A thermal protection structure for an aero-engine measurement probe with film gas vents

Abstract

The application provides an aero-engine measuring probe thermal protection structure with gas film holes, which comprises a flange plate, an outer shell and an inner shell, the outer shell and the inner shell are fixedly arranged on one side of the flange plate, the outer shell and the inner shell are arranged at intervals and form a sandwich space for cold gas input, and a plurality of gas film holes are arranged on the outer shell at intervals; and a gas guide pipeline is arranged on the other side of the flange plate, and the gas guide pipeline is communicated with the sandwich space. By introducing the cold gas in the compressor and arranging a plurality of gas film holes on the outer shell, a protective gas film is formed to isolate the probe from the high-temperature gas, so that the temperature of the probe is effectively reduced.

Description

Technical Field

[0001] This invention relates to the field of optical probe protection devices in high-temperature environments, and specifically to a thermal protection structure for an aero-engine measurement probe with an air film aperture. Background Technology

[0002] In the aviation, energy, and industrial sectors, gas turbines serve as core power equipment, and their performance and reliability are crucial for the stable operation of the entire system. As a key component of aero-engines, the turbine blades endure immense thermal and mechanical stresses in high-temperature, high-pressure, and high-speed operating environments, making them prone to deformation and wear, which in turn affects the overall performance and safety of the engine. Therefore, accurate measurement of turbine blade deformation is of great significance for assessing their health status, preventing failures, and optimizing performance.

[0003] To monitor the deformation of gas turbine blades in real time, optical measurement technology is commonly used in modern aero-engine design. Optical measurement technology offers advantages such as high precision, non-contact operation, and good real-time performance, enabling dynamic monitoring of blade deformation. However, when optical measurement probes operate in the high-temperature flow field of a gas turbine, they face the impact of extreme temperatures and high-speed airflow, posing significant challenges to their stability and reliability. High-temperature environments can lead to performance degradation or even damage to the optical probes, severely affecting the accuracy and reliability of the measurement data. This places extremely high demands on the thermal protection of the probes.

[0004] To address this issue, researchers have proposed various cooling methods to protect optical probes. Among them, liquid cooling technology has been widely studied and applied due to its high thermal conductivity. PCN241333 provides a thermal protection structure for an aero-engine measurement probe. This structure incorporates water-cooling channels where circulating cooling water rapidly dissipates heat, maintaining the optical components within a suitable temperature environment. This ensures the engine measurement probe can operate stably in high-temperature environments and provides accurate and reliable measurement results. While this approach can achieve thermal protection for optical components to some extent, it has several limitations in practical applications.

[0005] (1) System complexity: Liquid cooling systems require complex piping and pumping systems to circulate coolant, which increases the complexity of the system and the difficulty of maintenance.

[0006] (2) Weight and space limitations: The piping and pumping equipment of the liquid cooling system increases the weight and volume of the overall system, and also limits the flexibility and applicability of the probe, which is a major challenge for space-constrained aero engines.

[0007] (3) Leakage risk: Seals and connectors in the liquid cooling system are prone to aging and damage in high-temperature environments, affecting measurement accuracy and probe reliability, and may increase the risk of coolant leakage.

[0008] (4) Environmental adaptability: Liquid cooling systems have strict requirements for the selection of coolant. Different working environments may require different coolants, which limits the adaptability of the system.

[0009] (5) Cost issues: The initial investment and operating costs of liquid cooling systems are high, especially when special coolants are required or frequent maintenance is needed. Summary of the Invention

[0010] In view of this, the present invention provides a thermal protection structure for an aero-engine measurement probe with a film-forming hole, so as to achieve both effective thermal protection for optical components and simplification of the measurement process, thereby improving the reliability and flexibility of the system.

[0011] The present invention provides the following technical solution: a thermal protection structure for an aero-engine measurement probe with film gas holes, comprising: a flange; an outer shell and an inner shell, both fixedly disposed on one side of the flange, and the outer shell and the inner shell are spaced apart to form a sandwich space for cold air input, and the outer shell is provided with a plurality of spaced film gas holes; an air duct, disposed on the other side of the flange, and the air duct is connected to the sandwich space.

