Film hole design method for improving cooling performance of multi-channel double-layer wall at front edge of turbine blade

Abstract

The invention discloses a film hole design method for improving the cooling performance of a multi-channel double-layer wall at the front edge of a turbine blade. The film hole design method is suitable for high-pressure turbine blades of aero-engines and heavy duty gas turbines. According to the method, a fillet structure in smooth transition is machined at an inlet of a film hole, the fillet structure is distributed in the arc length direction of a channel array and the height direction of a blade and is smoothly connected with the inner wall of the hole and the target surface of a channel, and streamline optimization of a cooling working medium inflow path is achieved under the condition that the diameter and the inclination angle of the film hole are not changed. According to the invention, backflow vortexes in holes can be effectively inhibited, the air film covering stability and uniformity are improved, and the cooling working medium flow resistance and the system air supply energy consumption are reduced, so that the blade leading edge cooling effect is improved, and the service life of parts is prolonged.

Description

Technical Field

[0001] This invention belongs to the field of mechanical engineering technology, specifically relating to an optimized design method for film cooling of multi-channel double-walled structures at the leading edge of high-pressure turbine blades of aero-engines or gas turbines, and for reducing the pressure loss of the cooling medium. Background Technology

[0002] The core of the development of modern advanced aero-engines and heavy-duty gas turbines (hereinafter referred to as "the two engines") lies in continuously improving the thermal-to-work conversion efficiency while ensuring the long service life and high reliability of hot-end components. According to the Brayton cycle theory, increasing the inlet gas temperature of the high-pressure turbine can significantly improve the cycle's thermal efficiency. However, excessively high temperatures can cause severe thermal loads on hot-end components such as turbine blades, affecting their service safety and durability. Therefore, how to improve thermal efficiency while ensuring the structural integrity of hot-end components has become a key challenge in the development of the two engine technologies.

[0003] The leading edge of high-pressure turbine blades directly bears the impact and stagnant heat transfer of high-temperature combustion gases, making it a region of concentrated heat load. This makes it susceptible to failure modes such as ablation, cracking, coating peeling, and sulfide corrosion, threatening the overall operational safety of the turbine. While a multi-channel, double-walled cooling structure (hereinafter referred to as "multi-channel structure") can effectively improve blade cooling efficiency, its small aspect ratio film cooling holes are easily affected by internal backflow vortices, leading to decreased film coverage stability and insufficient cooling uniformity. Simultaneously, the complex flow channel structure and the additional resistance from backflow vortices increase the pressure loss of the cooling medium, forcing the gas supply system to increase its supply pressure, increasing compression system energy consumption, and indirectly reducing overall turbine efficiency. Summary of the Invention

[0004] To overcome the shortcomings of the prior art, the present invention aims to provide a method for designing film cooling holes to improve the cooling performance of multi-channel double-walled turbine blades at the leading edge. This method addresses the problems of backflow vortex interference, high flow resistance of the cooling medium, and high system energy consumption inherent in multi-channel double-walled structures due to small aspect ratio film cooling holes. Cooling performance is optimized by rounding the inlet of the film cooling hole. This method effectively suppresses the development of backflow vortices within the hole, improves the film cooling coverage, reduces the flow resistance of the cooling medium, thereby reducing the additional energy consumption of the compression system and indirectly improving the overall thermal efficiency of the turbine.

[0005] To achieve the above objectives, the technical solution adopted by the present invention is as follows: A method for designing film cooling holes to improve the cooling performance of multi-channel double-walled structures at the leading edge of turbine blades involves machining a smooth-transition rounded structure at the inlet of each row of leading-edge film cooling holes corresponding to each double-walled channel in the multi-channel double-walled structure of the high-pressure turbine blade leading edge region. This rounded structure is distributed along the arc length direction of the double-walled channel array and the blade height direction, and further smoothly connects with the inner wall of the leading-edge film cooling holes and the target surface of the double-walled channels. This achieves streamlined optimization of the cooling medium inflow path, suppressing the generation and development of backflow vortices from a flow mechanism perspective.

