A new hole-slot combined spoiler cooling structure for turbine blade and application thereof

Abstract

The application discloses a novel hole-slot combined spoiler cooling structure for a turbine blade and application, and belongs to the cooling technical field of a turbine blade of a gas turbine engine. The novel hole-slot combined spoiler cooling structure comprises an upper end wall and a lower end wall forming a cooling channel and a spoiler array arranged in the cooling channel. The upper end surface of the spoiler array is fixedly connected with the upper end wall, and the lower end surface is fixedly connected with the lower end wall. The spoiler is a non-closed ring structure. The ring surface side of the spoiler faces the inlet of the cooling channel. The outer ring surface is a windward surface, and the inner ring surface is a leeward surface. The opening side of the spoiler faces the cold gas outlet of the cooling channel. A plurality of through holes and through slots are formed in the ring surface of the spoiler. The cooling air is injected to the end wall on the leeward side through the through holes, and the cold gas on the windward surface and the cold gas on the leeward surface are mixed with each other through the rectangular slots. The novel hole-slot combined spoiler cooling structure overcomes the deficiency of the traditional cylindrical spoiler column in the heat exchange capacity, and improves the problems of the weak heat exchange around the spoiler column and the uneven distribution of the wall heat exchange coefficient caused by the flow separation.

Description

Technical Field

[0001] This invention belongs to the field of cooling technology for gas turbine engine turbine blades, specifically relating to a novel perforated slotted spoiler cooling structure for turbine blades and its application. Background Technology

[0002] Gas turbine engines are widely used as power output devices in aviation propulsion, marine propulsion, and land-based power generation. To achieve higher engine thrust and efficiency, the turbine inlet gas temperature is continuously being increased. Currently, the turbine inlet gas temperature of some advanced aero engines exceeds 2000K, a temperature far exceeding the temperature resistance limit of turbine blade materials. To ensure the safe and stable operation of turbine blades, advanced and efficient blade cooling technologies must be explored. The trailing edge, as a critical part of the turbine blade, has a thinner structure that ensures good aerodynamic performance, but its weaker structural strength and higher wall temperature make it more prone to ablation and fracture. Cooling measures for the trailing edge typically involve arranging a group of turbulence-inducing columns throughout the internal channel. These columns strengthen the structural strength of the trailing edge and enhance the disturbance of the cold airflow near the wall, thereby improving the heat transfer coefficient of the inner wall and reducing the wall temperature.

[0003] Existing technologies disclose the placement of cylindrical turbulence columns within the trailing edge channel, and have studied the heat transfer and flow characteristics of these circular turbulence columns, analyzing the influence of flow and structural parameters. The impact of the arrangement of the turbulence columns on enhancing end-wall heat transfer has also been analyzed. However, research on turbulence elements in the industry over the past few decades indicates that traditional cooling structures typically enhance heat transfer by generating large-scale turbulent vortices in the flow field, at the cost of significant flow losses. Furthermore, the actual increase in heat transfer coefficient is not substantial; the enhanced heat transfer mainly relies on the increase in heat transfer area, resulting in an uneven distribution of the obtained heat transfer coefficient, which is the main cause of localized thermal stress. Summary of the Invention

[0004] The technical problem to be solved:

[0005] To overcome the shortcomings of existing technologies, this invention provides a novel perforated and slotted spoiler cooling structure and its application for turbine blades. The main feature of this structure is the presence of angled cylindrical through-holes and slots on the curved side of the spoiler. Specific structure and airflow pattern are shown in the attached figure. Figure 1As shown, after being accelerated out of the through-hole, the cold air impacts the end wall of the leeward side of the baffle at a certain angle, enhancing heat transfer on the leeward side. Simultaneously, the cold air on the windward and leeward sides can mix through the through-slit, increasing the turbulence of the airflow near the wall and improving the local cooling effect. This invention overcomes the shortcomings of traditional cylindrical baffle columns in enhancing heat transfer and improves the problems of weak heat transfer around the baffle column and uneven distribution of the heat transfer coefficient on the wall surface caused by flow separation.

[0006] The technical solution of the present invention is: a novel slotted-hole combined spoiler cooling structure for turbine blades, comprising an upper end wall and a lower end wall constituting a cooling channel, and a spoiler array disposed within the cooling channel, wherein the upper end face of the spoiler array is fixedly connected to the upper end wall of the cooling channel, and its lower end face is fixedly connected to the lower end wall of the cooling channel.

[0007] The baffle plate is a non-closed annular structure with its annular side facing the entrance of the cooling channel, the outer annular side being the windward side, and the inner annular side being the leeward side; its open side facing the cold air outlet of the cooling channel, opposite to the leeward side; its annular surface has several through holes and through slots, through which the cooling airflow is directed to the end wall on the leeward side, and through the rectangular slots to mix the cold air on the windward side and the cold air on the leeward side.

[0008] A further technical solution of the present invention is: the cross-section of the windward side of the spoiler is an arc; the middle part of the cross-section of the leeward side is an arc, and the two ends near the opening side are straight lines; and the arc of the windward side and the arc of the leeward side are concentric arcs.

[0009] A further technical solution of the present invention is: the central angle θ of the windward arc is 270°, and the central angle Θ of the leeward arc is 180°~190°.

