CN119309621A - A thermal protection structure for an aircraft engine measurement probe - Google Patents
A thermal protection structure for an aircraft engine measurement probe Download PDFInfo
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- CN119309621A CN119309621A CN202411361313.XA CN202411361313A CN119309621A CN 119309621 A CN119309621 A CN 119309621A CN 202411361313 A CN202411361313 A CN 202411361313A CN 119309621 A CN119309621 A CN 119309621A
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- G—PHYSICS
- G01—MEASURING; TESTING
- G01D—MEASURING NOT SPECIALLY ADAPTED FOR A SPECIFIC VARIABLE; ARRANGEMENTS FOR MEASURING TWO OR MORE VARIABLES NOT COVERED IN A SINGLE OTHER SUBCLASS; TARIFF METERING APPARATUS; MEASURING OR TESTING NOT OTHERWISE PROVIDED FOR
- G01D21/00—Measuring or testing not otherwise provided for
- G01D21/02—Measuring two or more variables by means not covered by a single other subclass
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F25—REFRIGERATION OR COOLING; COMBINED HEATING AND REFRIGERATION SYSTEMS; HEAT PUMP SYSTEMS; MANUFACTURE OR STORAGE OF ICE; LIQUEFACTION SOLIDIFICATION OF GASES
- F25D—REFRIGERATORS; COLD ROOMS; ICE-BOXES; COOLING OR FREEZING APPARATUS NOT OTHERWISE PROVIDED FOR
- F25D17/00—Arrangements for circulating cooling fluids; Arrangements for circulating gas, e.g. air, within refrigerated spaces
- F25D17/02—Arrangements for circulating cooling fluids; Arrangements for circulating gas, e.g. air, within refrigerated spaces for circulating liquids, e.g. brine
-
- G—PHYSICS
- G01—MEASURING; TESTING
- G01D—MEASURING NOT SPECIALLY ADAPTED FOR A SPECIFIC VARIABLE; ARRANGEMENTS FOR MEASURING TWO OR MORE VARIABLES NOT COVERED IN A SINGLE OTHER SUBCLASS; TARIFF METERING APPARATUS; MEASURING OR TESTING NOT OTHERWISE PROVIDED FOR
- G01D11/00—Component parts of measuring arrangements not specially adapted for a specific variable
-
- G—PHYSICS
- G01—MEASURING; TESTING
- G01D—MEASURING NOT SPECIALLY ADAPTED FOR A SPECIFIC VARIABLE; ARRANGEMENTS FOR MEASURING TWO OR MORE VARIABLES NOT COVERED IN A SINGLE OTHER SUBCLASS; TARIFF METERING APPARATUS; MEASURING OR TESTING NOT OTHERWISE PROVIDED FOR
- G01D11/00—Component parts of measuring arrangements not specially adapted for a specific variable
- G01D11/24—Housings ; Casings for instruments
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- Physics & Mathematics (AREA)
- Engineering & Computer Science (AREA)
- General Physics & Mathematics (AREA)
- Chemical & Material Sciences (AREA)
- Combustion & Propulsion (AREA)
- Mechanical Engineering (AREA)
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- General Engineering & Computer Science (AREA)
- Radiation Pyrometers (AREA)
Abstract
本发明提供了一种航空发动机测量探头热防护结构，用于对精密光学组件进行保护，包括：外壳、内壁和内壳，外壳、内壁和内壳依次同轴套设，且外壳与内壁之间以及内壁和内壳之间均形成夹层，内壳具有轴向空腔，光学组件为多个，且间隔安装在内壳的轴向空腔内；水冷通道，设置在夹层内。本发明实施例中通过布置水冷通道，以循环流动的冷却水带走热量，维持光学组件在合适的温度环境下，从而保证发动机测量探头能在高温环境下稳定工作、测量结果准确可靠。这种结构不仅提高了探头的热防护能力，还增强了其在恶劣环境下的耐用性，对于航空发动机的监测和维护具有重要意义。
The present invention provides a thermal protection structure of an aircraft engine measurement probe, which is used to protect precision optical components, including: an outer shell, an inner wall and an inner shell, wherein the outer shell, the inner wall and the inner shell are coaxially sleeved in sequence, and a sandwich is formed between the outer shell and the inner wall, and between the inner wall and the inner shell, the inner shell has an axial cavity, and there are multiple optical components, which are installed at intervals in the axial cavity of the inner shell; a water cooling channel is arranged in the sandwich. In the embodiment of the present invention, a water cooling channel is arranged to remove heat with circulating cooling water, and the optical components are maintained in a suitable temperature environment, thereby ensuring that the engine measurement probe can work stably in a high temperature environment and the measurement results are accurate and reliable. This structure not only improves the thermal protection capability of the probe, but also enhances its durability in harsh environments, which is of great significance for the monitoring and maintenance of aircraft engines.
