Gas turbine trailing edge heat exchange enhanced cooling structure design method based on choked flow characteristic analysis

Abstract

The invention discloses a gas turbine trailing edge heat exchange enhanced cooling structure design method based on choked flow characteristic analysis, relates to a gas turbine trailing edge cooling structure, aims to solve the problems of high manufacturing difficulty, high processing cost and inconvenience in engineering application in existing blade cooling, and provides a cooling structure based on choked flow characteristic analysis. Comprising a cooling air compressor used for providing high-pressure cooling air; a cooling channel is arranged in the blade, and at least two rows of turbulent flow columns are arranged in the cooling channel to increase the turbulence intensity of cooling air; the crack tail cavity is formed in the tail edge of the blade and used for forming a stable air film layer and reducing thermal erosion. According to the design method, the cooling channel structure is optimized through numerical simulation, the tortuosity degree of the serpentine channel is determined through an optimization algorithm, and the width and the outlet position of the slit tail cavity are optimized so that it can be guaranteed that cooling air is evenly distributed. The cooling structure is suitable for cooling work of the gas turbine blade, the cooling efficiency is remarkably improved, and the machining difficulty and cost are reduced.

Description

Technical Field

[0001] The invention relates to the field of heat dissipation structures at the trailing edge of a gas turbine, and in particular to an enhanced cooling structure for heat exchange at the trailing edge of a gas turbine and a design method thereof. Summary of the invention

[0002] During the design and manufacturing process of gas turbines, the cooling efficiency of the blades directly affects the overall performance and life of the gas turbine. Gas turbine blades, especially the trailing edge, are subjected to extremely high temperatures and harsh environments, so effective cooling technology is essential. Existing gas turbine cooling technologies can be mainly divided into two categories: film cooling and internal cooling.

[0003] In terms of film cooling, common technologies include blade film cooling, end wall film cooling and trailing edge film cooling. This cooling method forms a protective film layer on the blade surface to reduce the direct thermal erosion of the high-temperature combustion gas on the blade substrate. However, the effect of film cooling is limited by the arrangement of the cooling airflow, especially in high-temperature areas, where the film is easily torn and cannot maintain a stable cooling layer for a long time, affecting the cooling efficiency.

[0004] Internal cooling technology absorbs heat and removes it by flowing cooling air inside the blade. Common methods include jet cooling and fin cooling. Jet cooling has a good cooling effect by directly impacting the high-temperature surface with a jet of cold air. However, in actual engineering, its complex flow channel design makes the manufacturing cost high and the processing difficult. Fin cooling enhances heat transfer by arranging fins in the cooling channel. The design of the fins can effectively increase the heat exchange area and disturb the flow, thereby improving the cooling efficiency. However, its heat exchange capacity is still limited by the channel layout and the geometry of the fins.

[0005] In recent years, in order to further improve the cooling effect of gas turbine blades, academia and industry have invested a lot of research, such as optimizing blade cooling structure through deep learning, or designing complex cooling channels by topology optimization and structural model modification. Although these methods have achieved certain results in improving heat transfer performance, they often face the following major problems:

[0006] 1. The structural design is too complex: Many high-efficiency cooling structures use complex geometric shapes and topological layouts, which makes the manufacturing and processing process very difficult, and the processing cost remains high, which is not conducive to practical engineering applications.

[0007] 2. Limited research on heat transfer mechanisms: Currently, most studies focus on improving heat transfer performance by improving existing structural models, while research on new composite cooling structures, especially on heat transfer enhancement based on the flow resistance characteristics of spoiler columns, is relatively scarce, and the advantages of both internal cooling and film cooling mechanisms have not been effectively combined.

[0008] 3. Lack of easy-to-implement solutions: Although many existing research results have good heat transfer performance in theory, due to the complexity of structural design and difficulty in processing, these solutions are difficult to implement in practical applications and cannot meet the needs of low cost and easy processing. Summary of the invention

[0010] In order to solve the technical defects existing in the prior art, the cooling work of the existing gas turbine blades has the following problems: the manufacturing and processing process is very difficult, the processing cost is high, which is not conducive to practical engineering applications, and the research on heat exchange enhancement based on the flow resistance characteristics of the spoiler column is relatively rare, and the advantages of the two mechanisms of internal cooling and film cooling cannot be effectively combined. The technical solution provided by the present invention is:

[0011] A heat exchange enhanced cooling structure for a gas turbine trailing edge based on resistance characteristic analysis, comprising:

[0012] A cooling air compressor for providing high-pressure cooling air;

[0013] A blade, wherein a cooling passage is provided between the blade and the compressor;

[0014] At least two rows of spoiler columns are arranged in the cooling channel to increase the turbulence intensity;

[0015] The split tail cavity is arranged in the trailing edge area of ​​the blade and is used to guide cooling air to form an air film layer.

