CN119160380A - A channel actively cooled leading edge structure for high-speed aircraft - Google Patents
A channel actively cooled leading edge structure for high-speed aircraft Download PDFInfo
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- CN119160380A CN119160380A CN202411671480.4A CN202411671480A CN119160380A CN 119160380 A CN119160380 A CN 119160380A CN 202411671480 A CN202411671480 A CN 202411671480A CN 119160380 A CN119160380 A CN 119160380A
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- cooling
- pipeline
- leading edge
- wing
- units
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- 238000001816 cooling Methods 0.000 claims abstract description 96
- 239000002826 coolant Substances 0.000 claims abstract description 25
- 238000007599 discharging Methods 0.000 claims description 3
- 230000007423 decrease Effects 0.000 claims description 2
- 239000000463 material Substances 0.000 abstract description 5
- 230000000694 effects Effects 0.000 description 5
- 238000010438 heat treatment Methods 0.000 description 5
- 238000005192 partition Methods 0.000 description 5
- 239000000498 cooling water Substances 0.000 description 4
- 229910001069 Ti alloy Inorganic materials 0.000 description 3
- 239000000956 alloy Substances 0.000 description 2
- 229910052751 metal Inorganic materials 0.000 description 2
- 239000002184 metal Substances 0.000 description 2
- 239000000654 additive Substances 0.000 description 1
- 230000000996 additive effect Effects 0.000 description 1
- 230000015572 biosynthetic process Effects 0.000 description 1
- 230000007613 environmental effect Effects 0.000 description 1
- 230000007774 longterm Effects 0.000 description 1
- 238000004519 manufacturing process Methods 0.000 description 1
- 150000002739 metals Chemical class 0.000 description 1
- 239000010453 quartz Substances 0.000 description 1
- 230000001105 regulatory effect Effects 0.000 description 1
- VYPSYNLAJGMNEJ-UHFFFAOYSA-N silicon dioxide Inorganic materials O=[Si]=O VYPSYNLAJGMNEJ-UHFFFAOYSA-N 0.000 description 1
- 238000006467 substitution reaction Methods 0.000 description 1
- 230000008646 thermal stress Effects 0.000 description 1
- 238000009827 uniform distribution Methods 0.000 description 1
- 230000008016 vaporization Effects 0.000 description 1
- 238000009834 vaporization Methods 0.000 description 1
Classifications
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64C—AEROPLANES; HELICOPTERS
- B64C3/00—Wings
- B64C3/36—Structures adapted to reduce effects of aerodynamic or other external heating
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64C—AEROPLANES; HELICOPTERS
- B64C3/00—Wings
- B64C3/10—Shape of wings
- B64C3/14—Aerofoil profile
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64C—AEROPLANES; HELICOPTERS
- B64C3/00—Wings
- B64C3/10—Shape of wings
- B64C3/14—Aerofoil profile
- B64C3/141—Circulation Control Airfoils
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64C—AEROPLANES; HELICOPTERS
- B64C3/00—Wings
- B64C3/28—Leading or trailing edges attached to primary structures, e.g. forming fixed slots
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64C—AEROPLANES; HELICOPTERS
- B64C3/00—Wings
- B64C3/10—Shape of wings
- B64C3/14—Aerofoil profile
- B64C2003/143—Aerofoil profile comprising interior channels
Landscapes
- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- Aviation & Aerospace Engineering (AREA)
- Physics & Mathematics (AREA)
- Fluid Mechanics (AREA)
Abstract
本申请属于散热器配置技术领域，特别涉及一种用于高速飞机的通道主动冷却前缘结构，包括：分布在机翼上下翼面内侧的多个上冷却管路单元与下冷却管路单元；与多个上冷却管路单元进口以及多个下冷却管路单元进口连接的进口管路；分别与多个上冷却管路单元出口连接的上出口管路；分别与多个下冷却管路单元出口连接的下出口管路；机翼前缘冷却管路，沿展向布置于机翼最前缘，其入口与进口管路连接，出口与上出口管路或/和下出口管路连接；其中，冷却介质从进口管路流入，由上出口管路与下出口管路流出，本申请相较于被动热防护结构，本申请可以选用较为轻质的材料，从而在保证温度可控的情况下，减轻结构重量，提高承载能力。