Description
Technical Field
The invention relates to the technical field of heat protection, in particular to a heat protection structure of an aeroengine measuring probe.
Background
Aeroengines operate as heart sites in aircraft in extremely demanding environments of high rotational speed, high load and high temperature. The high-pressure turbine rotor is used as a hot end component with the worst working condition, and the interval between the high-pressure and low-pressure two-pole blades is small, so that synchronous measurement of a temperature field and a strain field of the high-pressure turbine rotor becomes very difficult, and the design, verification and fault analysis of the hot end component are hindered. In order to realize non-contact temperature/strain synchronous measurement of the hot end part of the engine, the measuring sensor needs to be designed into a probe type and extend into the engine to collect surface image information and radiation information of the hot end part. However, the front end of the sensor probe goes deep into the combustion chamber and is in a high-temperature and high-pressure environment, and meanwhile, the front end of the sensor probe faces to the scouring of polluted fuel gas, so that the working environment is extremely bad. To prevent failure of the sensor probe in such severe operating environments, thermal protection techniques at high temperatures are becoming increasingly important.
The current thermal protection technology for the aeronautical environment mainly comprises a high-temperature-resistant thermal barrier coating, an internal cooling technology and an air film cooling technology. The high temperature resistant heat barrier coating belongs to passive heat protection technology, and heat transfer is blocked by depositing a coating system with good heat insulation effect on the metal surface. The internal cooling technology and the film cooling technology belong to active heat protection technology, and heat in a heated area is taken away in a heat exchange mode by introducing cooling gas or liquid. However, the conventional heat protection structure generally adopts only passive cooling or active cooling, does not realize the combination between the two and uses the heat protection structure for large temperature difference heat protection (the external environment is up to 1500 ℃ and the internal temperature is not higher than 100 ℃).
Disclosure of Invention
In view of the above, the invention provides a thermal protection structure of an aeroengine measurement probe, so as to achieve the purpose of stably and efficiently protecting the measurement probe.
The embodiment of the invention provides a thermal protection structure of an aeroengine measuring probe, which is used for protecting an optical assembly and comprises an outer shell, an inner wall and an inner shell, wherein the outer shell, the inner wall and the inner shell are coaxially sleeved in sequence, interlayers are formed between the outer shell and the inner wall and between the inner wall and the inner shell, the inner shell is provided with a plurality of axial cavities, the optical assemblies are arranged in the axial cavities of the inner shell at intervals, and a water cooling channel is arranged at the interlayers.
Further, the water cooling channels are distributed uniformly along the circumferential direction of the interlayer at intervals or are staggered relative to the water cooling channels between the inner wall and the inner wall.
Further, a heat transfer enhancement structure is arranged in the water cooling channel and is respectively connected with the top wall and the bottom wall of the water cooling channel.
Further, a plurality of heat transfer enhancement structures arranged along the liquid flowing direction of the water cooling channel form a group of heat transfer enhancement groups, at least two groups of heat transfer enhancement groups are arranged in the water cooling channel, and the heat transfer enhancement structures between the adjacent heat transfer enhancement groups are symmetrically arranged or staggered.