[0012] Furthermore, an optical lens module is installed inside the inner shell, and an optical window is installed at the end of the outer and inner shells away from the flange.

[0013] Furthermore, the interlayer space includes a front cooling airflow channel, the air inlet of which is connected to the air intake pipe, and the air outlet of which is located below the optical window.

[0014] Furthermore, the interlayer space also includes a rear cooling air duct, the air inlet of which is connected to the ductwork, and the air outlet of which is located above the optical window.

[0015] Furthermore, the interlayer space also includes a first cooling channel, symmetrically arranged on both sides of the front cooling channel.

[0016] Furthermore, the interlayer space also includes a second cooling channel, symmetrically arranged on both sides of the rear cooling channel.

[0017] Furthermore, baffles are provided between the front cooling airflow channel and the first cooling airflow channel, between the second cooling airflow channels, and between the rear cooling airflow channel and the second cooling airflow channel.

[0018] Furthermore, the outer wall of the inner shell is provided with a wedge-shaped flow guiding structure.

[0019] Furthermore, there are multiple air intake pipes, evenly distributed along the circumference of the outer shell.

[0020] Furthermore, the outer shell and the inner shell form a shell assembly, which includes a large-diameter section, a transition section and a small-diameter section connected in sequence, with the large-diameter section connected to the flange.

[0021] Compared with the prior art, the beneficial effects that the at least one technical solution adopted by the present invention can achieve include at least the following:

[0022] 1. Excellent thermal protection: By introducing cool air from the compressor and setting multiple film vents on the outer casing, a protective film is formed, isolating the probe from the high-temperature gas combustion gas, thereby effectively reducing the probe temperature. Calculations show that, under the same high-temperature gas turbine flow field environment, the probe temperature using the thermal protection structure of this invention is reduced by more than 40% compared to the probe temperature without the thermal protection structure.

[0023] 2. Simplified Measurement Process: The thermal protection structure of this invention eliminates the need for external cooling liquid; air is drawn directly from the compressor, significantly simplifying the measurement process. This not only reduces system complexity and maintenance costs but also improves measurement reliability and accuracy.

[0024] 3. Improved probe reliability: Since a complex liquid cooling system is unnecessary, probe installation and disassembly are more convenient, and its application range is wider. Simultaneously, the formation of a protective gas film effectively reduces friction and wear between the probe surface and the high-temperature gas, extending the probe's service life.

[0025] 4. High Flexibility: The thermal protection structure of this invention can be adjusted and optimized according to actual needs. For example, the thickness and temperature of the protective gas film can be adjusted by changing the size and number of gas film holes; the flow rate of cold air in the cold air chamber can be changed by adjusting the opening of the flow regulating valve, etc. This makes the thermal protection structure of this invention more flexible and adaptable.

[0026] 5. Suitable for long-term monitoring: Due to the excellent stability and reliability of the thermal protection structure of this invention, long-term monitoring of aero engines can be achieved. This is of great significance for evaluating engine performance, predicting its lifespan, and timely detection and troubleshooting. Attached Figure Description

[0027] To more clearly illustrate the technical solutions of the embodiments of this application, the drawings used in the embodiments will be briefly introduced below. Obviously, the drawings described below are only some embodiments of this application. For those skilled in the art, other drawings can be obtained based on these drawings without creative effort.

[0028] Figure 1This is a three-dimensional structural schematic diagram of the present invention;

[0029] Figure 2 yes Figure 1 A longitudinal sectional view;

[0030] Figure 3 yes Figure 1 A partial sectional view;

[0031] Figure 4 This is a schematic diagram of the internal cooling flow channel of the present invention;

[0032] Figure 5 This is a schematic diagram of the conical expansion air film orifice of the present invention.