The present application belongs to the technical field of radiator configuration, and particularly relates to a channel active cooling leading edge structure for high-speed aircraft, comprising: a plurality of upper cooling pipe units and a lower cooling pipe unit distributed on the inner sides of the upper and lower wing surfaces of the wing; an inlet pipe connected to the inlets of the plurality of upper cooling pipe units and the inlets of the plurality of lower cooling pipe units; an upper outlet pipe connected to the outlets of the plurality of upper cooling pipe units respectively; a lower outlet pipe connected to the outlets of the plurality of lower cooling pipe units respectively; a wing leading edge cooling pipe arranged at the leading edge of the wing along the span direction, with its inlet connected to the inlet pipe, and its outlet connected to the upper outlet pipe or/and the lower outlet pipe; wherein the cooling medium flows in from the inlet pipe and flows out from the upper outlet pipe and the lower outlet pipe. Compared with a passive thermal protection structure, the present application can select lighter materials, thereby reducing the structural weight and improving the bearing capacity while ensuring that the temperature is controllable.
Description
Technical Field
The application belongs to the technical field of radiator configuration, and particularly relates to a channel active cooling front edge structure for a high-speed aircraft.
Background
When flying at high altitude and high speed, high-speed aircraft have surfaces subjected to severe pneumatic heating, which presents a great challenge to the temperature resistance and load carrying capacity of the structure. The front edge of the wing is used as a part with very serious pneumatic heating, and is locally subjected to megawatt heat flow, the equilibrium temperature can reach more than six hundred ℃, and common materials are difficult to use. At the same time, the high temperature can reduce the strength limit and the bearing capacity of the material, and the structure of the aircraft can be damaged when the high temperature is serious, so that the safe operation of the aircraft is influenced, and a certain heat protection design is needed.
Disclosure of Invention
In order to solve the above problems, the present application provides a channel active cooling leading edge structure for a high-speed aircraft, comprising:
A plurality of upper cooling pipeline units distributed on the inner side of the upper wing surface of the wing;
a plurality of lower cooling pipeline units distributed on the inner side of the lower wing surface of the wing;
the inlet pipeline is positioned on the inner side of the front edge of the wing and is respectively connected with a plurality of inlets of the upper cooling pipeline units and a plurality of inlets of the lower cooling pipeline units;
an upper outlet pipe connected to the outlets of the plurality of upper cooling pipe units, respectively;
lower outlet pipelines respectively connected with the outlets of the plurality of lower cooling pipeline units;
the wing front edge cooling pipeline is arranged at the forefront edge of the wing along the spanwise direction, an inlet of the wing front edge cooling pipeline is connected with the inlet pipeline, and an outlet of the wing front edge cooling pipeline is connected with the upper outlet pipeline or/and the lower outlet pipeline;
wherein the cooling medium flows in from the inlet pipeline and flows out from the upper outlet pipeline and the lower outlet pipeline.
Preferably, the plurality of upper cooling pipeline units are located on the same plane to form an upper cooling network, the plurality of lower cooling pipeline units are located on the same plane to form a lower cooling network, and the upper cooling network, the lower cooling network and the inlet pipeline form a wedge-shaped structure.
Preferably, each upper cooling line unit and each lower cooling line unit comprises a plurality of mutually parallel sub-channels arranged in a reverse fractal configuration structure, wherein the sub-channels are divided into a plurality of stages from front to back along the chord direction, and each two sub-channels of the upper stage are converged into a sub-channel of the lower stage.
Preferably, the sub-channels meet at an angle of 60 °.
