Thermal protection structure of aero-engine measuring probe

Abstract

The invention provides an aero-engine measuring probe thermal protection structure, which is used for protecting precise optical components, and comprises an outer shell, an inner wall and an inner shell, the outer shell, the inner wall and the inner shell are sequentially and coaxially sleeved, an interlayer is formed between the outer shell and the inner wall and an interlayer is formed between the inner wall and the inner shell, the inner shell is provided with an axial cavity, and a plurality of optical components are arranged in the axial cavity. The inner shell and the outer shell are arranged in the axial cavity of the inner shell at intervals; and the water cooling channel is arranged in the interlayer. According to the embodiment of the invention, by arranging the water cooling channel, heat is taken away by circularly flowing cooling water, and the optical assembly is maintained in a proper temperature environment, so that the engine measuring probe can stably work in a high-temperature environment, and the measuring result is accurate and reliable. The structure not only improves the thermal protection capability of the probe, but also enhances the durability of the probe in a severe environment, and is of great significance to monitoring and maintenance of an aero-engine.

Description

Technical Field

[0001] The invention relates to the technical field of thermal protection, and in particular to a thermal protection structure of an aero-engine measuring probe. Background Art

[0002] As the heart of an aircraft, an aircraft engine works in an extremely harsh environment of high speed, high load and high temperature. As the hot end component with the worst working conditions, the high-pressure turbine rotor has a small gap between the high and low pressure blades, which makes it very difficult to synchronously measure its temperature field and strain field, hindering the design, verification and fault analysis of the hot end components. In order to achieve non-contact temperature / strain synchronous measurement of the hot end components of the engine, it is necessary to design the measurement sensor into a "probe type" and extend it into the engine to collect the surface image information and radiation information of the hot end components. However, the front end of the sensor probe is deep inside the combustion chamber, in a high temperature and high pressure environment, and is also facing the scouring of polluting fuel gas, and the working environment is extremely harsh. In order to prevent the sensor probe from failing in such a harsh working environment, its thermal protection technology at high temperature becomes increasingly important.

[0003] The thermal protection technologies currently used in aeroengine environments mainly include high-temperature resistant thermal barrier coatings, internal cooling technology, and air film cooling technology. High-temperature resistant thermal barrier coatings are passive thermal protection technologies that block the transfer of heat by depositing a coating system with good thermal insulation effect on the metal surface. Internal cooling technology and air film cooling technology are active thermal protection technologies that introduce cooling gas or liquid to remove heat from the heated area in a heat exchange manner. However, traditional thermal protection structures generally only use passive cooling or active cooling, and do not achieve a combination of the two and use them for large temperature difference thermal protection (external environment up to 1500°C, internal temperature not higher than 100°C). Summary of the invention

[0004] In view of this, the present invention provides a thermal protection structure for an aircraft engine measuring probe, so as to achieve the purpose of providing stable and efficient protection for the measuring probe.

[0005] An embodiment of the present invention provides a thermal protection structure of an aircraft engine measurement probe for protecting an optical component, comprising: 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 an interlayer 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 arranged at intervals in the axial cavity of the inner shell; and a water cooling channel is arranged at the interlayer.

[0006] Furthermore, there are multiple water cooling channels, which are evenly spaced along the circumference of the interlayer, or the water cooling channel between the outer shell and the inner wall is staggered relative to the water cooling channel between the inner wall and the inner shell.

[0007] Furthermore, a heat transfer enhancement structure is provided in the water cooling channel, and the heat transfer enhancement structure is respectively connected to the top wall and the bottom wall of the water cooling channel.

[0008] Furthermore, a plurality of heat transfer enhancement structures arranged along the liquid flow direction of the water cooling channel form a heat transfer enhancement group, at least two heat transfer enhancement groups are arranged in the water cooling channel, and the heat transfer enhancement structures between adjacent heat transfer enhancement groups are symmetrically arranged or staggered.

[0009] Furthermore, the outer shell, the inner wall and the inner shell form a stepped shaft-shaped protective device body, the protective device body has a small diameter section and a large diameter section connected to each other, and an optical component is arranged inside the small diameter section.

