CN114151137B - Combined cooling thermal management system for high Mach number aeroengine cabin and turbine disk - Google Patents
Combined cooling thermal management system for high Mach number aeroengine cabin and turbine disk Download PDFInfo
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- CN114151137B CN114151137B CN202111222096.2A CN202111222096A CN114151137B CN 114151137 B CN114151137 B CN 114151137B CN 202111222096 A CN202111222096 A CN 202111222096A CN 114151137 B CN114151137 B CN 114151137B
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- air
- heat exchanger
- turbine
- control valve
- fuel oil
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Classifications
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
- F01D5/12—Blades
- F01D5/14—Form or construction
- F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
- F01D5/185—Liquid cooling
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
- F01D25/08—Cooling; Heating; Heat-insulation
- F01D25/12—Cooling
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
- F01D5/12—Blades
- F01D5/14—Form or construction
- F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
- F01D5/187—Convection cooling
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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)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Engine Equipment That Uses Special Cycles (AREA)
Abstract
The invention provides a combined cooling and thermal management system of a high Mach number aeroengine cabin and a turbine disc, which comprises an engine air main flow path, wherein the engine air main flow path comprises a cold air inlet channel, a gas compressor, a main combustion chamber, a turbine and an afterburner which are sequentially communicated in the air flow direction, and also comprises an air-air heat exchanger and an air-oil heat exchanger, wherein the cold air inlet end of the air-air heat exchanger is communicated with the cold air inlet channel, the hot air inlet end of the air-air heat exchanger is communicated with the hot air outlet end of the gas compressor, the hot air outlet end of the air-air heat exchanger is communicated with the hot air inlet end of the air-oil heat exchanger, the air-oil heat exchanger further comprises a cold fuel oil input end, and the air output end of the air-oil heat exchanger is connected with the cooling air inlet end of the turbine. The thermal management system can simultaneously solve the problems that the temperature in the cabin of the aeroengine in the high-speed aircraft is high, accessories cannot work normally, and the temperature of the air bleed of the air compressor is high, and the turbine disc is over-temperature.
Description
Technical Field
The disclosure relates to the technical field of aeroengine thermal management, in particular to a high Mach number aeroengine cabin and turbine disc combined cooling thermal management system.
Background
As aircraft flying speeds increase, thermal problems become more and more pronounced in aircraft engine designs. The role of thermal management systems in aircraft engines is becoming increasingly important, being an important factor in ensuring that the various systems of the engine work properly. However, conventional simpler thermal management schemes have failed to meet the operational requirements of existing high speed flight conditions. The turbine disk cooling bleed air adopts compressor outlet air, and because the compressor outlet temperature is higher and higher, if the air entering the turbine disk is not cooled, the turbine material temperature resistance level is limited at the moment, the turbine disk overtemperature problem exists, and meanwhile, the cooling requirement of the movable blades cannot be met. The temperature in the engine cabin is higher and higher along with the higher and higher Mach number of the flight, and the normal working temperature requirements of accessories of the engine cannot be met.
According to the prior art, when solving the turbine disk cooling problem, the prior art generally adopts a heat exchanger to cool the air entering the turbine disk by using fuel oil or external air, or uses an air cooler with liquid nitrogen or helium cooling to cool the air, thereby increasing the complexity of the thermal management system. Another solution is to use the extra air to cool the incoming turbine disc air, which is only applicable to relatively low speed aircraft. In addressing the problem of excessive temperatures in the engine compartment, it is desirable to use an oil-to-gas heat exchanger in the ram combustion chamber circuit to cool the high mach number ram air prior to entering the engine.
In existing thermal management system solutions, separate cooling solutions are generally employed for cooling the engine compartment and for turbine disk super-temperature cooling. For the thermal management scheme at high Mach number, most of the thermal management schemes adopt an air-oil heat exchanger, and the heat sinks of the fuel oil are used for cooling engine cabin cooling bleed air and turbine disc cooling bleed air respectively, so that the problem is that the temperature of the fuel oil is too high, and the risk of fuel oil over-temperature coking exists.