[0016] Furthermore, a preferred embodiment is provided, wherein the cooling channel is a serpentine cooling channel.

[0017] Furthermore, a preferred embodiment is provided, wherein the cross-section of the spoiler column is cylindrical or elliptical.

[0018] Furthermore, a preferred embodiment is provided, wherein the spoiler columns are arranged at periodic intervals.

[0019] Furthermore, a preferred embodiment is provided, wherein the inner wall of the cooling channel is a rough surface.

[0020] Furthermore, a preferred embodiment is provided, wherein the slit tail cavity has at least two outlets.

[0021] Based on the same inventive concept, the present invention also provides a method for designing a heat exchange enhanced cooling structure at the trailing edge of a gas turbine based on resistance characteristic analysis, which is used to design the structure, including:

[0022] Steps of optimizing the structure of the cooling channel by numerical simulation;

[0023] The step of determining the tortuosity of the serpentine cooling channel by using an optimization algorithm based on the flow path of the cooling air;

[0024] The step of optimizing the structural parameters of the slit tail cavity includes the slit width and the outlet position.

[0025] Based on the same inventive concept, the present invention also provides a computer storage medium for storing a computer program, and when the computer reads the computer program, the computer executes the described method.

[0026] Based on the same inventive concept, the present invention also provides a computer, including a processor and a storage medium. When the processor reads a computer program stored in the storage medium, the computer executes the method described.

[0027] Based on the same inventive concept, the present invention also provides a computer program product, which is a computer program. When the computer program is executed, the method described above is implemented.

[0028] Compared with the prior art, the technical solution provided by the present invention is beneficial in that:

[0029] By arranging spoiler columns inside the blade, the disturbance effect of cooling air on the inner wall of the blade is significantly enhanced, and the heat exchange efficiency is improved. Compared with traditional jet cooling, the arrangement of spoiler columns not only increases the contact area between the cooling air and the inner wall, but also makes the cooling effect more significant by destroying the development of the boundary layer, thereby greatly improving the overall heat exchange capacity.

[0030] Combined with the design of the serpentine cooling channel, the flow path of the cooling air inside the blade is effectively extended, so that the cooling air can fully absorb the heat of the blade. Compared with the existing straight cooling channel, the serpentine channel design increases the turbulence intensity, further improves the heat exchange efficiency, and maintains a low pressure loss.

[0031] The application of the split tail cavity further optimizes the cooling effect of the trailing edge area. The split tail cavity guides the cooling air to form an efficient air film cooling layer in the trailing edge area. Compared with the existing air film cooling method, the split tail cavity not only improves the stability of the cooling layer, but also effectively reduces the direct erosion of the high-temperature airflow on the blade substrate, thereby extending the service life of the blade.

[0032] In general, the present invention, through the combination of the above three parts, not only solves the problems of complex cooling structure and high manufacturing difficulty in the prior art, but also finds a good balance between cooling efficiency and structural simplification, thereby significantly improving the overall performance and reliability of the gas turbine.

[0033] Suitable for cooling gas turbine blades. BRIEF DESCRIPTION OF THE DRAWINGS

[0034] Figure 1 Schematic diagram of the heat exchange enhanced cooling structure at the trailing edge of a gas turbine based on the analysis of resistance characteristics.

[0035] Among them, 1-blade front cavity, 2-membrane hole, 3-blade body, 4-flow basin spoiler column, 5-slit tail cavity, 6-wedge-shaped internal cooling channel, 7-thick lip wall, 8-rib plate, 9-symmetrical side spoiler column, 10-periodic symmetrical wall. DETAILED DESCRIPTION

[0036] In order to make the advantages and benefits of the technical solution provided by the present invention more clearly reflected, the technical solution provided by the present invention is now further described in detail in conjunction with the accompanying drawings, specifically:

[0037] Embodiment 1: This embodiment provides a gas turbine trailing edge heat exchange enhanced cooling structure based on resistance characteristic analysis, including:

[0038] A cooling air compressor for providing high-pressure cooling air;

[0039] A blade, wherein a cooling passage is provided between the blade and the compressor;

[0040] At least two rows of spoiler columns are arranged in the cooling channel to increase the turbulence intensity;

[0041] The split tail cavity is arranged in the trailing edge area of ​​the blade and is used to guide cooling air to form an air film layer.