[0006] Furthermore, the rounded structure does not change the original leading edge film cooling hole diameter and the leading edge film cooling hole tilt angle and outlet direction, maintaining the original film cooling design characteristics and ensuring that the film adhesion path and coverage effect on the blade surface are not affected.

[0007] Furthermore, under the premise of satisfying the blade wall strength and process feasibility, by reasonably matching the ratio of the rounding radius to the diameter of the leading edge film air hole (recommended range 10%–30%) and the ratio of the rounding radius to the average width of the double-wall channel (recommended range 5%–20%), the effect of suppressing backflow vortices and optimizing drag can be further enhanced.

[0008] Furthermore, the diameter of the leading edge air film aperture is 0.5-0.7 mm, and the inclination angle of the leading edge air film aperture is 40°-60°.

[0009] The design method of this invention only rounds off the inlet region of the leading edge air film hole, while the double-walled channel and impact hole structure remain unchanged.

[0010] In existing technologies, the blade walls of double-walled channels are typically formed using casting methods, and the film gas vents are generally formed into sharp inlet structures through a single laser drilling process. The rounded structure described in this invention is achieved through a laser drilling process. Specifically, to achieve the rounded structure, based on the single-stage forming of the sharp inlet, the following multi-stage laser processing method is employed according to design dimensional requirements: 1. First, a large-scale, low-energy laser treatment is applied to the inlet area of ​​the air film pore to initially form a rounded inlet edge; here, low energy refers to the energy that can form a rounded inlet edge on the material. The selection varies depending on the material, but it should be common knowledge to those skilled in the art.

[0011] 2. Subsequently, the laser's effective range is gradually narrowed and the energy is increased to finely shape the entrance area, achieving a smooth, rounded structure.

[0012] 3. Finally, the laser's effective range is narrowed to match the diameter of the air film aperture, achieving a smooth connection between the rounded structure and the inner wall of the air film aperture, thus completing the high-precision machining of the rounded inlet air film aperture.

[0013] The present invention has the following beneficial effects: By rounding the inlet of the film cooling hole across the entire range, the backflow vortex within the hole can be significantly suppressed, eliminating its lifting effect on the cooling medium, improving the adhesion stability and uniformity of the film cooling on the blade surface, and enhancing the leading-edge thermal protection capability. Simultaneously, the reduction in flow resistance decreases the demand for air supply pressure and the energy consumption of the compression system, indirectly improving the overall thermal efficiency. Compared to traditional sharp inlet designs, this invention achieves synergistic optimization of internal flow characteristics and external cooling performance, significantly improving system cooling efficiency and economy. Attached Figure Description

[0014] Figure 1 A schematic diagram of a three-dimensional model of a high-pressure turbine blade with a multi-channel double-wall cooling structure in the leading edge region.

[0015] Figure 2 This is a cross-sectional and partial enlarged view of a blade with a sharp inlet film gas vent in the prior art (corresponding to...). Figure 1 (aa view).

[0016] Figure 3 This is a cross-sectional and partially enlarged view of the blade with rounded inlet film gas holes of the present invention (corresponding to...). Figure 1 (aa view).

[0017] Figure 4 The arrangement of air film pores in a solid wall in existing technology (corresponding to) Figure 2 (the bb section).

[0018] Figure 5 This is the arrangement of the rounded air film holes in the solid wall according to the present invention (corresponding to...) Figure 3 (cc section).

[0019] Figure 6 This is a three-dimensional schematic diagram of a double-walled channel with sharp air film pores in the prior art.

[0020] Figure 7 This is a three-dimensional schematic diagram of the double-walled channel with rounded air film holes of the present invention.

[0021] Figure 8 Comparison of relative total pressure and streamline distribution in a cross section of a double-walled channel under design conditions (MFR = 1.750%): (a) prior art, (b) the present invention.

