Thermal protection structure of aero-engine measuring probe with air film hole

Abstract

The invention provides an aero-engine measuring probe thermal protection structure with film holes. The aero-engine measuring probe thermal protection structure comprises a flange plate; the outer shell and the inner shell are both fixedly arranged on one side of the flange plate, the outer shell and the inner shell are arranged in a spaced mode and form an interlayer space used for cold air input, and a plurality of air film holes distributed at intervals are formed in the outer shell; and the air entraining pipeline is arranged on the other side of the flange plate, and the air entraining pipeline is communicated with the interlayer space. Cold air in the gas compressor is introduced, and the multiple gas film holes are formed in the shell, so that a layer of protective gas film is formed, the probe is isolated from high-temperature fuel gas, and the temperature of the probe is effectively reduced.

Description

Technical Field

[0001] The invention relates to the technical field of optical probe protection devices in high temperature environments, and in particular to a thermal protection structure of an aero-engine measuring probe with an air film hole. Background Art

[0002] In the fields of aviation, energy and industry, gas turbines are core power equipment, and their performance and reliability are crucial to the stable operation of the entire system. As one of the key components of aircraft engines, the blades of gas turbine turbines are subjected to great thermal and mechanical stress in a high-temperature, high-pressure and high-speed working environment, and are prone to deformation and wear, which in turn affects the overall performance and safety of the engine. Therefore, accurate measurement of the deformation of turbine blades is of great significance for evaluating their health, preventing failures and optimizing performance.

[0003] In order to monitor the deformation of gas turbine blades in real time, optical measurement technology is usually used in modern aircraft engine design. Optical measurement technology has the advantages of high precision, non-contact, and good real-time performance, and can realize dynamic monitoring of blade deformation. However, when the optical measurement probe works in the high-temperature gas turbine flow field, it faces the impact of extremely high temperature and high-speed airflow, and its stability and reliability face huge challenges. The high temperature environment will cause the performance of the optical probe to degrade or even be damaged, seriously affecting the accuracy and reliability of the measurement data, which puts extremely high demands on the thermal protection of the probe.

[0004] In order to solve this problem, researchers have proposed a variety of cooling methods to protect optical probes. Among them, liquid cooling technology has been widely studied and applied due to its efficient heat conduction performance. PCN241333 provides a thermal protection structure for aircraft engine measurement probes. The structure is equipped with a water cooling channel. The cooling water circulating in the channel quickly removes the heat and maintains the optical components 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. Although this solution can achieve thermal protection for optical components to a certain extent, it has many limitations in practical applications:

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

[0006] (2) Weight and space limitations: The pipes and pumping equipment of the liquid cooling system increase the weight and volume of the overall system and also limit the flexibility and applicability of the probe, which is a major challenge for space-constrained aircraft 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: The liquid cooling system has strict requirements on the selection of coolant. Different working environments may require different coolants, which limits the adaptability of the system.

[0009] (5) Cost issue: The initial investment and operating costs of the liquid cooling system are high, especially when special coolants or frequent maintenance are required. Summary of the invention

[0010] In view of this, the present invention provides an aircraft engine measurement probe thermal protection structure with an air film hole, so as to achieve the purpose of achieving effective thermal protection for optical components, simplifying the measurement process, and improving the reliability and flexibility of the system.

[0011] The present invention provides the following technical solution: a thermal protection structure of an aircraft engine measuring probe with an air film hole, comprising: a flange; an outer shell and an inner shell, both of which are fixedly arranged on one side of the flange, and the outer shell and the inner shell are spaced apart to form an interlayer space for cold air input, and a plurality of air film holes distributed at intervals are arranged on the outer shell; an air bleed pipe, which is arranged on the other side of the flange, and the air bleed pipe is connected to the interlayer space.

[0012] Furthermore, an optical lens module is installed inside the inner shell, and optical windows are installed at ends of the outer shell and the inner shell away from the flange.

[0013] Furthermore, the interlayer space includes a front cooling air flow channel, an air inlet end of the front cooling air flow channel is connected to the air duct, and an air outlet end of the front cooling air flow channel is located below the optical window.