[0010] Furthermore, the small diameter section includes an air film cooling structure with the same structure and a mirror-arranged arrangement; the thermal protection structure of the aircraft engine measurement probe also includes: a first gas cooling flow channel, extending from the large diameter section to the small diameter section, and connected to the air film cooling structure on the side close to the large diameter section; a second gas cooling flow channel, extending from the large diameter section to the small diameter section, and after passing through the end of the small diameter section, connected to the air film cooling structure on the side away from the large diameter section.

[0011] Furthermore, the first gas cooling channel and the second gas cooling channel each include an inlet section, an expansion section and an outlet section which are connected in sequence.

[0012] Furthermore, 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.

[0013] Furthermore, the outer surface of the shell is coated with a high temperature resistant thermal barrier coating.

[0014] Compared with the prior art, the beneficial effects achieved by at least one of the above technical solutions adopted in the embodiments of this specification include at least the following: in the embodiments of the present invention, by arranging a water cooling channel, the circulating cooling water takes away the heat, thereby maintaining the stability and accuracy of the optical components in a high temperature environment. 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. BRIEF DESCRIPTION OF THE DRAWINGS

[0015] In order to more clearly illustrate the technical solutions of the embodiments of the present application, the drawings required for use in the embodiments will be briefly introduced below. Obviously, the drawings described below are only some embodiments of the present application. For ordinary technicians in this field, other drawings can be obtained based on these drawings without paying creative work.

[0016] Figure 1The present invention is a schematic diagram of the overall three-dimensional structure of a thermal protection structure of an aero-engine measuring probe.

[0017] Figure 2 The invention discloses a Z-direction cross-sectional view of a thermal protection structure of an aero-engine measuring probe.

[0018] Figure 3 It is a Y-direction cross-sectional view of a thermal protection structure of an aero-engine measuring probe of the present invention.

[0019] Figure 4 The present invention is a structural schematic diagram of an air film cooling structure of a thermal protection structure of an aero-engine measuring probe.

[0020] Figure 5 It is an enlarged view of a part B of the thermal protection structure of an aero-engine measuring probe of the present invention.

[0021] Figure 6 The present invention is a schematic structural diagram of a double-layer water cooling channel of a thermal protection structure of an aero-engine measuring probe.

[0022] Figure 7 The present invention is a schematic structural diagram of a heat transfer enhancement structure of a thermal protection structure of an aero-engine measurement probe.

[0023] Figure 8 The present invention is a schematic diagram of the staggered arrangement of a heat protection structure and a heat transfer enhancement structure of an aero-engine measurement probe.

[0024] Fig. 9 The present invention is a schematic diagram of the symmetrical arrangement of a heat transfer enhancement structure of a thermal protection structure of an aero-engine measurement probe.

[0025] Fig.10 This is a graph showing the internal wall temperature variation of the probe at different water cooling temperatures in a 1500°C high temperature combustion chamber environment.

[0026] Reference numerals in the figure: 1. high temperature resistant thermal barrier coating; 2. air film cooling structure; 3. water cooling channel; 4. heat transfer enhancement structure; 5. outer shell; 6. inner wall; 7. inner shell; 8. cooling jet direction; 9. cooling water flow direction; 10. optical component. DETAILED DESCRIPTION

[0027] The embodiments of the present application are described in detail below with reference to the accompanying drawings.

[0028] The following describes the implementation methods of the present application through specific examples, and those skilled in the art can easily understand other advantages and effects of the present application from the contents disclosed in this specification. Obviously, the described embodiments are only a part of the embodiments of the present application, rather than all the embodiments. The present application can also be implemented or applied through other different specific implementation methods, and the details in this specification can also be modified or changed in various ways based on different viewpoints and applications without departing from the spirit of the present application. It should be noted that, in the absence of conflict, the following embodiments and the features in the embodiments can be combined with each other. Based on the embodiments in the present application, all other embodiments obtained by ordinary technicians in the field without making creative work belong to the scope of protection of the present application.