[0033] The attached figures are labeled as follows: 1. Outer shell; 2. Optical window; 3. Air film vent; 4. Flange; 5. Air intake pipe; 6. Inner shell; 7. Wedge-shaped flow guide structure; 8. Optical probe; 9. Partition; 10. Front cold air flow channel; 11. Rear cold air flow channel; 12. First cold air flow channel; 13. Second cold air flow channel. Detailed Implementation

[0034] The embodiments of this application will now be described in detail with reference to the accompanying drawings.

[0035] The following specific examples illustrate the implementation of this application. Those skilled in the art can easily understand other advantages and effects of this application from the content disclosed in this specification. Obviously, the described embodiments are only a part of the embodiments of this application, and not all of them. This application can also be implemented or applied through other different specific embodiments, and the details in this specification can also be modified or changed based on different viewpoints and applications without departing from the spirit of this application. It should be noted that, in the absence of conflict, the following embodiments and features in the embodiments can be combined with each other. Based on the embodiments in this application, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of this application.

[0036] like Figures 1 to 5 As shown, this embodiment of the invention provides a thermal protection structure for an aero-engine measurement probe with film vents, including a flange 4, an outer shell 1, an inner shell 6, and an air duct 5. The outer shell 1 and the inner shell 6 are both fixedly disposed on one side of the flange 4, and are spaced apart to form a sandwich space for cold air input. The outer shell 1 has multiple spaced film vents 3. The air duct 5 is disposed on the other side of the flange 4 and communicates with the sandwich space.

[0037] The technical principle of this invention is mainly based on film cooling. By introducing cool air from the compressor and setting multiple film cooling holes on the outer casing, the cool air flows within the cooling chamber and is ejected through these holes, forming a protective film. This protective film effectively isolates the probe from the high-temperature combustion gases, significantly reducing the probe's surface temperature. This allows the optical components to operate safely in high-temperature environments, extending their service life. Furthermore, since no coolant is required, the system's structural design is simplified, reducing production and maintenance costs. Simultaneously, the stability of film cooling contributes to providing more reliable measurement results, improving the monitoring accuracy of aero-engines under high-temperature conditions.

[0038] In this embodiment of the invention, the outer shell 1 serves as the first layer of protection against the impact of high-temperature gas combustion. Its surface is covered with numerous discrete gas film pores 3, which are arranged according to a specific pattern to ensure that the cold air can uniformly and stably cover the surface of the outer shell 1, forming an effective heat insulation barrier.

[0039] The inner shell 6 has a cavity inside, and inside the cavity is a stepped shaft-shaped optical lens module 8 for observation. The end of the inner shell 6 and the outer shell 1 away from the flange 4 is one end of the heat protection structure, and an optical window 2 is installed at one end of the heat protection structure.

[0040] The optical lens module 8, also known as the optical probe, is responsible for receiving the reflected light from the gas turbine blades and converting it into electrical signals for subsequent analysis and processing. The optical window 2 primarily functions to transmit light and protect the optical lens module 8 from external environmental influences.

[0041] It should be noted that the outer shell 1 and the inner shell 6 form a shell assembly, which includes a large-diameter section, a transition section, and a small-diameter section connected in sequence. The large-diameter section is connected to the flange 4. The inner shell 6 is a stepped shaft structure with the outer diameter decreasing from bottom to top. At the connection between the large-diameter section and the small-diameter section of the outer shell 1, the outer diameter of the inner shell 6 changes linearly to form a transition section. The outer shell 1 and the inner shell 6 are coaxially fitted together, and the sandwich structure between them is a cavity structure. The outer diameter of this cavity structure has the same trend as the outer diameter of the inner shell, and it is also divided into a large-diameter section, a transition section, and a small-diameter section from bottom to top above the flange 4.

[0042] The interlayer space in this embodiment of the invention includes a front cooling airflow channel 10, the air inlet of the front cooling airflow channel 10 is connected to the air intake pipe 5, and the air outlet of the front cooling airflow channel 10 is located below the optical window 2.

[0043] The interlayer space also includes a rear cooling air duct 11, the air inlet of which is connected to the air intake pipe 5, and the air outlet of the rear cooling air duct 11 is located above the optical window 2.