Preferably, all sub-channels are rectangular in cross-section.
Preferably, the ratio of the width of the sub-channel of the upper stage to the width of the sub-channel of the lower stage is 1:1.4.
Preferably, the adjacent upper cooling pipe units are parallel to each other, and the adjacent lower cooling pipe units are parallel to each other.
Preferably, the inlet line is connected to a main inlet line for supplying the cooling medium, and the upper outlet line and the lower outlet line are each connected to a main outlet line for discharging the cooling medium.
Preferably, the cross-sectional area of the wing leading edge cooling pipeline is gradually reduced along the flow direction of the cooling medium, the cross-sectional areas of the upper outlet pipeline and the lower outlet pipeline are gradually increased along the flow direction of the cooling medium, and the main output pipeline is connected with the position with the largest cross-sectional areas of the upper outlet pipeline and the lower outlet pipeline.
The advantages of the application include:
1. Compared with a passive heat protection structure, the application can select lighter materials, thereby reducing the weight of the structure and improving the bearing capacity under the condition of ensuring controllable temperature.
2. The application can ensure that the structure operates at a lower temperature in a high-temperature environment, and realizes the application of the low-temperature resistant material in the high-temperature environment.
3. Compared with other wing leading edge structure schemes, the application ensures that the whole wing leading edge structure has good temperature uniformity and lower thermal stress level.
4. The application has small flow resistance, more uniform distribution of cooling effect on the front edge, difficult vaporization of cooling medium, higher utilization efficiency, no flow stagnation area and difficult blockage.
The present invention thus provides significant advantages over existing solutions in high speed flight leading edge structure applications and in heat carrying characteristics.
Drawings
FIG. 1 is a perspective view of a channel actively cooled leading edge structure for a high speed aircraft in accordance with a preferred embodiment of the present application.
Detailed Description
In order to make the technical solution of the present application and its advantages more clear, the technical solution of the present application will be further and completely described in detail with reference to the accompanying drawings, it being understood that the specific embodiments described herein are only some of the embodiments of the present application and are for explanation of the present application and not for limitation of the present application. It should be noted that, for convenience of description, only the part related to the present application is shown in the drawings, and other related parts may refer to the general design, and the embodiments of the present application and the technical features of the embodiments may be combined with each other to obtain new embodiments without conflict.
As shown in fig. 1, in order to solve the above problems, the present application provides a channel active cooling leading edge structure for a high-speed aircraft, comprising:
A plurality of upper cooling line units 7 distributed inside the upper wing surface of the wing;
a plurality of lower cooling line units 8 distributed inside the lower wing surface of the wing;
an inlet pipeline 5 is positioned on the inner side of the front edge of the wing and is respectively connected with inlets of a plurality of upper cooling pipeline units 7 and inlets of a plurality of lower cooling pipeline units 8;
an upper outlet pipe 3 connected to outlets of the plurality of upper cooling pipe units 7, respectively;
a lower outlet pipe 2 connected to outlets of the plurality of lower cooling pipe units 8, respectively;
The wing leading edge cooling pipeline 6 is arranged at the forefront edge of the wing along the spanwise direction, the inlet of the wing leading edge cooling pipeline is connected with the inlet pipeline 5, and the outlet of the wing leading edge cooling pipeline is connected with the upper outlet pipeline 3 or/and the lower outlet pipeline 2;
wherein the cooling medium flows in from the inlet pipe 5 and flows out from the upper outlet pipe 3 and the lower outlet pipe 2.
Preferably, the upper cooling pipeline units 7 are positioned on the same plane to form an upper cooling network, the lower cooling pipeline units 8 are positioned on the same plane to form a lower cooling network, the upper cooling network, the lower cooling network and the inlet pipeline 5 form a wedge-shaped structure, the whole body is in mirror symmetry along a central line, the inlet pipeline 5 is arranged on a wing closer to the aircraft nose and is positioned near the windward side of the front edge of the wing, the upper outlet pipeline 3 and the lower outlet pipeline 2 are arranged at the rear part far away from the aircraft nose, the purpose is that the temperature of cooling medium at an inlet is lower so that the cooling medium preferentially contacts the front edge of the wing with higher temperature, and most of the areas of the upper surface and the lower surface of the front edge structure are used for cooling the whole heat load of the front edge structure.