[0029] like Figures 1 to 10 As shown, an embodiment of the present invention provides a thermal protection structure of an aircraft engine measurement probe, which is used to protect an optical component 10, and includes 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, and a sandwich is 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 has an axial cavity, and there are multiple optical components 10, which are arranged in the axial cavity of the inner shell 7 at intervals; the water cooling channel 3 is arranged at the sandwich.

[0030] In the embodiment of the present invention, the water cooling channel 3 is arranged to remove heat with circulating cooling water, thereby maintaining the stability and accuracy of the optical component 10 in a high temperature environment. 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.

[0031] There are multiple water cooling channels 3, which are evenly spaced along the circumference of the interlayer, or the water cooling channels 3 between the outer shell 5 and the inner wall 6 are staggered relative to the water cooling channels 3 between the inner wall 6 and the inner shell 7.

[0032] This layout optimizes the flow path of the cooling water and the heat exchange efficiency. In this way, the cooling water can more effectively absorb and disperse the heat generated inside the probe, thereby reducing the temperature of the probe more evenly. This design not only improves the cooling efficiency of the thermal protection structure, but also helps reduce the risk of local overheating and ensure the long-term stable operation of the probe in high temperature and high pressure environments. In addition, the staggered water cooling channels can increase the contact area between the cooling water and the inner wall of the probe, further improving the heat exchange efficiency, which has significant benefits for protecting sensitive optical components and improving the reliability and durability of the entire measurement system.

[0033] Preferably, in one embodiment, a total of 8 water-cooling channels 3 are included, each layer includes 4 water-cooling channels 3, and a baffle is included inside. The inlet and outlet of the water-cooling channel 3 are arranged at the tail of the probe. The flow directions of adjacent flow channels in the inner and outer layers can be arranged to be opposite to achieve a better heat exchange effect. However, actual use still needs to be adjusted according to the requirements of pipeline installation.

[0034] A heat transfer enhancement structure 4 is provided in the water cooling channel 3 , and the heat transfer enhancement structure 4 abuts against the top wall and the bottom wall of the water cooling channel 3 respectively.

[0035] A heat transfer enhancement structure 4 is cleverly arranged inside the water cooling channel 3. This design significantly improves the heat exchange efficiency between the cooling water and the wall surface, because the heat transfer enhancement structure 4 increases the turbulence of the fluid and reduces the thickness of the thermal boundary layer, thereby accelerating the heat transfer. Such a layout not only optimizes the cooling effect, but also helps to more evenly distribute the temperature inside the probe and prevent local overheating, which is crucial for protecting sensitive optical components and ensuring the accuracy of measurement data.

[0036] Preferably, the heat transfer enhancement structure 4 is a NACA airfoil heat transfer enhancement structure. A plurality of heat transfer enhancement structures 4 arranged along the liquid flow direction of the water cooling channel 3 form a heat transfer enhancement group, at least two heat transfer enhancement groups are arranged in the water cooling channel 3, and the heat transfer enhancement structures 4 between adjacent heat transfer enhancement groups are symmetrically arranged or staggered.

[0037] The streamlined design of the NACA airfoil reduces fluid flow resistance, increases turbulence, effectively destroys the thermal boundary layer, and promotes rapid heat transfer. In addition, by staggering or symmetrically arranging these airfoil structures in the water cooling channel 3, the fluid flow is further optimized, the flow dead zone is reduced, and it is ensured that the cooling water can flow evenly through the entire channel.

[0038] In a specific embodiment of the 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 structure length is not less than 1:1, and the aspect ratio is 10:1-10:3. Such a design optimizes the fluid dynamics and heat exchange efficiency. A larger horizontal spacing helps to reduce the flow interference that may occur between adjacent structures, thereby reducing the flow resistance and maintaining the continuity and uniformity of the fluid flow. At the same time, a moderate aspect ratio ensures that the structure maximizes the disturbance of the fluid without increasing too much resistance, enhances the turbulence, and effectively improves the heat exchange area and efficiency.