[0022] Figure 9 Comparison of the effective distribution of the adiabatic film in the leading edge region of the blade under design conditions: (a) prior art, (b) the present invention.

[0023] Figure 10 Comparison of the average effectiveness of the adiabatic air film in the leading edge region of the blade under different operating conditions.

[0024] Figure 11 It is a comparison of the relative total pressure of the air supply to the front cooling chamber of the blade under different operating conditions.

[0025] Explanation of the labels in the diagram: 1-Turbine blade; 2-Front cooling chamber; 3-Leading edge stagnation line; 4-Leading edge film cooling hole; 5-Leading edge region; 6-Suction side; 7-Suction side film cooling hole; 8-Pressure side; 9-Pressure side film cooling hole; 10-Impact hole; 11-Double-walled channel; 12-High-temperature combustion side; 13-Sharp inlet of film cooling hole; 14-Average width of double-walled channel; 15-Rounded structure; 16-Diameter of leading edge film cooling hole; 17-Target surface of double-walled channel; 18-Inclination angle of leading edge film cooling hole; 19-Diameter of impact hole; 20-Thickness of double-walled channel; 21-Rounding radius; 22-Recirculation vortex. Detailed Implementation

[0026] The present invention will now be described in further detail with reference to the accompanying drawings. This embodiment uses a high-pressure turbine blade of a certain type of aero-engine as an example, but it does not constitute a limitation on the scope of protection of the present invention.

[0027] See Figure 1 The first-stage stator blade of a certain type of aero-engine high-pressure turbine blade 1 has a large number of suction-side film cooling holes 7 on the suction side 6 and a large number of pressure-side film cooling holes 9 on the pressure side 8. Four sets of multi-channel double-wall cooling structures are arranged in the leading edge region 5. Each set of multi-channel double-wall cooling structures includes several leading edge film cooling holes 4. In the figure, one set of leading edge film cooling holes 4 is arranged along the leading edge stagnation line 3.

[0028] See Figure 2 , Figure 4 and Figure 6 Each multi-channel double-walled cooling structure consists of an impact hole 10, a double-walled channel 11, and a leading-edge film cooling hole 4, connecting the high-temperature combustion gas side 12 to the front cooling chamber 2 of the turbine blade 1. The figure shows the diameter 19 of the impact hole and the thickness 20 of the double-walled channel. As can be seen from the figure, the leading-edge film cooling hole 4 has a sharp inlet 13, meaning that the leading-edge film cooling hole 4 and the target surface 17 of the double-walled channel have a sharp angle. This sharp inlet 13 is commonly used in existing technologies, mainly because: 1. Design Factors: Although multi-channel double-wall cooling structures are existing technology, they are still a relatively new cooling layout developed in recent years, and their design concepts are not yet fully mature. Adopting a sharp inlet design for film cooling holes is a feasible solution adopted under current design thinking to meet cooling requirements.

[0029] 2. Processing technology factors: Existing air film holes are mostly processed by electrical discharge machining, which is a mature technology with low cost.

[0030] In summary, since the influence of the leading edge film air hole inlet structure on the cooling performance of multi-channel structures has not yet been discovered or paid attention to, and given the influence of current mainstream processing technology, the current leading edge film air hole 4 generally adopts the structure of sharp film air hole inlet 13.

[0031] This invention is the first to discover that, through reasonable optimization of the film cooling hole inlet structure, backflow vortices can be significantly suppressed and flow resistance reduced, which has significant engineering value for improving the cooling performance and system economy of multi-channel structures. However, existing electrical discharge machining (EDM) methods are prone to remelting, resulting in limited machining accuracy and making it difficult to achieve precision hole manufacturing. Therefore, this invention further discovers that ultrafast lasers, water-guided lasers, and other laser drilling methods can be applied to film cooling hole manufacturing. Laser processing has no thermal effects and can achieve high-precision hole manufacturing, meeting the processing requirements of this invention. In summary, the main contribution of this application lies in discovering the relationship between the film cooling hole inlet structure and the cooling performance of multi-channel structures, modifying the existing sharp inlet structure of film cooling holes, and providing a new processing technology. Ultimately, without significantly increasing manufacturing costs and process complexity, the cooling performance of multi-channel structures is improved by optimizing the film cooling hole inlet structure.