Further, the outer shell, the inner wall and the inner shell form a stepped shaft-shaped protector body, the protector body is provided with a small-diameter section and a large-diameter section which are connected with each other, and an optical assembly is arranged inside the small-diameter section.
The air engine measuring probe heat protection structure further comprises a first air cooling flow passage which extends from the large-diameter section to the small-diameter section and is connected with the air film cooling structure on one side close to the large-diameter section, and a second air cooling flow passage which extends from the large-diameter section to the small-diameter section and is connected with the air film cooling structure on one side far away from the large-diameter section after passing through the end part of the small-diameter section.
Further, the first gas cooling flow passage and the second gas cooling flow passage each comprise an inlet section, an expansion section and an outlet section which are connected in sequence.
Further, the ratio of the length of the expansion section to the height of the inlet section is less than or equal to 6:1, the ratio of the height of the outlet section to the height of the inlet section is 1:5-3:5, and the ratio of the width of the outlet section to the width of the inlet section is greater than or equal to 2:1.
Further, the outer surface of the housing is coated with a high temperature resistant thermal barrier coating.
Compared with the prior art, the at least one technical scheme adopted by the embodiment of the specification has the beneficial effects that at least the water cooling channels are arranged to remove heat by circulating cooling water, so that the stability and the accuracy of the optical assembly in a high-temperature environment are maintained. The structure not only improves the heat protection capability of the probe, but also enhances the durability of the probe in severe environments, and has important significance for monitoring and maintaining the aeroengine.
Drawings
In order to more clearly illustrate the technical solutions of the embodiments of the present application, 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 application, and that other drawings can be obtained according to these drawings without inventive effort for a person skilled in the art.
FIG. 1 is a schematic view of the three-dimensional structure of the heat shield structure of the aircraft engine measurement probe of the present invention.
Fig. 2 is a Z-sectional view of a thermal protection structure for an aircraft engine measurement probe in accordance with the present invention.
Fig. 3 is a Y-sectional view of a thermal protection structure for an aircraft engine measurement probe in accordance with the present invention.
Fig. 4 is a schematic structural view of an air film cooling structure of a heat protection structure of an aeroengine measurement probe.
Fig. 5 is an enlarged view of a portion B of a thermal protection structure for an aircraft engine measurement probe according to the present invention.
FIG. 6 is a schematic view of a double-layer water cooling channel of a heat protection structure of an aircraft engine measurement probe.
FIG. 7 is a schematic structural view of a heat transfer enhancement structure of an aircraft engine measurement probe thermal protection structure of the present invention.
FIG. 8 is a schematic illustration of a staggered arrangement of heat transfer enhancement structures of a heat shield structure for an aircraft engine measurement probe in accordance with the present invention.
FIG. 9 is a schematic illustration of a symmetrical arrangement of heat transfer enhancement structures of a heat shield structure for an aircraft engine measurement probe in accordance with the present invention.
FIG. 10 is a graph of the internal wall temperature change of a probe at various water cooling temperatures in a 1500C high temperature combustor environment.
The drawing comprises the reference numerals of 1, a high-temperature-resistant heat barrier coating, 2, a gas film cooling structure, 3, a water cooling channel, 4, a heat transfer enhancement structure, 5, an outer shell, 6, an inner wall, 7, an inner shell, 8, a cooling jet flow direction, 9, a cooling water flow direction and 10, and an optical component.
Detailed Description
Embodiments of the present application will be described in detail below with reference to the accompanying drawings.
Other advantages and effects of the present application will become apparent to those skilled in the art from the following disclosure, which describes the embodiments of the present application with reference to specific examples. It will be apparent that the described embodiments are only some, but not all, embodiments of the application. The application may be practiced or carried out in other embodiments that depart from the specific details, and the details of the present description may be modified or varied from the spirit and scope of the present application. 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 those skilled in the art based on the embodiments of the application without making any inventive effort, are intended to be within the scope of the application.