[0039] like Figure 1As shown, the outer shell 5, the inner wall 6 and the inner shell 7 form a stepped shaft-shaped protective device body, which has a small diameter section and a large diameter section connected to each other, and an optical component 10 is arranged inside the small diameter section. Correspondingly, the small diameter section includes an air film cooling structure 2 with the same structure and mirror arrangement; the thermal protection structure of the aircraft engine measurement probe also includes:

[0040] The first gas cooling flow channel extends from the large diameter section to the small diameter section and is connected to the air film cooling structure 2 close to the large diameter section;

[0041] The second gas cooling flow channel extends from the large diameter section to the small diameter section, and after passing through the end of the small diameter section, is connected to the air film cooling structure 2 on the side away from the large diameter section.

[0042] By forming a uniform and complete air film at the optical window (small diameter section) at the front end of the probe, the high-temperature fuel gas and pollutants are effectively isolated and purged, protecting the optical window from erosion and pollution. This structural design allows the cooling gas to be sprayed at a specific angle to form an isolation layer with the high-temperature fuel gas, reducing direct heat conduction and convection heat exchange. At the same time, through the purging effect of the air film, the fuel gas pollutants contained in the engine can be prevented from adhering to the optical window, maintaining the cleanliness and transparency of the optical window, and ensuring the measurement accuracy and reliability of the measuring probe. In addition, the design of the air film cooling structure 2 also helps to reduce the temperature of the front end of the probe, improve its durability and stability in extremely high temperature environments, and is of great significance for ensuring the safe operation and performance monitoring of aircraft engines.

[0043] Preferably, the optical component 10 in the embodiment of the present invention further comprises a plurality of optical lenses arranged in the large diameter section, and the plurality of optical lenses in the large diameter section are arranged at intervals along the axial direction.

[0044] Furthermore, the present invention can achieve further optimized cooling of the overall device by providing the first gas cooling channel and the second gas cooling channel, realize coordinated cooling of air cooling and water cooling, and improve the cooling efficiency of the overall device.

[0045] The first gas cooling channel and the second gas cooling channel each include an inlet section, an expansion section and an outlet section 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.

[0046] The ratio of the length of the expansion section to the height of the inlet section is less than or equal to 6:1. This design helps the gas form a uniform air film coverage in the flow channel while avoiding the decrease in efficiency caused by 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 while maintaining the integrity of the air film, the gas flow rate can be effectively controlled to adapt to different cooling needs. 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 enhance the protection of the optical window at the front end of the probe. These precise proportional controls enable the air film cooling structure to optimize the distribution and use efficiency of the cooling gas while providing effective thermal protection, thereby reducing the consumption of cooling gas.

[0047] Preferably, the inlet cross-section of the air film cooling structure 2 can be selected in a groove shape, cylindrical shape or the like according to actual needs. Compared with the groove structure, the cylindrical structure is more difficult to process, but has a more uniform air film coverage effect and a better cooling effect at low gas flow rates; while the groove structure has an air film coverage effect and cooling effect very close to that of the cylindrical structure at high flow rates. In actual use, excessive gas flow will affect the external gas environment and the measurement of the probe, so it is necessary to select the cross-sectional shape according to the degree of influence of the cooling gas on the external gas.

[0048] 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 passes through the first gas cooling channel and the second gas cooling channel respectively to enter the corresponding air film cooling structure 2, and its outlet jet direction is 90° to the gas flushing direction.

[0049] The outer surface of the shell 5 is coated with a high temperature resistant thermal barrier coating 1. Specifically, the high temperature resistant thermal barrier coating 1 includes a ceramic layer (TC layer) and a bonding layer (BC layer), the ceramic layer material is a 6-8wt% yttria partially stabilized zirconia material, a pyrochlore material LMA material or a perovskite material, and the bonding layer material is a MNiCrAl alloy component, wherein M can be an element such as Co, Fe, or Y.