Disclosure of Invention
In view of the above, the embodiments of the present disclosure provide a combined cooling thermal management system for an aircraft engine nacelle and a turbine disk with high mach numbers, which can solve the problems that the temperature in the aircraft engine nacelle is high, accessories cannot work normally, and the temperature of the air bleed gas of a compressor is high, and the turbine disk is over-heated.
In order to achieve the above object, the present invention provides the following technical solutions:
the utility model provides a high Mach number aeroengine cabin and turbine dish joint cooling thermal management system, includes engine air mainstream flow path, engine air mainstream flow path includes air conditioning intake duct, compressor, main combustion chamber, turbine and afterburner that communicate in proper order in the air flow direction, still includes air-air heat exchanger and air oil heat exchanger, air conditioning intake end of air-air heat exchanger with air conditioning intake duct intercommunication, air-air heat exchanger's steam intake end with the hot air exhaust end of compressor communicates, air-air heat exchanger's hot air exhaust end with the hot air intake end of air oil heat exchanger communicates, air oil heat exchanger still includes cold fuel input, air output of air oil heat exchanger connects the cooling air intake end of turbine.
Further, the air turbine pump is arranged between the air inlet channel and the air-air heat exchanger, the air inlet end of the air turbine pump is communicated with the air inlet channel, and the first air outlet end of the air turbine pump is communicated with the air inlet end of the air-air heat exchanger.
Further, the engine compartment is further included, and the second exhaust end of the air turbine pump is communicated with the engine compartment.
Further, the air conditioner further comprises a first control valve, wherein the first control valve is arranged on a communicating pipeline between the cold air inlet channel and the air turbine pump.
Further, the air-to-air heat exchanger further comprises a second control valve, wherein the second control valve is arranged on a communication pipeline between the air compressor and the air-to-air heat exchanger.
Further, the air compressor further comprises a third control valve, the hot air exhaust end of the air compressor is further communicated with the cooling air inlet end of the turbine, and the third control valve is arranged on the communicating pipeline.
Further, the cold fuel oil input end of the air-fuel oil heat exchanger comprises an oil tank, a booster pump, a high-flow oil path part, an accessory and a fuel oil heat exchanger which are sequentially connected in the fuel oil flowing direction, and the output end of the fuel oil heat exchanger is connected with the cold fuel oil input end of the air-fuel oil heat exchanger.
Further, the fuel oil distributor is characterized by further comprising a fuel oil distributor, wherein the input end of the fuel oil distributor is connected with the fuel oil output end of the air-fuel heat exchanger, and the output end of the fuel oil distributor is connected with the afterburner.
Further, the air-oil heat exchanger further comprises a pre-spinning nozzle and a turbine disc cavity, wherein the pre-spinning nozzle and the turbine disc cavity are sequentially arranged on a connecting pipeline between the air output end of the air-oil heat exchanger and the cooling air inlet end of the turbine.
Further, the air-air heat exchanger further comprises a cold air exhaust end, and the cold air exhaust end is connected to the external environment.
The invention can solve the problems of overhigh engine room and over-temperature turbine disk during high-speed flight simultaneously on the basis of adding the air turbine pump, the air-air heat exchanger and the air-oil heat exchanger, and has the advantages of easy implementation, easy transformation and high efficiency.
The air in the air inlet channel is cooled by the air turbine pump, and one path of the air enters the engine cabin to realize cooling in the engine cabin, so that the problem of over-temperature of the engine cabin at a high Mach number is solved. The other path of cooling gas of the air turbine pump enters the air-air heat exchanger to realize preliminary cooling of the bleed air of the air compressor, and then is discharged to the external environment. After the bleed air of the air compressor is primarily cooled by the air-air heat exchanger, the final cooling of the bleed air of the air compressor is realized by utilizing a heat sink of fuel oil through the air-oil heat exchanger, so that the turbine disc is cooled.