The wing leading edge cooling duct 6 is a separate path of cooling passage located at the forefront for cooling the high heat load at the forefront.
Preferably, each upper cooling line unit 7 and lower cooling line unit 8 comprises a plurality of mutually parallel sub-channels arranged in a reverse fractal configuration, which sub-channels are divided into a plurality of stages from front to back in the chord direction, and each two sub-channels of the upper stage merge into a sub-channel of the lower stage.
Preferably, all the sub-channels are rectangular in cross section, and the upper outlet duct 3, the lower outlet duct 2, the wing leading edge cooling duct 6 and the inlet duct 5 are rectangular, so that the contact area with the inner surface of the wing can be increased.
Preferably, the ratio of the width of the sub-channel of the upper stage to the width of the sub-channel of the lower stage is 1:1.4, the included angle of the sub-channels at the junction is 60 degrees, the adjacent upper cooling pipeline units 7 are parallel to each other, the adjacent lower cooling pipeline units 8 are parallel to each other, the formation of a flow stagnation area can be avoided, and the flow resistance is small.
Preferably, the inlet line 5 is connected to a main inlet line 4 for supplying the cooling medium, and the upper outlet line 3 and the lower outlet line 2 are each connected to a main outlet line 1 for discharging the cooling medium.
Preferably, the cross-sectional area of the wing leading edge cooling pipeline 6 gradually decreases along the flow direction of the cooling medium, the cross-sectional areas of the upper outlet pipeline 3 and the lower outlet pipeline 2 gradually increase along the flow direction of the cooling medium, and the main output pipeline 1 is connected with the position where the cross-sectional areas of the upper outlet pipeline 3 and the lower outlet pipeline 2 are the largest.
The cooling medium enters the inlet pipeline 5 and the wing front edge cooling pipeline 6 from the main input pipeline 4, the cooling medium of the inlet pipeline 5 sequentially enters the upper cooling pipeline unit 7 and the lower cooling pipeline unit 8 respectively, the cooling medium of the upper cooling pipeline unit 7 and the cooling medium of the lower cooling pipeline unit 8 sequentially enter the upper outlet pipeline 3 and the lower outlet pipeline 2 respectively, and the wing front edge cooling pipeline 6 also enters the upper outlet pipeline 3 and the lower outlet pipeline 2 and finally is discharged through the main output pipeline 1.
The performance of the actively cooled leading edge structure is analyzed in connection with specific embodiments below. To verify that the structure has good temperature control effect, in a specific embodiment, an additive manufacturing is used, and a test piece of metal is manufactured by using a TC4 titanium alloy material.
In order to verify the cooling characteristics of the channel actively cooled leading edge structure, a test system was constructed, and 12 heated quartz lamps were arranged on the surface of the test piece to heat the test piece to simulate pneumatic heating. The upper and lower surfaces of the test piece, which are close to the front edge half part, are divided into a first partition, the upper and lower surfaces of the test piece, which are close to the rear edge half part, are divided into a second partition, three lamps are respectively arranged on the upper and lower surfaces of the first partition and the second partition, the power of the lamps in each partition is equal, 5L of cooling water is adopted as cooling medium, the supply flow rate of the cooling medium is regulated to be 3.5L/min, the test piece is supplied with cooling water, the temperature of the cooling water inlet and outlet is measured through a thermocouple, the cooling water flow rate is measured through a flowmeter, and the temperatures of a plurality of positions on the surface of the active cooling front edge structure are measured through the thermocouple.