[0050] The high-temperature resistant thermal barrier coating 1 can effectively block the heat transfer of external high-temperature combustion gas and reduce the impact of heat on the optical components 10 inside the probe. By selecting materials with excellent thermal insulation properties, such as partially stabilized zirconium oxide, pyrochlore materials or perovskite materials, the coating can withstand extreme temperatures of up to 1500°C while maintaining a low thermal conductivity, thereby establishing a temperature gradient between the probe surface and internal components. In addition, the coating's bonding layer uses MNiCrAl alloy components to enhance the bonding between the coating and the substrate, ensuring the stability and durability of the coating in high-temperature environments. The application of this high-temperature resistant thermal barrier coating 1 significantly improves the thermal protection capability of the measuring probe in harsh environments, extends its service life, and ensures the accuracy and reliability of aircraft engine measurement data.

[0051] Fig.10 This is the change trend of the internal wall temperature of the probe at different water cooling temperatures in a 1500℃ high temperature combustion chamber environment. Fig.10 It shows that when the cooling water temperature is 6.0℃-19.8℃, the temperature of the inner wall of the probe rises from 52.7℃ to 73.0℃, not exceeding 100℃, which far meets the temperature resistance requirements of the internal optical components.

[0052] The above is only a specific implementation of the present application, but the protection scope of the present application is not limited thereto. Any changes or substitutions that can be easily thought of by a person skilled in the art within the technical scope disclosed in the present application should be included in the protection scope of the present application. Therefore, the protection scope of the present application shall be based on the protection scope of the claims.

Claims

1. A thermal protection structure of an aircraft engine measurement probe, used for protecting an optical component (10), characterized in that: include: An outer shell (5), an inner wall (6) and an inner shell (7), wherein the outer shell (5), the inner wall (6) and the inner shell (7) are coaxially sleeved in sequence, and sandwich layers 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) has an axial cavity, and a plurality of optical components (10) are arranged in intervals in the axial cavity of the inner shell (7); A water cooling channel (3) is arranged at the interlayer.

2. The thermal protection structure of the aircraft engine measurement probe according to claim 1, characterized in that: There are a plurality of water cooling channels (3), which are evenly spaced along the circumference of the interlayer, or the water cooling channels (3) between the outer shell (5) and the inner wall (6) are staggered relative to the water cooling channels (3) between the inner wall (6) and the inner shell (7).

3. The thermal protection structure of the aircraft engine measurement probe according to claim 1, characterized in that: A heat transfer enhancement structure (4) is provided in the water cooling channel (3), and the heat transfer enhancement structure (4) is respectively connected to the top wall and the bottom wall of the water cooling channel (3).

4. The thermal protection structure of the aircraft engine measurement probe according to claim 3, characterized in that: A 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 the heat transfer enhancement groups are arranged in the water cooling channel (3), and the heat transfer enhancement structures (4) between adjacent heat transfer enhancement groups are symmetrically arranged or staggered.

5. The thermal protection structure of the aircraft engine measurement probe according to claim 1, characterized in that: The outer shell (5), the inner wall (6) and the inner shell (7) form a stepped shaft-shaped protective device body, the protective device body having a small diameter section and a large diameter section connected to each other, and an optical component (10) is arranged inside the small diameter section.

6. The thermal protection structure of the aircraft engine measurement probe according to claim 5, characterized in that: The small diameter section includes an air film cooling structure (2) with the same structure and mirror-image arrangement; the aero-engine measurement probe thermal protection structure also includes: A first gas cooling flow channel extends from the large diameter section to the small diameter section and is connected to the air film cooling structure (2) close to one side of the large diameter section; The second gas cooling flow channel extends from the large diameter section to the small diameter section and is connected to the air film cooling structure (2) at a side away from the large diameter section after passing through the end of the small diameter section.

7. The thermal protection structure of the aircraft engine measurement probe according to claim 6, characterized in that: The first gas cooling channel and the second gas cooling channel each include an inlet section, an expansion section and an outlet section which are connected in sequence.

8. The thermal protection structure of the aircraft engine measurement probe according to claim 7, characterized in that: 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.

9. The thermal protection structure of the aircraft engine measurement probe according to claim 1, characterized in that: The outer surface of the housing (5) is coated with a high-temperature resistant thermal barrier coating (1).

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