As shown in fig. 1 to 10, the embodiment of the present invention provides a heat protection structure of an aeroengine measurement probe for protecting an optical assembly 10, comprising an outer shell 5, an inner wall 6, an inner shell 7 and a water cooling channel 3. The outer shell 5, the inner wall 6 and the inner shell 7 are coaxially sleeved in sequence, interlayers are formed between the outer shell 5 and the inner wall 6 and between the inner wall 6 and the inner shell 7, the inner shell 7 is provided with a plurality of axial cavities, the optical components 10 are arranged in the axial cavities of the inner shell 7 at intervals, and the water cooling channel 3 is arranged at the interlayer.
In the embodiment of the invention, the water cooling channel 3 is arranged to carry heat away by the circulating cooling water, so that the stability and the accuracy of the optical assembly 10 in a high-temperature environment are maintained. The structure not only improves the heat protection capability of the probe, but also enhances the durability of the probe in severe environments, and has important significance for monitoring and maintaining the aeroengine.
The water cooling channels 3 are a plurality of, are uniformly distributed along the circumferential direction of the interlayer at intervals or are arranged in a staggered manner relative to the water cooling channels 3 between the inner wall 6 and the inner shell 7 and the water cooling channels 3 between the outer shell 5 and the inner wall 6.
This arrangement optimizes the flow path and heat exchange efficiency of the cooling water. In this way, the cooling water can more effectively absorb and disperse heat generated inside the probe, thereby more uniformly reducing the temperature of the probe. The design not only improves the cooling efficiency of the heat protection structure, but also helps to reduce the risk of local overheating and ensures the long-term stable operation of the probe under the high-temperature and high-pressure environment. In addition, the water cooling channels arranged in a staggered manner can increase the contact area between cooling water and the inner wall of the probe, further improve the heat exchange efficiency, and have remarkable benefits for protecting sensitive optical components and improving the reliability and durability of the whole measuring system.
Preferably, in one embodiment, the water cooling device comprises 8 water cooling channels 3, each layer comprises 4 water cooling channels 3, the inside of each layer comprises a flow baffle plate, the inlet and the outlet of each water cooling channel 3 are arranged at the tail part of the probe, and better heat exchange effect can be achieved by arranging the water cooling channels so that the flow directions of the inner layer and the outer layer of adjacent flow channels are opposite, but the water cooling channels are actually used and are adjusted according to the requirement of pipeline installation.
The water cooling channel 3 is internally provided with a heat transfer enhancement structure 4, and the heat transfer enhancement structure 4 is respectively abutted with the top wall and the bottom wall of the water cooling channel 3.
The heat transfer enhancement structure 4 is skillfully arranged in the water cooling channel 3, and the heat exchange efficiency between the cooling water and the wall surface is remarkably improved by the design, because the heat transfer enhancement structure 4 increases the turbulence of the fluid, reduces the thickness of a thermal boundary layer, and accelerates the heat transfer. Such an arrangement not only optimizes the cooling effect but also helps to more evenly distribute the temperature inside the probe, preventing local overheating, which is critical for protecting sensitive optical components and ensuring the accuracy of the measurement data.
Preferably, the heat transfer enhancing structure 4 is a NACA airfoil heat transfer enhancing structure. The plurality of heat transfer enhancement structures 4 arranged along the liquid flow direction of the water cooling channel 3 form a group of heat transfer enhancement groups, at least two groups of heat transfer enhancement groups are arranged in the water cooling channel 3, and the heat transfer enhancement structures 4 between the adjacent heat transfer enhancement groups are symmetrically arranged or staggered.
The streamline design of NACA airfoil reduces fluid flow resistance, increases turbulence, effectively breaks the thermal boundary layer, and promotes rapid heat transfer. In addition, by staggering or symmetrically arranging these airfoil structures within the water cooling channels 3, fluid flow is further optimized, flow dead zones are reduced, and cooling water is ensured to flow uniformly through the entire channels.
