may be positioned one revolution 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 comprises a mode switching device 3 mounted in the cold side flow passage 14, the mode switching device 3 being capable of adjusting the opening of the first flow passage 141 and adjusting the opening of the second flow passage 142. The mode switching device 3 can adjust the opening state of the flow passage.
For example, in a low speed flight condition of the aircraft, the mode switching device 3 may control the second flow passage 142 to be opened, the first flow passage 141 to be closed, and the air flow to pass through the air 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 air flow enters the main combustion chamber of the hot end flow passage 15 after being depressurized through the first turbine 2.
Further, when the aircraft transitions from the low speed flight state to the 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. The mode switching device 3 may control and maintain the opening degrees of the first and second flow passages 141 and 142 in accordance with the speed ratio when the speed of the aircraft travels at a speed between the low speed flight state and the high speed flight state. The opening degree of the first flow passage 141 and the second flow passage 142 may be adjusted according to the experience of the speed of the aircraft.
Therefore, the depressurization aeroengine disclosed by the invention can adjust the opening state or opening degree of the first flow passage 141 according to the flight speed of the aircraft, adjust the opening state or opening degree of the second flow passage 142, can conveniently realize the large-range adjustment of the cyclic pressure ratio, can greatly expand the flight speed limit of the engine, and can consider the performance of the engine in a high-low speed state.
In some implementations, referring to fig. 2, a first compressor 4 is installed in the second flow channel 142, and the first compressor 4 is used for pressurizing the gas in the cold end flow channel 14 and then introducing the gas into the hot end flow channel 15. By arranging the first air compressor 4, the mode switching device 3 controls the second flow passage 142 to be opened and the first flow passage 141 to be closed in the low-speed flight state of the aircraft, and air flow enters the main combustion chamber of the hot end flow passage 15 after being pressurized by the first air compressor 4. A wide range of adjustment of the cyclic 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 sheet 31 may be a separate valve sheet disposed in the first flow passage 141 and the second flow passage 142, the valve sheet in the first flow passage 141 being controlled to be opened and closed, or the valve sheet in the second flow passage 142 being controlled to be opened and closed, respectively, by the mode switching unit; for another example, the control mode switching valve plate 31 may be a single valve plate, and 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 passage 142, the second flow passage 142 is closed, and the first flow passage 141 is opened;
the mode switching valve sheet 31 may be disposed at inlet positions of the first and second flow passages 141 and 142; the mode switching valve sheet 31 may 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 provided at both inlet and outlet positions of the first and second flow passages 141 and 142.
In a specific embodiment, the first flow channel 141 and the second flow channel 142 are both annular in cross section, 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 an outer ring of the annular second flow channel 142, the second flow channel 142 may also be disposed on an 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 includes a first annular valve plate 311 and a second annular valve plate 312, where 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 at a position between the inlet of the first flow channel 141 and the inlet of the second flow channel 142, and the second annular valve plate 312 is hinged at a position between the outlet of the first flow channel 141 and the outlet of the second flow channel 142. The first annular valve piece 311 and the second annular valve piece 312 are each capable of being 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 plate 31, including,
when the aircraft is in a high-speed flight state, the first flow passage 141 is opened and the second flow passage 142 is closed when the mode switching valve plate 31 is in the first position; i.e., the mode switching valve sheet 31 is turned 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, the first flow passage 141 is closed and the second flow passage 142 is opened when the mode switching valve plate 31 is in the second position; i.e., the mode switching valve sheet 31 is turned upward, the inlet and outlet of the first flow path 141 are closed, and the inlet and outlet of the second flow path 142 are opened.
When the aircraft transitions from the low-speed flight state to the high-speed flight state, the mode switching valve plate 31 is located between the first position and the second position, and the ratio of the opening of the first flow passage 141 to the opening of the second flow passage 142 can be adjusted.
The disclosure further provides a control method based on the step-down aero-engine, wherein a first preset speed and a second preset speed can be preset, 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 this 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 switching 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 the first preset speed and the second preset speed (Ma 2-3), and the speed of the aircraft is increased, increasing the opening of the first flow channel 141 and reducing the opening 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 of the first flow passage 141 and increasing the opening of the second flow passage 142; the opening of the first flow passage 141 and the opening of the second flow passage 142 may be adjusted based on a 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 (Ma 3 or more), the mode switching device controls the first flow passage to be opened, and the second flow passage to be closed.
The disclosure also provides an aircraft comprising the above-described reduced pressure aircraft engine. The aircraft may be an aircraft, spacecraft, rocket, missile, guided weapon, or the like.