[0042] Specifically, the design method is:

[0043] Step 1: Cooling air introduction and establishment of flow path inside the blade

[0044] The cooling air is first ejected by the high-pressure compressor and enters the cooling channel inside the blade through the bottom of the blade. Inside the blade, the cooling air flows along the pre-arranged serpentine cooling channel. The serpentine cooling channel effectively prolongs the residence time of the cooling air inside the blade through the tortuous flow path, ensuring that the air can absorb more heat. This design lays the foundation for the subsequent turbulent cooling of the cooling air, and also further improves the heat exchange efficiency through the turbulent effect in the initial stage.

[0045] Step 2: Arrangement of spoiler columns and enhanced heat exchange of cooling air

[0046] Multiple rows of spoiler columns are arranged in the serpentine cooling channel. When the cooling air flows in the serpentine channel, it will be disturbed by the spoiler columns, forming strong turbulence. The design and layout of the spoiler columns are precisely calculated to ensure that the cooling air can disturb the wall boundary layer to the greatest extent when passing through, thereby improving the efficiency of heat transfer. The geometry, diameter and layout angle of the spoiler columns are optimized so that they can effectively increase the heat exchange area without causing excessive pressure loss.

[0047] Step 3: Cooling design of the split tail cavity

[0048] After the cooling air passes through the serpentine channel and the spoiler column area, it enters the split tail cavity area. The split tail cavity is precisely designed to form a stable air film layer when the cooling air passes through the trailing edge area. This air film layer can effectively reduce the direct thermal erosion of the high-temperature combustion gas on the blade substrate and further improve the cooling effect of the trailing edge area. Compared with the traditional air film cooling method, the design of the split tail cavity enhances the stability of the cooling layer and allows the cooling air to be evenly distributed in the trailing edge area of ​​the blade, thereby achieving a better cooling effect.

[0049] Step 4: Realization of the overall effect of the system

[0050] With the combination of the above cooling methods, the cooling air forms a continuous cooling path inside the blade from the serpentine channel to the spoiler column and then to the split tail cavity. The cooling effect of each step provides a good foundation for the next step. Through this coherent cooling process, the overall cooling effect of the blade is greatly improved. The strong turbulence effect of the spoiler column combined with the air film cooling of the split tail cavity significantly improves the heat exchange efficiency of the blade and reduces the heat load of the high temperature on the blade material, ultimately achieving the effect of extending the service life of the blade and improving the overall reliability of the gas turbine.

[0051] Embodiment 2: This embodiment further limits the heat exchange enhanced cooling structure of the gas turbine trailing edge based on the resistance characteristic analysis provided in Embodiment 1, and the cooling channel is a serpentine cooling channel.

[0052] Embodiment 3: This embodiment further limits the heat exchange enhanced cooling structure for the trailing edge of a gas turbine based on the analysis of the flow resistance characteristics provided in Embodiment 1, and the cross-section of the spoiler column is cylindrical or elliptical.

[0053] Embodiment 4: This embodiment further limits the heat exchange enhanced cooling structure for the trailing edge of a gas turbine based on the analysis of flow resistance characteristics provided in Embodiment 1, and the spoiler columns are arranged at periodic intervals.

[0054] Embodiment 5: This embodiment further limits the heat exchange enhanced cooling structure for the trailing edge of a gas turbine based on the analysis of flow resistance characteristics provided in Embodiment 1, and the inner wall of the cooling channel is a rough surface.

[0055] Embodiment 6: This embodiment further limits the heat exchange enhanced cooling structure of the gas turbine trailing edge based on the analysis of the flow resistance characteristics provided in Embodiment 1, and the split tail cavity has at least two outlets.

[0056] Embodiment 7: This embodiment provides a method for designing a heat exchange enhanced cooling structure of a gas turbine trailing edge based on resistance characteristic analysis, which is used to design the structure provided in embodiment 1, including:

[0057] Steps of optimizing the structure of the cooling channel by numerical simulation;

[0058] The step of determining the tortuosity of the serpentine cooling channel by using an optimization algorithm based on the flow path of the cooling air;

[0059] The step of optimizing the structural parameters of the slit tail cavity includes the slit width and the outlet position.