[0032] For details, please refer to Figure 3 , Figure 5 and Figure 7 In the multi-channel double-wall structure of the leading edge region 5 of the turbine blade 1, the present invention provides a rounded structure 15 at the inlet position of each column of leading edge film holes 4 corresponding to each double-wall channel 11.

[0033] See Figure 3 and Figure 5 In this invention, the rounded structure 15 is distributed along the arc length direction and the blade height direction of the double-walled channel array with respect to the holes.

[0034] See Figure 3 and Figure 5 In this invention, the rounded structure 15 is smoothly connected to the inner wall of the leading edge air film hole 4 and the double-walled channel target surface 17, and no longer has a sharp angle.

[0035] See Figure 3 and Figure 5 The rounded structure 15 in this invention does not change the connection between the leading edge gas film hole 4 and the high-temperature gas combustion side 12 outside the blade, thus ensuring the original outflow path of the cooling working fluid.

[0036] See Figure 3 and Figure 5 In this invention, the rounding radius 21 is 5%-20% of the average width 14 of the double-walled channel, and can be adjusted according to the processing accuracy. An appropriate rounding radius 21 is beneficial to optimizing the flow structure of the cooling medium.

[0037] See Figure 4 and Figure 5 In this invention, the rounding radius 21 is 10%-30% of the diameter 16 of the leading edge air film hole. The appropriate value can be selected according to the cooling performance requirements and processing costs. An appropriate rounding radius 21 is beneficial to balancing the backflow vortex suppression effect and structural strength.

[0038] See Figure 4 and Figure 5 The diameter of the rounded inlet air film hole of the present invention is consistent with the diameter of the sharp inlet air film hole (i.e., the diameter of the leading edge air film hole 16) in the prior art, and the value range is 0.5-0.7mm, which is compatible with the average width 14 of the double-wall channel.

[0039] See Figure 4 and Figure 5 The inclination angle of the rounded inlet film orifice of the present invention is consistent with the inclination angle of the sharp inlet film orifice (i.e., the inclination angle of the leading edge film orifice 18) in the prior art, and the value range is 40°-60°, ensuring that the cold air outflow direction remains unchanged.

[0040] The technical principle of this invention is as follows: Step 1: Construct a three-dimensional model of the high-pressure turbine blade 1 with a multi-channel double-wall cooling design in the leading edge region 5.

[0041] In an embodiment of the present invention, the thickness 20 of the double-walled channel is 0.3 mm, the diameter 16 of the leading edge air film hole is 0.5 mm, the inclination angle 18 of the leading edge air film hole is 40°, the diameter 19 of the impact hole is 0.5 mm, and the radius 21 of the inlet rounding 15 of the air film hole is 0.125 mm.

[0042] Step 2: Generate computational meshes for the 3D models of the sharp inlet blade and the rounded inlet blade, respectively.

[0043] In embodiments of the present invention, Ansys-Fluent Meshing software is used to generate polyhedral unstructured meshes for CFD calculations. Specifically, the fluid computation domain is extracted, and prismatic boundary layer meshes are generated for the mesh near the wall to satisfy Realizable properties. k - ε Turbulence model for y + To meet the requirement of ≈1, check the mesh quality.

[0044] Step 3: Numerical simulations are performed on the film cooling performance of sharp inlet blades and rounded inlet blades under different relative flow rates (MFR) of the cooling medium. The adiabatic temperature of the leading edge region 5, the streamlines inside the double-walled channel, and the relative total pressure of the air supplied to the front cooling chamber are calculated. C pt Obtain the effective response value of the thermal insulation membrane. η .