In addition, after the air turbine pump is added, 200K temperature drop of air inlet bleed air can be realized in a high-speed flight state. After the air-air heat exchanger and the air-oil heat exchanger are added, the temperature of the outlet of the air compressor can be reduced by 100K in a high-speed flight state, and the cooling requirement of the turbine disk is met. The thermal load of the fuel can be reduced by approximately 20K compared with the prior thermal management scheme.
Drawings
In order to more clearly illustrate the technical solutions of the embodiments of the present disclosure, the drawings that are needed in the embodiments will be briefly described below, and it is obvious that the drawings in the following description are only some embodiments of the present disclosure, and other drawings may be obtained according to these drawings without inventive effort to a person of ordinary skill in the art.
FIG. 1 is a schematic illustration of an air flow path in a thermal management system according to an embodiment of the present invention;
FIG. 2 is a schematic diagram of a fuel flow path in a thermal management system according to an embodiment of the invention.
Detailed Description
Embodiments of the present disclosure are described in detail below with reference to the accompanying drawings.
Other advantages and effects of the present disclosure will become readily apparent to those skilled in the art from the following disclosure, which describes embodiments of the present disclosure by way of specific examples. It will be apparent that the described embodiments are merely some, but not all embodiments of the present disclosure. The disclosure may be embodied or practiced in other different specific embodiments, and details within the subject specification may be modified or changed from various points of view and applications without departing from the spirit of the disclosure. It should be noted that the following embodiments and features in the embodiments may be combined with each other without conflict. All other embodiments, which can be made by one of ordinary skill in the art without inventive effort, based on the embodiments in this disclosure are intended to be within the scope of this disclosure.
It is noted that various aspects of the embodiments are described below within the scope of the following claims. It should be apparent that the aspects described herein may be embodied in a wide variety of forms and that any specific structure and/or function described herein is merely illustrative. Based on the present disclosure, one skilled in the art will appreciate that one aspect described herein may be implemented independently of any other aspect, and that two or more of these aspects may be combined in various ways. For example, an apparatus may be implemented and/or a method practiced using any number of the aspects set forth herein. In addition, such apparatus may be implemented and/or such methods practiced using other structure and/or functionality in addition to one or more of the aspects set forth herein.
It should also be noted that the illustrations provided in the following embodiments merely illustrate the basic concepts of the disclosure by way of illustration, and only the components related to the disclosure are shown in the drawings and are not drawn according to the number, shape and size of the components in actual implementation, and the form, number and proportion of the components in actual implementation may be arbitrarily changed, and the layout of the components may be more complicated.
In addition, in the following description, specific details are provided in order to provide a thorough understanding of the examples. However, it will be understood by those skilled in the art that the aspects may be practiced without these specific details.
As shown in fig. 1-2, embodiments of the present disclosure provide a high mach number aircraft nacelle and turbine disk combined cooling thermal management system comprising an air turbine pump 2, an air-to-air heat exchanger 6, and an air-to-oil heat exchanger 7; the air turbine pump 2 is arranged after the air is introduced into the air inlet channel 8; the air turbine pump 2 has two air outlets, one into the engine compartment 3 and one into the external environment 5.
When the aero-engine flies at a low speed, the first control valve 1 and the second control valve 4 are closed, and the third control valve 20 is opened; when the aeroengine flies at a high Mach number, the first control valve 1 and the second control valve 4 are opened, and the third control valve 20 is closed;
the bleed air in the air inlet channel 8 is cooled through the air turbine pump 2, so that the engine compartment 3 is cooled, and the problem of overhigh temperature of the engine compartment 3 is solved. An air-air heat exchanger 6 is arranged in the cabin, and an air-oil heat exchanger 7 is arranged at the lower side of the casing in the main combustion chamber of the engine to solve the problem of overtemperature of the cavity of the turbine disc 17. The specific technical measures are that when the high Mach number flight is carried out, the first control valve 1 and the second control valve 4 are opened, the third control valve 20 is closed, the air turbine pump 2 is adopted to cool the air introduced into the air inlet channel 8, the outlet of the air turbine pump 2 is divided into two paths, one path enters the engine compartment 3 to cool the engine compartment 3, the other path enters the air-air heat exchanger 6 to primarily cool the air introduced by the air compressor 9, and then the air is discharged to the external environment 5. For the bleed air of the air compressor 9, after the air-air heat exchanger 6 is used for cooling by utilizing the outlet of the air turbine pump 2, the final cooling of the bleed air of the air compressor 9 is realized by utilizing the heat sink of the fuel oil through the air-oil heat exchanger 7, so that the cavity of the turbine disc 17 is cooled.