[0060] Specifically, they include:

[0061] 1. Structural design of cooling device

[0062] 1.1 Cooling air compressor

[0063] The cooling device includes a cooling air compressor, which is used to provide high-pressure cooling air. The cooling air is ejected through the compressor and enters the bottom of the blade, and then enters the cooling channel inside the blade. The flow rate of the cooling air can be adjusted according to the actual operating conditions to optimize the cooling effect and reduce pressure loss.

[0064] 1.2 Cooling channel inside the blade

[0065] The cooling channel is designed to guide the cooling air through the high-temperature area of ​​the blade. The cooling channel design includes multiple spoiler columns and serpentine cooling channels to maximize the contact area between the cooling air and the blade surface, improving the overall heat exchange efficiency.

[0066] 1.3 Design of spoiler column

[0067] Multiple spoiler columns are arranged in the cooling channel inside the blade. The spoiler columns can be cylindrical or elliptical, and are arranged at periodic intervals to enhance the turbulence effect of the cooling air and evenly distribute the heat. The spoiler columns disturb the flow of cooling air, destroy the stability of the boundary layer, increase the turbulence intensity, and improve the heat transfer efficiency. At the same time, the spoiler columns are made of metal materials with good thermal conductivity to further improve the overall heat dissipation effect of the blade.

[0068] 1.4 Serpentine cooling channel

[0069] The serpentine cooling channel is located inside the blade. The cooling air flows along the tortuous channel, which prolongs the residence time of the air inside the blade, thus ensuring that the cooling air can fully absorb heat. The tortuosity of the serpentine channel is optimized to ensure that the cooling efficiency is improved without significantly increasing the pressure loss. The inner wall of the channel is provided with a rough surface to further increase the turbulence intensity of the cooling air, thereby improving the heat exchange effect.

[0070] 1.5 Design of the split tail cavity

[0071] A split tail cavity is provided in the trailing edge area of ​​the blade, and the cooling air enters the split tail cavity after flowing through the spoiler column and the serpentine cooling channel. The split tail cavity is designed to guide the cooling air to form a stable air film layer at the trailing edge to reduce the direct thermal erosion of the high-temperature airflow on the blade substrate. The width of the split tail cavity is precisely designed to ensure that the cooling air is evenly distributed in the trailing edge area to form a stable air film layer. The split tail cavity has multiple outlets, which can reduce the accumulation of cooling air at the trailing edge and improve the discharge efficiency of the cooling air.

[0072] 2. Algorithmic Implementation of Design Method

[0073] The design method of this embodiment optimizes the structure of the cooling channel inside the blade through numerical simulation. The specific implementation steps are as follows:

[0074] 2.1 Geometry optimization of spoiler column

[0075] The geometry, position and arrangement of the spoiler pins are determined through numerical simulation to maximize turbulence intensity and improve heat transfer efficiency. The diameter, length and arrangement spacing of the spoiler pins are optimized through simulation to ensure the strongest turbulence effect on the cooling air without significantly increasing pressure loss.

[0076] 2.2 Optimization design of serpentine cooling channel

[0077] Based on the flow path of the cooling air, an optimization algorithm is used to determine the degree of tortuosity of the serpentine cooling channel to improve the heat exchange efficiency of the cooling air without significantly increasing the pressure loss. The tortuosity design of the serpentine cooling channel increases the residence time of the cooling air, thereby ensuring the maximum cooling effect.

[0078] 2.3 Structural optimization of the split tail cavity

[0079] According to the simulation results, the structural parameters of the split tail cavity are optimized, including the width and outlet position of the split. The optimized split tail cavity can effectively guide the cooling air to be evenly distributed in the trailing edge area and form a stable air film layer to minimize the erosion of the high-temperature airflow on the blade substrate.

[0080] 2.4 Iterative calculation and optimization selection

[0081] Through iterative calculation, the influence of each cooling channel and spoiler column layout scheme on the overall heat transfer efficiency and pressure loss is evaluated. The optimal scheme is selected to maximize the cooling effect and ensure the processing feasibility and economy of the cooling structure design.