In the description of the present specification, reference to the terms "one embodiment/manner," "some embodiments/manner," "example," "specific example," or "some examples," etc., means that a particular feature, structure, material, or characteristic described in connection with the embodiment/manner or example is included in at least one embodiment/manner or example of the present application. In this specification, the schematic representations of the above terms are not necessarily for the same embodiment/manner 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/modes or examples described in this specification and the features of the various embodiments/modes or examples can be combined and combined by persons skilled in the art without contradiction.
Furthermore, the terms "first," "second," and the like, are used for descriptive purposes only and are not to be construed as indicating or implying a relative importance or implicitly indicating the number of technical features indicated. Thus, a feature defining "a first" or "a second" may explicitly or implicitly include at least one such feature. In the description of the present application, the meaning of "plurality" is at least two, such as two, three, etc., unless explicitly defined otherwise.
It will be appreciated by those skilled in the art that the above-described 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 will be apparent to persons skilled in the art from the foregoing disclosure, and such variations or modifications are intended to be within the scope of the present disclosure.
Claims (10)
1. The utility model provides a step-down aeroengine, its characterized in that includes the engine body, have the cross-section in the engine body and be annular sprue, sprue is including cold junction runner and the hot junction runner of intercommunication 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 extremely the hot junction runner still includes:
the first turbine is arranged in the first flow channel and is used for reducing the pressure of gas in the cold end flow channel and then introducing the gas into the hot end flow channel;
and the mode conversion device is arranged in the cold end flow channel and can adjust the opening of the first flow channel and the opening of the second flow channel.
2. The buck aircraft engine of claim 1, wherein: and a first air compressor is arranged in the second flow passage and is used for boosting the gas in the cold end flow passage and then introducing the boosted gas into the hot end flow passage.
3. The buck aircraft engine of claim 2, wherein: the mode conversion device comprises a mode conversion valve plate and a mode conversion unit for controlling the action of the mode conversion valve plate;
when the mode switching valve plate is at a first position, the first flow channel is opened, and the second flow channel is closed;
when the mode switching valve plate is at a second position, the first flow channel is closed, and the second flow channel is opened;
the mode switching valve plate is capable of adjusting the ratio of the first flow passage opening to the second flow passage opening when the mode switching valve plate is between a first position and a second position.
4. A buck aircraft engine according to claim 3, wherein: the sections of the first flow channel and the second flow channel are 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 conversion 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 conversion unit, the first annular valve plate is hinged at a position between the first flow channel inlet and the second flow channel inlet, and the second annular valve plate is hinged at a position between the first flow channel outlet and the second flow channel outlet.
5. The buck aircraft engine of claim 4, wherein: the annular first flow passage is arranged on the outer ring of the annular second flow passage.
6. The buck aircraft 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 buck aircraft engine of claim 1, wherein: the inlet position of the cold end runner is provided with a second compressor, the outlet of the cold end runner is provided with a third compressor, and the first turbine is arranged between the second compressor and the third compressor.
8. A control method based on the step-down aeroengine of any one of claims 1 to 7, characterized in that:
when the speed of the aircraft is smaller than or equal to a first preset speed, the mode switching device controls the second flow passage to be opened, and the first flow passage 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 of the first flow channel and decreases the opening 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 of the first flow channel and increases the opening of the second flow channel;
when the speed of the aircraft is greater than or equal to a second preset speed, the mode switching device controls the first flow passage to be opened, and the second flow passage to be closed;
the first preset speed is less than the second preset speed.
9. The control method of a step-down aircraft engine according to claim 8, characterized in that: the first preset speed is Mach 1.8-2.2, and the second preset speed is Mach 2.8-3.2.
10. An aircraft, characterized in that: a reduced pressure aircraft engine comprising any one of claims 1-7.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202210217441.1A CN114542285B (en) | 2022-03-07 | 2022-03-07 | Pressure-reducing aeroengine, control method thereof and aircraft |
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| Application Number | Priority Date | Filing Date | Title |
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| CN202210217441.1A CN114542285B (en) | 2022-03-07 | 2022-03-07 | Pressure-reducing aeroengine, control method thereof and aircraft |
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| CN114542285A CN114542285A (en) | 2022-05-27 |
| CN114542285B true CN114542285B (en) | 2024-03-19 |
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| FR2728937A1 (en) * | 1994-12-28 | 1996-07-05 | Aisin Seiki | VALVE VALVE STRUCTURE FOR TURBOCHARGER |
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| 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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| CN114542285A (en) | 2022-05-27 |
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