When the air turbine pump 2 does not exist in high Mach number flight, the temperature of incoming air of the air inlet channel 8 is too high, and the temperature in the engine compartment 3 is high, so that the normal working temperature requirement of an accessory cannot be met. Meanwhile, when the air-air heat exchanger 6 and the air-oil heat exchanger 7 are not present, the air bleed air of the air compressor 7 passes through the pre-rotation nozzle 18 and finally enters the turbine disc cavity 17, and at the moment, the outlet temperature of the air compressor is too high and exceeds the temperature resistance level of the turbine disc.
In response to this problem, the present embodiment provides a high Mach number aircraft nacelle and turbine disk combined cooling thermal management system. The specific implementation mode is as follows:
(1) When the aero-engine flies at a low speed, the first control valve 1 and the second control valve 4 are closed, the third control valve 20 is opened, at the moment, the system scheme does not work, the air flow path is the same as that of a conventional engine, the air flows into the air compressor 9 from the air inlet channel 8, the main flow of the air compressor enters the main combustion chamber 10, then enters the turbine 11, and then enters the afterburner 12. In addition, the gas introduced by the compressor directly enters the turbine 11 for cooling.
(2) During high mach number flight of the aeroengine, the first control valve 1 and the second control valve 4 are opened and the third control valve 20 is closed, at which time the system scheme works.
(3) In the schematic air flow path in the thermal management system of fig. 1, for the bleed air flow path of the air inlet channel, the bleed air of the air inlet channel 8 enters the air turbine pump 2 for cooling, the outlet of the air turbine pump is divided into two paths, one path enters the engine compartment 3 for cooling the interior of the engine compartment, the other path enters the air-air heat exchanger 6 for primarily cooling the bleed air of the air compressor 9, and then is discharged to the external environment 5;
(4) For the air flow path of the thermal management system of fig. 1, the air bleed of the air compressor 9 firstly enters the air-air heat exchanger 6, after the air bleed is primarily cooled by the air turbine pump 2 through the air inlet channel 8, enters the air-oil heat exchanger 7, and after the air bleed of the air compressor 9 is cooled by the heat sink of the high-flow fuel, the air bleed flows through the pre-rotation nozzle 18, enters the turbine disc cavity 17 and finally enters the turbine 11;
(5) In the fuel flow path of the thermal management system of fig. 2, an air-fuel heat exchanger 7 is added in the fuel flow path of the fuel system, and the fuel in the fuel tank 13 passes through the booster pump 14 and then enters the large-flow oil path part to form an accessory 15, then flows into the fuel-oil heat exchanger 16, flows into the air-fuel heat exchanger 7, flows into the fuel distributor 19, and finally enters the afterburner 12.
The foregoing is merely specific embodiments of the disclosure, but the protection scope of the disclosure is not limited thereto, and any changes or substitutions that can be easily conceived by those skilled in the art within the technical scope of the disclosure are intended to be covered by the protection scope of the disclosure. Therefore, the protection scope of the present disclosure shall be subject to the protection scope of the claims.