[0044] The front cooling airflow channel 10 provides cooling protection to the area in front of the optical window 2 near the large-diameter section (below the optical window 2 in the figure), ensuring stable transmission of the laser signal. The cylindrical structure of the front cooling airflow channel 10 provides an ideal channel for the introduction of optical fibers and the conduction of laser light. The rear cooling airflow channel 11 connects to the film cooling structure near the small-diameter section at the top, and its outlet is also located in front of the optical window 2 (above the optical window 2 in the figure), effectively supplementing the protective film and effectively purging high-temperature combustion gases and contaminants, protecting the optical window 2 from corrosion and contamination. The film cooling hole 3 drawing air from the rear cooling airflow channel 11 has an outlet angle of 90°. The film cooling hole 3 at this location has a relatively small impact on thermal protection, but it can, to some extent, fill the wake region behind the thermal protection structure, reducing the impact of the thermal protection structure on the flow field.

[0045] The interlayer space also includes a first cooling channel 12, symmetrically arranged on both sides of the front cooling channel 10. The interlayer space also includes a second cooling channel 13, symmetrically arranged on both sides of the rear cooling channel 11.

[0046] The first cooling channel 12 and the second cooling channel 13 not only maintain structural consistency in the large-diameter section, transition section, and small-diameter section, but also expand into a disc-shaped structure at the top of the small-diameter section to fit the overall structure. The first cooling channel 12, closer to the optical window, appropriately narrows in the small-diameter section, reserving ample space for the optical window. The second cooling channel 13, farther from the optical window, uses a specific aperture design with tapered expansion holes (see reference). Figure 5 As shown in the figure, it enhances the adhesion of cold air to the surface of the outer casing 1 and improves the thermal protection performance.

[0047] It should be noted that baffles 9 are provided between the front cooling airflow channel 10 and the first cooling airflow channel 12, between the second cooling airflow channels 13 and 13, and between the rear cooling airflow channel 11 and 13. The baffles 9 can maximize the efficiency of cooling air utilization and reduce energy loss in the cooling airflow.

[0048] Preferably, the outer wall of the inner shell 6 is provided with a wedge-shaped flow guiding structure 7. The wedge-shaped flow guiding structure 7 can further enhance the outflow effect of cold air and improve the thermal protection effect. The distribution of the wedge-shaped flow guiding structure 7 is related to the distribution of the film vents 3.

[0049] Multiple air intake pipes 5 are evenly distributed circumferentially along the outer shell 1. As a key component for cold air supply, the air intake pipes 5 are evenly distributed along the lower circumference of the outer shell 1 to ensure that cold air is evenly distributed to each cold air channel and cold air chamber. By connecting flow regulating valves and pressure monitoring sensors, operators can precisely control the flow rate and pressure of the cold air, thereby optimizing the coverage effect of the protective air film and ensuring that the thermal protection structure always operates at its best. The pressure monitoring sensors are used to monitor the pressure changes of the cold air entering the cold air chamber in real time to ensure that the thermal protection structure operates within the normal range.

[0050] It should be noted that the size and number of the film gas holes 3 can be adjusted according to actual needs to achieve the best thermal protection effect. Taking into account both space and cooling effect, except for the rear cold air flow channel 11 and the top of the thermal protection structure, the angle between the axis of the film gas holes 3 and the surface of the outer shell of the thermal protection structure is between 30° and 50°, that is, the outflow angle of the cold air is between 30° and 50°. After the cold air flows out, a film gas is formed on the surface of the outer shell 1 of the thermal protection structure, which isolates the thermal protection structure from the mainstream high-temperature combustion gas.

[0051] The present invention also provides a specific embodiment, as follows:

[0052] During operation, the cold air from the compressor enters the cooling channel through the intake pipe 5. As the cold air flows, it is ejected through multiple discrete film cooling holes 3 on the outer casing 1, forming a protective film. This protective film effectively isolates the thermal protection structure from the high-temperature combustion gas, effectively reducing the probe temperature. Simultaneously, by adjusting the opening of the flow regulating valve connected to the intake pipe 5, the flow rate of the cold air in the cooling channel can be changed, thereby optimizing the coverage effect of the protective film.