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

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

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

[0017] Furthermore, partitions are provided between the front cooling flow channel and the first cooling flow channel, between the second cooling flow channel and the second cooling flow channel, and between the rear cooling flow channel and the second cooling flow 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 ducts, which are evenly spaced along the circumference of the shell.

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

[0021] Compared with the prior art, the beneficial effects that can be achieved by at least one of the above technical solutions adopted by the present invention include at least:

[0022] 1. Good thermal protection effect: By introducing cold air from the compressor and setting multiple air film holes on the shell, a protective air film is formed to isolate the probe from the high-temperature gas, thereby effectively reducing the temperature of the probe. Calculations show that under the same high-temperature gas turbine turbine flow field environment, the temperature of the probe using the thermal protection structure of the present invention is reduced by more than 40% compared with the probe without the thermal protection structure.

[0023] 2. Simplify the measurement process: The thermal protection structure of the present invention does not need to introduce cooling liquid from the outside, but only needs to draw air from the compressor, which greatly simplifies the measurement process. This not only reduces the complexity and maintenance cost of the system, but also improves the reliability and accuracy of the measurement.

[0024] 3. Improve the reliability of the probe: Since there is no need for a complex liquid cooling system, the installation and removal of the probe is more convenient and has a wider range of applications. At the same time, the formation of a protective gas film can also effectively reduce the friction and wear between the probe surface and the high-temperature gas, extending the service life of the probe.

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

[0026] 5. Suitable for long-term monitoring: Since the thermal protection structure of the present invention has good stability and reliability, it can realize long-term monitoring of aircraft engines. This is of great significance for evaluating the performance of the engine, predicting its life, and timely discovering and eliminating faults. BRIEF DESCRIPTION OF THE DRAWINGS

[0027] 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.

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

[0029] Figure 2 yes Figure 1 A longitudinal cross-sectional view of

[0030] Figure 3 yes Figure 1 A partial cross-sectional view of

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

[0032] Figure 5 Schematic diagram of the conical expansion air film hole of the present invention.

[0033] Reference numerals in the figure: 1. outer shell; 2. optical window; 3. air film hole; 4. flange; 5. air duct; 6. inner shell; 7. wedge-shaped guide structure; 8. optical probe; 9. partition; 10. front cold air flow duct; 11. rear cold air flow duct; 12. first cold air flow duct; 13. second cold air flow duct. DETAILED DESCRIPTION

[0034] The embodiments of the present application are described in detail below with reference to the accompanying drawings.

[0035] 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.

[0036] like Figures 1 to 5 As shown, an embodiment of the present invention provides a thermal protection structure of an aircraft engine measuring probe with air film holes, comprising a flange 4, an outer shell 1, an inner shell 6 and an air bleed pipe 5. The outer shell 1 and the inner shell 6 are both fixedly arranged on one side of the flange 4, and the outer shell 1 and the inner shell 6 are arranged at intervals to form an interlayer space for cold air input, and a plurality of air film holes 3 distributed at intervals are arranged on the outer shell 1; the air bleed pipe 5 is arranged on the other side of the flange 4, and the air bleed pipe 5 is connected to the interlayer space.

[0037] The technical principle of the present invention is mainly based on the principle of air film cooling. By introducing cold air from the compressor and setting a plurality of air film holes on the outer shell, the cold air is ejected through the air film holes when flowing in the cold air cavity to form a protective air film. This protective air film can effectively isolate the probe from the high-temperature combustion gas, significantly reduce the temperature of the probe surface, enable the optical components to operate safely in a high-temperature environment, and extend its service life. In addition, since there is no need to introduce coolant, the structural design of the system is simplified, and the production and maintenance costs are reduced. At the same time, the stability of air film cooling helps to provide more reliable measurement results and improve the monitoring accuracy of aircraft engines under high-temperature conditions.