[0082] 3. Working principle and advantages

[0083] In this embodiment, high-pressure cooling air is provided by a cooling air compressor, and the cooling air passes through a continuous cooling path of the spoiler column, the serpentine cooling channel and the split tail cavity, so that the overall cooling effect of the blade is greatly improved. The combination of the strong turbulence effect of the spoiler column, the extended cooling path of the serpentine cooling channel and the air film cooling layer formed by the split tail cavity effectively improves the heat exchange efficiency of the blade, reduces the heat load of the high temperature on the blade material, and thus extends the service life of the blade.

[0084] Compared with the prior art, the cooling structure design of this embodiment has the following advantages:

[0085] 1. High cooling efficiency: The combined design of spoiler columns, serpentine cooling channels and split tail cavities significantly improves the cooling efficiency, ensuring the stable operation of the blades in high temperature environments.

[0086] 2. Simple structure and easy processing: Compared with the complex topology optimization structure, the cooling channel and spoiler column of this embodiment are relatively simple in design, easy to process and manufacture, and reduce the manufacturing cost.

[0087] 3. Strong adjustability: The output flow rate of the cooling air compressor is adjustable to adapt to different operating conditions and further optimize the cooling effect.

[0088] Embodiment 8: This embodiment provides a computer storage medium for storing a computer program. When the computer reads the computer program, the computer executes the method provided in embodiment 7.

[0089] Embodiment 9: This embodiment provides a computer, including a processor and a storage medium. When the processor reads a computer program stored in the storage medium, the computer executes the method provided in embodiment 7.

[0090] Embodiment 10: This embodiment provides a computer program product, which is a computer program. When the computer program is executed, the method provided in embodiment 7 is implemented.

[0091] Implementation Method XI: Combination Figure 1This embodiment is described in detail by using specific examples to further describe the technical solution provided above. Specifically:

[0092] Blade front cavity 1: This is the internal cavity at the front of the blade, one of the paths through which cooling air enters. The front cavity is used to reduce the temperature of the leading edge portion of the blade to avoid damage to the blade material due to high temperature.

[0093] Air film holes 2: These holes are located on the blade surface, and the cooling air discharged through these holes forms a protective air film layer on the blade surface. The purpose of air film cooling is to form an insulating layer on the blade surface to reduce the direct thermal shock of the gas on the blade.

[0094] Blade body 3: The main part of the blade is the main load-bearing structure of the entire cooling system. It contains multiple cooling channels and cavities for effectively dispersing the cooling air flow.

[0095] Incoming flow spoiler 4: These spoiler columns are located in the incoming flow area of ​​the blades. Their purpose is to increase the turbulence intensity of the cooling air to improve the heat transfer efficiency. They enhance the overall heat transfer effect by destroying the boundary layer of the cooling air.

[0096] Split tail cavity 5: This is a split structure located at the trailing edge of the blade, used to guide the cooling air to form an air film layer at the trailing edge to reduce the direct thermal erosion of the combustion gas on the trailing edge area of ​​the blade, protect the blade structure, and improve the efficiency and stability of the trailing edge cooling.

[0097] Wedge-shaped internal cooling channel 6: This is a special cooling channel designed in a wedge-shaped structure to guide the cooling air to flow along a specific path inside the blade. The wedge-shaped structure helps control the flow rate and flow direction of the cooling air, further improving the heat exchange efficiency.

[0098] Thick lip wall 7: This is a thicker structural portion in the trailing edge area of ​​the blade, which is used to enhance the mechanical strength of the blade and withstand the high pressure stress caused by the high-speed airflow, while also increasing the stability of the cooling system.

[0099] Rib 8: Located inside the blade, the rib is used to increase the heat exchange area of ​​the cooling channel, promote the flow of cooling air inside the blade and improve the heat exchange effect. The design of the rib ensures that the cooling air can absorb the heat inside the blade to the maximum extent.

[0100] Symmetrical side spoiler columns 9: These spoiler columns are located on the symmetrical sides of the blades to further disturb the cooling air and distribute it more evenly inside the blades, thereby improving the uniformity and effect of cooling. The symmetrical design ensures the symmetry and consistency of the overall cooling effect.

[0101] Periodic symmetric wall 10: This is a periodic symmetric wall design, which is usually used to improve the air flow characteristics in the cooling channel so as to increase turbulence without increasing excessive pressure loss, thereby improving heat transfer efficiency. This symmetrical periodic wall design also helps to improve structural stability and cooling efficiency.