According to the test result, under the effect of channel cooling, the temperatures of different positions of the front edge structure can be effectively controlled below 320 ℃, the temperature is far lower than the high temperature of 700 ℃ which can be achieved by environmental heating, a good structure temperature control effect is achieved, and meanwhile, the temperatures of the structure surface temperature measuring points are similar, and the fact that the channel active cooling front edge structure has a good temperature equalizing effect is also proved.
The test result also shows that under the severe heating condition, the TC4 titanium alloy material (the upper limit of the long-term temperature resistance is 400 ℃) can be stably applied in a high-temperature environment, and compared with other high-temperature resistant metals, the TC4 titanium alloy has lower density, so that the scheme has the potential of reducing the weight of the structure.
The wing leading edge cooling pipeline 6 has stronger heat exchange capacity, meanwhile, the phenomenon that a flow stagnation area appears on the leading edge with concentrated heat load is avoided, and the matching of the cooling capacity and the heat load and the efficient heat exchange are realized.
Thus, by this example, it is further demonstrated that the channel actively cooled leading edge structure proposed by the present invention has excellent performance and practicality.
The foregoing is merely illustrative of the present application, and the present application is not limited thereto, and any changes or substitutions easily contemplated by those skilled in the art within the scope of the present application should be included in the present application. Therefore, the protection scope of the present application shall be subject to the protection scope of the claims.
Claims (10)
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202411671480.4A CN119160380A (en) | 2024-11-21 | 2024-11-21 | A channel actively cooled leading edge structure for high-speed aircraft |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202411671480.4A CN119160380A (en) | 2024-11-21 | 2024-11-21 | A channel actively cooled leading edge structure for high-speed aircraft |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CN119160380A true CN119160380A (en) | 2024-12-20 |
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Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN202411671480.4A Pending CN119160380A (en) | 2024-11-21 | 2024-11-21 | A channel actively cooled leading edge structure for high-speed aircraft |
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| Country | Link |
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| CN (1) | CN119160380A (en) |
Citations (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN1773080A (en) * | 2004-11-09 | 2006-05-17 | 联合工艺公司 | Heat transferring cooling features for an airfoil |
| CN114162308A (en) * | 2021-12-28 | 2022-03-11 | 沈阳飞机设计研究所扬州协同创新研究院有限公司 | Structure for cooling wing leading edge shell |
| US20220185446A1 (en) * | 2020-12-14 | 2022-06-16 | General Electric Company | System and method for cooling a leading edge of a high speed vehicle |
| CN115108019A (en) * | 2021-03-23 | 2022-09-27 | 空中客车德国运营有限责任公司 | Heated leading edge structure for an aircraft, slat, wing and aircraft |
| CN117734930A (en) * | 2023-12-19 | 2024-03-22 | 南昌大学 | An active cooling structure for the leading edge of hypersonic aircraft |
-
2024
- 2024-11-21 CN CN202411671480.4A patent/CN119160380A/en active Pending
Patent Citations (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN1773080A (en) * | 2004-11-09 | 2006-05-17 | 联合工艺公司 | Heat transferring cooling features for an airfoil |
| US20220185446A1 (en) * | 2020-12-14 | 2022-06-16 | General Electric Company | System and method for cooling a leading edge of a high speed vehicle |
| CN115108019A (en) * | 2021-03-23 | 2022-09-27 | 空中客车德国运营有限责任公司 | Heated leading edge structure for an aircraft, slat, wing and aircraft |
| US20220306301A1 (en) * | 2021-03-23 | 2022-09-29 | Airbus Operations Gmbh | Heated leading edge structure for an aircraft |
| CN114162308A (en) * | 2021-12-28 | 2022-03-11 | 沈阳飞机设计研究所扬州协同创新研究院有限公司 | Structure for cooling wing leading edge shell |
| CN117734930A (en) * | 2023-12-19 | 2024-03-22 | 南昌大学 | An active cooling structure for the leading edge of hypersonic aircraft |
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