[0038] In the embodiment of the present invention, the outer shell 1 serves as the first layer of protection against the impact of high-temperature combustion gas, and a large number of discrete air film holes 3 are distributed on its surface. These air film holes 3 are distributed according to a specific arrangement pattern to ensure that the cold air can evenly and stably cover the surface of the outer shell 1 to form an effective heat insulation barrier.

[0039] The inner shell 6 has a cavity inside, and the cavity contains 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 thermal protection structure, and an optical window 2 is installed at one end of the thermal protection structure.

[0040] The optical lens module 8 is an optical probe, which is responsible for receiving the reflected light from the turbine blades of the gas turbine and converting it into an electrical signal for subsequent analysis and processing. The main function of the optical window 2 is to transmit light and protect the optical lens module 8 from the influence of the external environment.

[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, and the large diameter section is connected to the flange 4. The inner shell 6 is a stepped shaft structure with a decreasing outer diameter 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 sleeved, and the sandwich structure between the two is a cavity structure. The outer diameter of the cavity structure has the same change trend as the outer diameter of the inner shell, and 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 the embodiment of the present invention comprises a front cooling air flow channel 10 , an air inlet end of the front cooling air flow channel 10 is connected to the air duct 5 , and an air outlet end of the front cooling air flow channel 10 is located below the optical window 2 .

[0043] The interlayer space further includes a rear cooling air flow channel 11 , an air inlet end of the rear cooling air flow channel 11 is connected to the air duct 5 , and an air outlet end of the rear cooling air flow channel 11 is located above the optical window 2 .

[0044] The front cold air flow channel 10 is responsible for providing cold air protection to the front of the optical window 2 near the large diameter section (below the optical window 2 in the figure) to ensure the stable transmission of the laser signal. The cylindrical structure design of the front cold air flow channel 10 provides an ideal channel for the introduction of optical fiber and the conduction of laser. The rear cold air flow channel 11 is connected to the air film cooling structure near the small diameter section at the top of the small diameter section, and its outlet is also located in front of the optical window 2 (above the optical window 2 in the figure), which effectively supplements the protective air film and can effectively blow away high-temperature combustion gas and pollutants to protect the optical window 2 from erosion and pollution. The air film hole 3 that draws air from the rear cold air flow channel 11 has an outlet angle of 90°. The air film hole 3 at this location has little effect on thermal protection, but it can enrich the wake area behind the thermal protection structure to a certain extent and reduce the impact of the thermal protection structure on the flow field.

[0045] The interlayer space further includes a first cooling channel 12 symmetrically arranged on both sides of the front cooling channel 10 . The interlayer space further 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 pancake-shaped structure at the top of the small diameter section to adapt to the overall structure. The first cooling channel 12 near the optical window is appropriately contracted in the small diameter section to reserve sufficient space for the optical window. The second cooling channel 13 away from the optical window is designed with a specific hole shape of a conical expansion hole (refer to Figure 5 As shown), the adhesion effect of cold air to the surface of the housing 1 is enhanced, and the thermal protection performance is improved.

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

[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 heat protection effect. The distribution of the wedge-shaped flow guiding structure 7 is related to the distribution of the air film holes 3.

[0049] There are multiple air ducts 5, which are evenly spaced along the circumference of the outer shell 1. As a key component for cold air supply, the air duct 5 is evenly distributed along the circumference of the lower part of the outer shell 1 to ensure that the cold air can be evenly distributed to each cold air flow channel and cold air cavity. By connecting the flow regulating valve and the pressure monitoring sensor, the operator can accurately control the flow and pressure of the cold air, thereby optimizing the coverage effect of the protective air film and ensuring that the thermal protection structure always works in the best state. The pressure monitoring sensor is used to monitor the pressure changes of the cold air entering the cold air cavity in real time to ensure that the thermal protection structure works within the normal range.

[0050] It should be noted that the size and number of the air film holes 3 can be adjusted according to actual needs to achieve the best thermal protection effect. Taking into account the 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 air film hole 3 and the surface of the thermal protection structure shell 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 layer of air film is formed on the surface of the thermal protection structure shell 1, isolating the thermal protection structure from the mainstream high-temperature gas.