[0102] Among them, the cooling air is ejected from the high-pressure compressor and flows through the blade body 3 of the turbine blade, and enters the cooling channel inside the blade through the bottom of the blade. Inside the blade, the cooling air flows through the serpentine channel with ribs 8, and the inner wall of the blade is cooled by increasing the heat exchange area and turbulence intensity. The cooling air will form three airflows: one airflow mainly acts on the high-temperature area of ​​the leading edge of the blade for impact cooling, and finally flows out through the air film hole 2 at the leading edge of the blade to form an air film cooling of the front cavity 1 of the blade; the second airflow passes through the air film hole 2 in the middle area of ​​the blade to form an air film protection layer on the surface of the blade; the research focus of this embodiment is on the third cooling airflow. The strong cooling air impacts the spoiler column at the trailing edge of the blade, and is discharged from the split tail cavity 5 after flowing through the wedge-shaped internal cooling channel 6 for disturbance and enhanced heat exchange. In the heat exchange process in the trailing edge area, the cooling air flows through the spoiler column to generate a disturbance of wake-enhanced flow, destroying the development of the wall boundary layer, thereby improving the heat exchange efficiency. The spoiler column 4 itself is also a heat conductor, which can transfer the heat from the periodically symmetrical wall surface 10 of the blade to the spoiler column, and then take the heat away through the flow of fluid. In addition, the column rib spoiler cooling in the trailing edge area uses the spoiler columns 9 connecting the walls on both sides of the channel to enhance the heat exchange in the channel, which is mainly arranged in the trailing edge area of ​​the turbine blade where the lip wall 7 is thinner, and also plays a supporting role.

[0103] This embodiment is based on the analysis of the flow resistance characteristics of spoiler columns of different shapes, and proposes a new enhanced heat exchange cooling structure. As shown in the virtual frame line in the figure, for the non-circular fluid flow channel, a wall surface 10 with periodic symmetry is used, mainly for the symmetry of the distribution area of ​​the trailing edge spoiler column. The incoming flow spoiler column 4 is placed in the main incoming flow area, and the symmetrical side spoiler column 9 is located on the side symmetrical wall surface. In addition to the reassembly of the shape, a series of analysis and discussion were also carried out on the diameter of the middle incoming flow spoiler column 4, and the diameter D was finally selected. 4 =8mm,D 9=10mm. This spoiler assembly method is used to analyze its heat transfer characteristics. After a series of standardization processing and Fluent simulation calculations, it is found that when Re = 10,000, the Nu representing the overall heat transfer efficiency of this structure is about 2.84% higher than that of the vertical spoiler used in actual applications, the friction coefficient f is reduced by about 44.91%, and the comprehensive heat transfer coefficient TP increases by about 16.38%. The above quantitative analysis data show that the design of this enhanced cooling structure greatly improves the comprehensive heat transfer efficiency of the spoiler at the trailing edge, and it is not obtained under the condition of large friction resistance, which is of great significance for actual engineering applications.

[0104] The main purpose of this embodiment is to address the situation that the heat load and thermal stress at the trailing edge of the gas turbine are large. A research based on the latest spoiler column structure at the trailing edge is proposed, and the arrangement, tilt angle, spoiler column diameter, composite form, and tilt direction in the flow channel are used to design a cooling structure to improve the optimization design of the cooling structure of the trailing edge of the gas turbine. The advantages of this structure are high comprehensive heat exchange efficiency, high convection heat transfer coefficient on the flow channel surface, small friction coefficient, strong comprehensive heat exchange capacity, simple structure, easy processing and design, and low production cost. This embodiment has greatly improved many problems existing at the trailing edge of the gas turbine, improved the cooling efficiency at the trailing edge, and extended the service life of the gas turbine.