In one embodiment of the present invention, the ratio of the horizontal spacing between adjacent heat transfer enhancement structures 4 along the cooling water flow direction 9 to the length of the structures is not less than 1:1, and the aspect ratio is 10:1-10:3. Such a design optimizes hydrodynamic performance and heat exchange efficiency. The larger horizontal spacing helps to reduce flow disturbances that may occur between adjacent structures, thereby reducing flow resistance and maintaining fluid flow continuity and uniformity. Meanwhile, the moderate length-width ratio ensures that the structure can maximally disturb fluid, enhance turbulence and effectively improve heat exchange area and efficiency on the premise of not increasing excessive resistance.
As shown in fig. 1, the outer housing 5, the inner wall 6 and the inner housing 7 form a stepped shaft-like protector body having a small diameter section and a large diameter section connected to each other, and an optical assembly 10 is provided inside the small diameter section. Correspondingly, the small-diameter section comprises an air film cooling structure 2 which is identical in structure and arranged in a mirror image mode, and the aeroengine measuring probe heat protection structure further comprises:
The first gas cooling flow passage extends from the large-diameter section to the small-diameter section and is connected with the gas film cooling structure 2 close to one side of the large-diameter section;
The second gas cooling flow passage extends from the large-diameter section to the small-diameter section, passes through the end part of the small-diameter section and is connected with the gas film cooling structure 2 at one side far away from the large-diameter section.
Through forming even and complete air film in probe front end light window department (path section department), effectively isolated and sweep away high temperature gas and pollutant, protection light window does not receive erosion and pollution. The structure design ensures that cooling gas is sprayed at a specific angle to form an isolation layer with high-temperature fuel gas, reduces direct heat conduction and convection heat exchange, and simultaneously can prevent fuel gas pollutants contained in the engine from adhering to the light window through the blowing action of the air film, so that the cleanness and transparency of the light window are maintained, and the measurement accuracy and reliability of the measuring probe are ensured. In addition, the design of the air film cooling structure 2 is also beneficial to reducing the temperature of the front end of the probe, improving the durability and stability of the probe in an extremely high-temperature environment, and has important significance for guaranteeing the safe operation and performance monitoring of the aeroengine.
Preferably, the optical assembly 10 in the embodiment of the present invention further includes a plurality of optical lenses disposed at the large diameter section, the plurality of optical lenses being disposed at intervals in the axial direction.
Furthermore, the invention can realize the further optimized cooling of the integral device by arranging the first gas cooling flow passage and the second gas cooling flow passage, realize the cooperative cooling of gas cooling and water cooling and improve the cooling efficiency of the integral device.
The first gas cooling flow passage and the second gas cooling flow passage comprise an inlet section, an expansion section and an outlet section which are connected in sequence, the ratio of the length of the expansion section to the height of the inlet section is less than or equal to 6:1, the ratio of the height of the outlet section to the height of the inlet section is 1:5-3:5, and the ratio of the width of the outlet section to the width of the inlet section is greater than or equal to 2:1.
The ratio of the length of the expansion section to the height of the inlet section is less than or equal to 6:1, and such a design facilitates uniform gas film coverage of the gas within the flow passage while avoiding efficiency degradation due to excessive expansion. The ratio of the height of the outlet section to the height of the inlet section is controlled between 1:5 and 3:5, which ensures that the gas flow can be effectively controlled to accommodate different cooling requirements while maintaining the gas film integrity. The ratio of the width of the outlet section to the width of the inlet section is greater than or equal to 2:1, which helps to form a wider coverage area at the outlet and enhances the protection effect of the light window at the front end of the probe. These precise proportional controls enable the film cooling structure to optimize the distribution and use efficiency of the cooling gas while providing effective thermal protection, reducing the consumption of the cooling gas.
Preferably, the inlet cross section of the film cooling structure 2 can be shaped like a groove or a cylinder according to actual requirements. Compared with the groove-shaped structure, the cylindrical structure is difficult to process, but has more uniform air film covering effect and more excellent cooling effect at low air flow, and the groove-shaped structure is very similar to the air film covering effect and the cooling effect of the cylindrical structure at high air flow. When in actual use, the excessive gas flow can influence the external gas environment to influence the measurement of the probe, so that the selection of the cross section shape is required according to the influence degree of the cooling gas on the external gas.