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

[0052] During operation, the cold air in the compressor enters the cooling flow channel through the air bleed pipe 5. The cold air is ejected through the multiple discrete air film holes 3 on the shell 1 during the flow process to form a protective air film. This protective air film can effectively isolate the thermal protection structure from the high-temperature fuel gas and effectively reduce the temperature of the probe. At the same time, by adjusting the opening of the flow control valve connected to the air bleed pipe 5, the cold air flow in the cold air flow channel can be changed, thereby optimizing the coverage effect of the protective air film.

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

[0054]

[0055] Among them, T aw is the adiabatic wall temperature, T c is the inlet temperature of the cold air flow channel. The larger the value, the better the air film coverage effect. CFD simulation is used to optimize the air film hole position, angle, aperture, hole shape and cold air distribution.

[0056] The CFD simulation results are then applied as temperature loads to the optical probe 8 inside the inner shell 6 for fluid-solid coupling heat transfer analysis. The temperatures of various areas on the optical probe are generally below 80°C, indicating that the temperature of the probe using the thermal protection structure of the present invention is reduced by more than 40% compared with the probe not using the thermal protection structure, thereby achieving effective thermal protection for the optical components.

[0057] 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 measuring probe with an air film hole, characterized in that: include: Flange (4); The outer shell (1) and the inner shell (6) are both fixedly arranged on one side of the flange (4), and the outer shell (1) and the inner shell (6) are arranged at intervals to form an interlayer space for cold air input, and the outer shell (1) is provided with a plurality of air film holes (3) distributed at intervals; An air bleed pipe (5) is arranged on the other side of the flange (4), and the air bleed pipe (5) is communicated with the interlayer space.

2. The thermal protection structure of the aircraft engine measuring probe with air film holes according to claim 1, characterized in that: An optical lens module (8) is installed inside the inner shell (6), and an optical window (2) is installed on the outer shell (1) and the end of the inner shell (6) away from the flange (4).

3. The thermal protection structure of the aircraft engine measuring probe with air film holes according to claim 2 is characterized in that: The interlayer space comprises a front cold air flow channel (10), the air inlet end of the front cold air flow channel (10) is connected to the air duct (5), and the air outlet end of the front cold air flow channel (10) is located below the optical window (2).

4. The thermal protection structure of the aircraft engine measuring probe with air film holes according to claim 3 is characterized in that: The interlayer space also includes a rear cooling air flow channel (11), the air inlet end of the rear cooling air flow channel (11) is connected to the air duct (5), and the air outlet end of the rear cooling air flow channel (11) is located above the optical window (2).

5. The thermal protection structure of the aircraft engine measuring probe with air film holes according to claim 4, characterized in that: The interlayer space also includes a first cooling flow channel (12) symmetrically arranged on both sides of the front cooling flow channel (10).

6. The thermal protection structure of the aircraft engine measuring probe with air film holes according to claim 5, characterized in that: The interlayer space also includes a second cooling flow channel (13) which is symmetrically arranged on both sides of the rear cooling flow channel (11).

7. The thermal protection structure of the aircraft engine measuring probe with air film holes according to claim 6, characterized in that: Partition plates (9) are provided between the front cooling flow channel (10) and the first cooling flow channel (12), between the second cooling flow channel (13) and the second cooling flow channel (13), and between the rear cooling flow channel (11) and the second cooling flow channel (13).

8. The thermal protection structure of an aircraft engine measuring probe with air film holes according to any one of claims 1 to 7, characterized in that: The outer wall of the inner shell (6) is provided with a wedge-shaped flow guide structure (7).

9. The thermal protection structure of an aircraft engine measuring probe with air film holes according to any one of claims 1 to 7, characterized in that: There are a plurality of air bleed pipelines (5) which are evenly spaced along the circumference of the outer shell (1).

10. The thermal protection structure of an aircraft engine measuring probe with air film holes according to any one of claims 1 to 7, characterized in that: The outer shell (1) and the inner shell (6) form a shell assembly, which comprises a large diameter section, a transition section and a small diameter section which are connected in sequence, and the large diameter section is connected to the flange (4).

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