[0105] In the specific implementation work, in order to solve the problem of excessively high local temperature at the trailing edge of the spoiler column in actual engineering applications, this embodiment adopts the numerical simulation method to perform Fluent numerical calculation. The specific application scenario or boundary condition is that the velocity value of each node on the inlet boundary is given in the flow inlet boundary condition, and the formula is used. Substitute the inlet velocity values ​​under different Re conditions into the boundary conditions for calculation. The reference system adopts the method of adjacent boundary unit areas, and the heat value at the inlet is set to 298K. The pressure value given by the working condition is 10,1325Pa. The pressure is always expressed as a relative value. The actual pressure solved is not an absolute value, but relative to the inlet pressure. Therefore, in some cases, the given inlet pressure can be set to 0 to solve the pressure at other locations. The setting of the outlet boundary condition is often used in conjunction with the inlet boundary condition, and is generally set far enough away from the geometric disturbance. The setting of the recirculation zone is that the direction is from the nearby unit area, the static pressure recirculation pressure is used, and the average pressure is set. Different from the inlet, the outlet heat value is 300K. The no-slip wall boundary condition is used on the wall, and the wall velocity is 0. And the Y+ value is controlled within 1 to accurately capture the area that is more obviously affected by viscosity. At this time, it is required to divide a very precise grid in the spoiler column area, that is, to do local encryption processing. In the grid interface, the symmetry and periodic boundary conditions are used in combination, which is very common in practical engineering applications. According to the excellent symmetry of this physical model, the symmetry condition can effectively avoid solving the entire calculation domain, thereby greatly reducing the calculation scale. Moreover, on the symmetric boundary, the velocity of the vertical wall is 0, and its boundary value is equal inside and outside the flow channel. The periodic boundary condition is also proposed for the symmetry problem. The setting of residuals in calculation monitoring is crucial to the convergence of the calculation results. In order to ensure the calculation accuracy and convergence, the energy equation is used with 1x10 -6 Convergence criteria, for the continuity equation and turbulent kinetic energy equation, 1x10 -4 The heat transfer effect is very obvious. Specifically, when Re = 10,000, Nu, which represents the overall heat transfer efficiency, is about 2.84% higher than the vertical spoiler column used in actual applications, the friction coefficient f is reduced by about 44.91%, and the comprehensive heat transfer coefficient TP increases by about 16.38%.

[0106] The technical solution provided by the present invention is further described in detail above through several specific implementation modes in order to highlight the advantages and benefits of the technical solution provided by the present invention. However, the several specific implementation modes described above are not intended to be used as limitations on the present invention. Any reasonable modification and improvement of the present invention, combination of implementation modes and equivalent substitution within the spirit and principle of the present invention shall be included in the protection scope of the present invention.

Claims

1. A gas turbine trailing edge heat exchange enhanced cooling structure based on resistance characteristic analysis, characterized in that: include: A cooling air compressor for providing high-pressure cooling air; A blade, wherein a cooling passage is provided between the blade and the compressor; At least two rows of spoiler columns are arranged in the cooling channel to increase the turbulence intensity; The split tail cavity is arranged in the trailing edge area of ​​the blade and is used to guide cooling air to form an air film layer.

2. The gas turbine trailing edge heat exchange enhanced cooling structure based on resistance characteristic analysis according to claim 1, characterized in that: The cooling channel is a serpentine cooling channel.

3. The gas turbine trailing edge heat exchange enhanced cooling structure based on resistance characteristic analysis according to claim 1, characterized in that: The cross section of the spoiler column is cylindrical or elliptical.

4. The gas turbine trailing edge heat exchange enhanced cooling structure based on resistance characteristic analysis according to claim 1, characterized in that: The spoiler columns are arranged at periodic intervals.

5. The gas turbine trailing edge heat exchange enhanced cooling structure based on resistance characteristic analysis according to claim 1, characterized in that: The inner wall of the cooling channel is a rough surface.

6. The gas turbine trailing edge heat exchange enhanced cooling structure based on resistance characteristic analysis according to claim 1, characterized in that: The split tail cavity has at least two outlets.

7. A method for designing a heat exchange enhanced cooling structure for a gas turbine trailing edge based on resistance characteristic analysis, characterized in that: Used to design the structure described in claim 1, comprising: Steps of optimizing the structure of the cooling channel by numerical simulation; The step of determining the tortuosity of the serpentine cooling channel by using an optimization algorithm based on the flow path of the cooling air; The step of optimizing the structural parameters of the slit tail cavity includes the slit width and the outlet position.

8. A computer storage medium for storing a computing program, characterized in that: When the computer reads the computer program, the computer executes the method of claim 7.

9. A computer, comprising a processor and a storage medium, characterized in that: When the processor reads the computer program stored in the storage medium, the computer executes the method of claim 7 .

10. A computer program product, being a computer program, characterized in that When the computer program is executed, the method of claim 7 is implemented.

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