[0045] In an embodiment of the present invention, Ansys-Fluent software was used to conduct numerical simulations on a high-pressure turbine blade model. The steady-state RANS method was employed to simulate the cooling process at different temperatures at the main high-temperature gas inlet and the cold gas chamber inlet. For example, the total temperature at the main gas inlet was controlled at 2048.53 K, and the total temperature at the cold gas chamber inlet was controlled at 745 K, both being actual operating parameters of the high-pressure turbine blade of this aero-engine.

[0046] In embodiments of the present invention, Ansys-CFD Post software is used for post-processing to obtain the streamline distribution, pressure distribution, and supply pressure of the front cooling chamber within the double-walled channel. The adiabatic temperature distribution on the blade surface under different relative flow rates (MFR) of the cooling fluid is then extracted, thereby calculating the corresponding adiabatic film effectiveness. η This allows us to determine the cooling performance of the rounded inlet blades.

[0047] In this embodiment, relative flow rate (MFR) is used as the control index for the variable flow rate of the cooling medium. Specifically, relative flow rate (MFR) can be expressed as: In the formula, m c This represents the total flow rate of the cooling working fluid through the leading edge film cooling holes. m g This represents the mainstream gas inlet flow rate.

[0048] This embodiment uses a relative total pressure. C pt As a key indicator for evaluating the magnitude of the air supply pressure in the front cooling chamber, specifically, the relative total pressure... C pt It can be represented as: In the formula, Pt c The total pressure of the local cooling working fluid. Ps g The total pressure of the main gas outlet.

[0049] This embodiment utilizes the effectiveness of the insulating air film. η As a key indicator for evaluating the performance of leading-edge film cooling, specifically, the effectiveness of the adiabatic film... η It can be represented as: In the formula, Tt g The main gas inlet total temperature, Tt c This is the total temperature at the inlet of the cold air chamber. T AdiThis is the adiabatic temperature of the blade surface.

[0050] Step 4: Perform post-processing analysis on the obtained blade cooling calculation results and compare the effectiveness of the adiabatic film on the blade surface. η The relative total pressure of the air supply to the front air chamber C pt The rounded inlet blade design resulted in a higher effectiveness of the adiabatic film in the leading edge region. η Distribution and lower front air chamber supply relative to total pressure C pt Under different relative flow rates (MFR) of the cooling medium, the film cooling performance of the leading edge region was significantly optimized, while the pressure loss of the cooling medium was effectively reduced.

[0051] See Figure 8 In (a), in the existing sharp-inlet film cooling design, the cooling medium first enters the inlet cooling chamber 2 inside the blade, and then forms an impinging jet through the impinging hole 10, impacting the target surface 17 of the double-walled channel for localized enhanced heat transfer. After completing the impinging cooling effect, the cooling medium flows out of the blade through the film cooling hole 4 at the sharp inlet 13, forming a cooling film. The results show that in the existing technology, the sharp inlet 13 of the film cooling hole 4 causes a sudden change in the flow area of ​​the cooling medium, forming an adverse pressure gradient, generating flow separation, and subsequently inducing a backflow vortex 22 inside the hole. This vortex interferes with the outflow of the cooling medium, resulting in poor film adhesion stability and uneven coverage, which limits the cooling performance and increases the flow resistance of the cooling medium.

[0052] See Figure 8 In (b), the present invention uses a rounded inlet 15 instead of a sharp inlet 13 to control the smooth transition of the cooling medium flow area, reduce the reverse pressure gradient, suppress the intensity of the backflow vortex 22 from the flow mechanism, optimize the flow path of the cooling medium, reduce flow resistance, thereby improving the uniformity of gas film coverage and reducing the pressure loss of the cooling medium.