Claims (5)
1. The utility model provides a high Mach number aeroengine cabin and turbine dish joint cooling thermal management system, includes engine air mainstream flow path, engine air mainstream flow path includes air conditioning intake duct, compressor, main combustion chamber, turbine and afterburner that communicate in proper order in the air flow direction, its characterized in that: the air-to-air heat exchanger is characterized by further comprising an air-to-air heat exchanger and an air-to-oil heat exchanger, wherein a cold air inlet end of the air-to-air heat exchanger is communicated with the cold air inlet channel, a hot air inlet end of the air-to-air heat exchanger is communicated with a hot air exhaust end of the air compressor, a hot air exhaust end of the air-to-air heat exchanger is communicated with a hot air inlet end of the air-to-oil heat exchanger, the air-to-oil heat exchanger further comprises a cold fuel oil input end, and an air output end of the air-to-oil heat exchanger is connected with a cooling air inlet end of the turbine;
the air turbine pump is arranged between the cold air inlet channel and the air-air heat exchanger, the air inlet end of the air turbine pump is communicated with the cold air inlet channel, and the first air outlet end of the air turbine pump is communicated with the cold air inlet end of the air-air heat exchanger;
the engine room is further included, and the second exhaust end of the air turbine pump is communicated with the engine room;
the air turbine pump also comprises a first control valve, wherein the first control valve is arranged on a communicating pipeline between the cold air inlet channel and the air turbine pump;
the air-to-air heat exchanger also comprises a second control valve, wherein the second control valve is arranged on a communication pipeline between the air compressor and the air-to-air heat exchanger;
the hot air exhaust end of the air compressor is also communicated with the cooling air inlet end of the turbine, and the communicating pipeline is provided with the third control valve;
when the aeroengine flies at a low speed, the first control valve and the second control valve are closed, and the third control valve is opened;
and when the aeroengine runs at a high Mach number, the first control valve and the second control valve are opened, and the third control valve is closed.
2. The high mach number aircraft nacelle and turbine disk combined cooling thermal management system of claim 1, wherein: the cold fuel oil input end of the air-fuel oil heat exchanger comprises an oil tank, a booster pump, a high-flow oil way part, an accessory and a fuel oil heat exchanger which are sequentially connected in the fuel oil flowing direction, and the output end of the fuel oil heat exchanger is connected with the cold fuel oil input end of the air-fuel oil heat exchanger.
3. The high mach number aircraft nacelle and turbine disk combined cooling thermal management system of claim 2, wherein: the fuel oil distributor is characterized by further comprising a fuel oil distributor, wherein the input end of the fuel oil distributor is connected with the fuel oil output end of the air-fuel heat exchanger, and the output end of the fuel oil distributor is connected with the afterburner.
4. A high mach number aircraft nacelle and turbine disc combined cooling thermal management system as recited in claim 3 wherein: the air-oil heat exchanger further comprises a pre-spinning nozzle and a turbine disc cavity, wherein the pre-spinning nozzle and the turbine disc cavity are sequentially arranged on a connecting pipeline between the air output end of the air-oil heat exchanger and the cooling air inlet end of the turbine.
5. The high mach number aircraft nacelle and turbine disk combined cooling thermal management system of claim 1, wherein: the air-air heat exchanger further comprises a cold air exhaust end, and the cold air exhaust end is connected to the external environment.
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| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202111222096.2A CN114151137B (en) | 2021-10-20 | 2021-10-20 | Combined cooling thermal management system for high Mach number aeroengine cabin and turbine disk |
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| Application Number | Priority Date | Filing Date | Title |
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| CN202111222096.2A CN114151137B (en) | 2021-10-20 | 2021-10-20 | Combined cooling thermal management system for high Mach number aeroengine cabin and turbine disk |
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| CN114151137A CN114151137A (en) | 2022-03-08 |
| CN114151137B true CN114151137B (en) | 2023-09-05 |
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| CN117905586B (en) * | 2024-03-19 | 2024-05-17 | 中国航发沈阳发动机研究所 | Lubricating oil cooling system and method for turbine support bearing of aero-engine |
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|---|---|
| CN114151137A (en) | 2022-03-08 |
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