It should be noted that, in the embodiment of the present invention, the gas enters from the end of the probe according to the cooling jet direction 8 and enters the corresponding film cooling structure 2 through the first gas cooling flow channel and the second gas cooling flow channel respectively, and the outlet jet direction of the gas cooling structure is 90 ° with the gas flushing direction.
The outer surface of the outer envelope 5 is coated with a high temperature resistant thermal barrier coating 1. Specifically, the high temperature resistant thermal barrier coating 1 comprises a ceramic layer (TC layer) and a bonding layer (BC layer), wherein the ceramic layer is made of 6-8wt% of yttria partially stabilized zirconia material, pyrochlore material LMA material or perovskite type material, and the bonding layer is MNiCrAl alloy component shapes, wherein M can be Co, fe, Y and other elements.
The high-temperature-resistant thermal barrier coating 1 can effectively block heat transfer of external high-temperature fuel gas, and reduces influence of heat on the optical component 10 in the probe. By selecting materials with excellent insulating properties, such as partially stabilized zirconia, pyrochlore or perovskite type materials, the coating is able to withstand extreme temperatures up to 1500 ℃ while maintaining a low thermal conductivity, thereby establishing a temperature gradient between the probe surface and the internal components. In addition, the bonding layer of the coating adopts MNiCrAl alloy components, so that the bonding force between the coating and the substrate is enhanced, and the stability and durability of the coating in a high-temperature environment are ensured. The application of the high-temperature-resistant thermal barrier coating 1 remarkably improves the thermal protection capability of the measuring probe in a severe environment, prolongs the service life of the measuring probe, and ensures the accuracy and reliability of the measurement data of the aero-engine.
Fig. 10 shows the trend of the internal wall temperature change of the probe under the condition of 1500 ℃ high temperature combustion chamber at different water cooling temperatures, and fig. 10 shows that the temperature of the inner wall surface of the probe is increased from 52.7 ℃ to 73.0 ℃ and does not exceed 100 ℃ when the temperature of cooling water is 6.0 ℃ to 19.8 ℃, so that the temperature-resistant requirement of an internal optical component is met.
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 application is subject to the protection scope of the claims.
Claims (9)
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| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202411361313.XA CN119309621A (en) | 2024-09-27 | 2024-09-27 | A thermal protection structure for an aircraft engine measurement probe |
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| Application Number | Priority Date | Filing Date | Title |
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| CN202411361313.XA CN119309621A (en) | 2024-09-27 | 2024-09-27 | A thermal protection structure for an aircraft engine measurement probe |
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| CN113280341A (en) * | 2021-05-25 | 2021-08-20 | 浙江大学 | Combustion chamber spray gun with air film hole |
| CN215415938U (en) * | 2021-06-05 | 2022-01-04 | 西斐（上海）工业控制有限公司 | Water-cooled optical detection structure |
| CN115623744A (en) * | 2022-10-10 | 2023-01-17 | 西安交通大学 | Spiral hot protective structure |
| CN116294703A (en) * | 2023-03-16 | 2023-06-23 | 广州大学 | A finned tube heat exchanger |
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2024
- 2024-09-27 CN CN202411361313.XA patent/CN119309621A/en active Pending
Patent Citations (9)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JP2009128011A (en) * | 2007-11-19 | 2009-06-11 | Central Res Inst Of Electric Power Ind | Cooling structure of optical measurement probe |
| CN106969872A (en) * | 2017-04-18 | 2017-07-21 | 北京航空航天大学 | A kind of pressure probe of use double-row hole gaseous film control |
| CN110231093A (en) * | 2019-07-02 | 2019-09-13 | 电子科技大学 | A kind of infrared radiation temperature meter reducing meter background radiation |
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| CN116294703A (en) * | 2023-03-16 | 2023-06-23 | 广州大学 | A finned tube heat exchanger |
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