[0053] See Figure 9 In (a) and (b), compared with the sharp inlet design of the prior art, the rounded inlet design of this embodiment significantly reduces the momentum of the cooling working fluid at the outlet of the film cooling hole in the blade height direction, and improves the effectiveness of the adiabatic film near the outlet of the film cooling hole. η Significantly improved, especially on the suction side.

[0054] See Figure 10 Compared with the sharp inlet design of the prior art, the rounded inlet design of this embodiment has a higher average adiabatic film effectiveness in the leading edge region under the five relative flow rates (MFR) of the cooling working fluid. ηWhen the relative flow rate (MFR) of the cooling working fluid is 0.875%, 1.3125%, 1.75%, 2.1875%, and 2.625%, respectively, the average adiabatic film effectiveness of the leading edge region of the rounded inlet design is... η Compared with the sharp entry design, the improvements were 5.79%, 7.09%, 6.39%, 3.17%, and 3.79%, respectively.

[0055] See Figure 11 Compared to the sharp inlet design of the prior art, the rounded inlet design of this embodiment has a lower relative total pressure of the front cooling chamber supply air under all five relative flow rates (MFR) of the cooling working fluid. C pt When the relative flow rate (MFR) of the cooling working fluid is 0.875%, 1.3125%, 1.75%, 2.1875%, and 2.625%, respectively, the relative total pressure of the air supplied to the front cooling chamber with the rounded inlet design is... C pt Compared with the sharp entry design, the costs were reduced by 1.10%, 3.20%, 5.64%, 7.75%, and 9.62%, respectively.

[0056] The above description is only a preferred embodiment of the present invention. For those skilled in the art, without departing from the technical principle of the present invention, parameters such as the rounding radius 21 can be adjusted according to the cooling performance requirements and actual processing capabilities. These improvements should be considered within the scope of protection of the present invention.

Claims

1. A method for designing film cooling holes to improve the cooling performance of multi-channel double-walled turbine blades at the leading edge, characterized in that: A rounded structure (15) is provided at the inlet of each column of leading edge film holes (4) of the multi-channel double-wall structure at the leading edge of the turbine blade (1). The rounded structure (15) is distributed along the arc length direction of the double-wall channel array and the blade height direction with the holes.

2. The air film pore design method according to claim 1, characterized in that: The radius (21) of the rounded structure (15) is 10%-30% of the diameter (16) of the leading edge air film hole.

3. The air film pore design method according to claim 1, characterized in that: The radius (21) of the rounded structure (15) is 5%-20% of the average width (14) of the double-walled channel.

4. The air film pore design method according to claim 1, characterized in that: The rounded structure (15) is smoothly connected to the inner wall of the leading edge air film hole (4) and the double-wall channel target surface (17), without changing the original leading edge air film hole diameter (16) and leading edge air film hole tilt angle (18).

5. The air film pore design method according to claim 4, characterized in that: The diameter (16) of the leading edge air film hole is 0.5-0.7 mm, and the inclination angle (18) of the leading edge air film hole is 40°-60°.

6. The air film pore design method according to claim 1, characterized in that: The design method only rounds the inlet area of ​​the leading edge air film hole (4), while the structure of the double-walled channel (11) and the impact hole (10) remains unchanged.

7. The air film pore design method according to claim 1, characterized in that: The rounded structure (15) is achieved by laser drilling.

8. The air film pore design method according to claim 7, characterized in that: The laser drilling process is as follows: First, a large-scale, low-energy laser treatment is applied to the inlet area of ​​the leading edge air film hole (4) to initially form a rounded inlet edge; Subsequently, the laser's effective range was gradually narrowed and the energy was increased to finely shape the entrance area and achieve a smooth, rounded structure (15). Ultimately, the laser action range is narrowed to be consistent with the diameter (16) of the leading edge air film hole, so as to achieve a smooth connection between the rounded structure (15) and the inner wall of the leading edge air film hole (4) and the double-walled channel target surface (17).

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