[0053] When the mainstream flow velocity v m =300m / s, temperature T m =400K, air conditioning temperature T c =300K, when the optical window faces the main incoming flow direction, by reasonably distributing the flow rate of each cooling air channel, the air film coverage effect can be optimized. The adiabatic air film cooling efficiency η is introduced to measure this effect, which is defined as follows:

[0054]

[0055] Among them, T aw T is the temperature of the adiabatic wall surface. c The inlet temperature of the cold air channel is denoted as . A higher value indicates a better film gas coverage effect. CFD simulation is used to optimize the position, angle, diameter, shape, and cold air volume distribution of the film gas holes.

[0056] The CFD simulation results were then applied as a temperature load to the optical probe 8 inside the inner shell 6 for fluid-structure interaction heat transfer analysis. The temperature of each region on the optical probe was generally below 80°C, indicating that the temperature of the probe using the thermal protection structure of the present invention was reduced by more than 40% compared with the probe without the thermal protection structure, thus achieving effective thermal protection for the optical components.

[0057] The above description is merely a specific embodiment of this application, but the scope of protection of this application is not limited thereto. Any variations or substitutions that can be easily conceived by those skilled in the art within the technical scope disclosed in this application should be included within the scope of protection of this application. Therefore, the scope of protection of this application should be determined by the scope of the claims.

Claims

1. An aeroengine measuring probe heat shield structure with gas film holes, characterized in that, The utility model relates to a kind of optical lens cooling device, including: Flange (4); Outer shell (1) and inner shell (6) are fixedly arranged on one side of flange (4), and outer shell (1) and inner shell (6) are spaced apart and form the sandwich space for cold gas input, and a plurality of spaced distribution gas film holes (3) are provided on outer shell (1); Air-bleed line (5) is arranged on the other side of flange (4), and air-bleed line (5) is communicated with the sandwich space; The sandwich space includes front cold gas flow channel (10), and the air inlet end of front cold gas flow channel (10) is connected with air-bleed line (5), and the air outlet end of front cold gas flow channel (10) is located below optical window (2);The sandwich space also includes rear cold gas flow channel (11), and the air inlet end of rear cold gas flow channel (11) is connected with air-bleed line (5), and the air outlet end of rear cold gas flow channel (11) is located above optical window (2);The sandwich space also includes rear cold gas flow channel (11), and the air inlet end of rear cold gas flow channel (11) is connected with air-bleed line (5), and the air outlet end of rear cold gas flow channel (11) is located above optical window (2);The sandwich space also includes first cooling flow channel (12), and is symmetrically arranged on the both sides of front cold gas flow channel (10);The sandwich space also includes second cooling flow channel (13), and is symmetrically arranged on the both sides of rear cold gas flow channel (11);The sandwich space also includes second cooling flow channel (13), and is symmetrically arranged on the both sides of rear cold gas flow channel (11);Front cold gas flow channel (10) and first cooling flow channel (12) are provided with baffle (9) between second cooling flow channel (13) and second cooling flow channel (13) between rear cold gas flow channel (11) and second cooling flow channel (13).

2. The aeroengine measurement probe heat shield structure with gas film orifice of claim 1, wherein, Optical lens module (8) is installed in inner shell (6), and optical window (2) is installed at the end of outer shell (1) and inner shell (6) away from flange (4).

3. The aeroengine measurement probe heat shield structure with gas film orifice of claim 1 or 2, wherein, The outer wall of inner shell (6) is provided with wedge-shaped flow guide structure (7).

4. The hot structure of a gas turbine engine measurement probe with gas film orifice according to claim 1 or 2, characterized in that, Air-bleed line (5) is multiple, and is uniformly distributed along the circumference of outer shell (1).

5. The hot gas path engine measurement probe thermal protection structure with gas film holes of claim 1 or 2, wherein, Outer shell (1) and inner shell (6) form a shell assembly, and the shell assembly includes large diameter section, transition section and small diameter section connected in sequence, and the large diameter section is connected with flange (4).

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