[0010] A further technical solution of the present invention is that the intersection of the windward and leeward sides of the spoiler is a rounded transition with a radius of r = 0.5-0.15 mm.

[0011] A further technical solution of the present invention is that the angle between the through hole and the upper / lower end wall is 20° to 30°, the center of the through hole on the windward surface is located on the stagnation line of the windward surface, and the axis of each through hole intersects with the axis of the windward surface.

[0012] A further technical solution of the present invention is: a plurality of through holes on the baffle plate form a rectangular array of through holes, which are evenly distributed along the axial and circumferential directions respectively. The circumferential angle β between adjacent through holes in the circumferential direction ranges from 20° to 30°, and the distance L between adjacent through holes in the axial direction ranges from 0.4mm to 0.7mm.

[0013] A further technical solution of the present invention is: the through hole array is symmetrically arranged with respect to the mid-section of the cooling channel, the through holes located between the upper end wall and the mid-section constitute the upper through hole array, and the axis of the through holes is inclined from the mid-section on the windward side to the upper end wall on the leeward side; the through holes located between the lower end wall and the mid-section constitute the lower through hole array, and the axis of the through holes is inclined from the mid-section on the windward side to the lower end wall on the leeward side.

[0014] A further technical solution of the present invention is: the through-slot is a rectangular through-slot that runs through the windward and leeward sides, and multiple through-slots are distributed at the junction of the arc segment and the straight segment on the leeward side; the circumferential width of the through-slot is W = 0.6mm-1.2mm, the axial height is h = 0.2mm-0.5mm, and the circumferential angle from the axis to the vertical symmetrical section of the spoiler is γ = 105°-110°.

[0015] A further technical solution of the present invention is: the number of through seams is four, the four through seams are symmetrical about the mid-section of the cold air channel, and are symmetrical about the plane where the stagnation line of the windward side of the baffle is located and the windward side axis is located; the two through seams on the baffle above the mid-section of the cold air channel are adjacent to the upper end wall, and the two through seams on the baffle below the mid-section of the cold air channel are adjacent to the lower end wall.

[0016] A novel slotted spoiler cooling structure for turbine blades is applied to the internal channel of the trailing edge of the turbine blade, wherein the distance between spanwise adjacent spoilers in the spoiler array is equal to the distance between flowwise adjacent spoilers.

[0017] Beneficial effects

[0018] The beneficial effects of this invention are as follows:

[0019] 1. Compared with cylindrical turbulence columns of the same hydraulic diameter, the perforated slit turbulence plate of the present invention has a larger near-wall windward area, which can generate greater disturbance to the incoming cold air flow, enhance the turbulence of the near-wall cold air flow, and have a greater strengthening effect on surface heat transfer.

[0020] 2. The through-hole design on the baffle plate of this invention allows some of the cold air to flow smoothly to the leeward side of the baffle plate through the through-hole, reducing the flow resistance of the cold air. At the same time, some of the cold air flowing out of the through-hole can create a second impact effect on the end wall surface of the leeward side of the baffle plate, enhancing the heat transfer of the leeward end wall.

[0021] 3. The through-slit design on the baffle plate of this invention creates a flow channel on the leeward side of the baffle plate to the windward side, which enhances the mixing of cold air and strengthens the local heat exchange intensity. Attached Figure Description

[0022] Figure 1: A schematic diagram of a novel perforated slotted spoiler cooling structure for turbine blades according to the present invention;

[0023] Figure 2 : Schematic diagram of the perforated and slotted spoiler structure of the present invention;

[0024] Figure 3 Cross-sectional view of the perforated and slotted spoiler structure of the present invention;

[0025] Figure 4 This invention discloses a novel arrangement of perforated slotted spoiler cooling structures for turbine blades.

[0026] Figure 5 The surface Nusselt number (Nu) of the perforated and slotted spoiler structure of this invention is compared with that of the cylindrical spoiler structure, Re = 10000;

[0027] Figure 6 The present invention relates to a perforated and slotted spoiler structure and a cylindrical spoiler column structure with a spanwise average Nusselt number (Nu). Ave In comparison, Re = 10000.

[0028] Explanation of reference numerals in the attached diagram: 1-Break-type spoiler body, 2-Upper end wall, 3-Lower end wall, 4-Through hole, 5-Through slot, 11-Windward side, 12-Leafward side, 13-Windward side arc, 14-Leafward side arc, 15-Leafward side straight line, 16-Rounded corner, 41-Upper through hole row, 42-Lower through hole row. Detailed Implementation

[0029] The embodiments described below with reference to the accompanying drawings are exemplary and intended to explain the invention, and should not be construed as limiting the invention.

[0030] In the description of this invention, it should be understood that the terms "center," "longitudinal," "lateral," "length," "width," "thickness," "upper," "lower," "front," "rear," "left," "right," "vertical," "horizontal," "top," "bottom," "inner," "outer," "clockwise," and "counterclockwise," etc., indicate the orientation or positional relationship based on the orientation or positional relationship shown in the accompanying drawings. They are only for the convenience of describing this invention and simplifying the description, and do not indicate or imply that the device or element referred to must have a specific orientation, or be constructed and operated in a specific orientation. Therefore, they should not be construed as limitations on this invention.

[0031] Addressing the shortcomings of existing technologies where enhanced heat transfer primarily relies on increased heat transfer area, resulting in uneven heat transfer coefficient distribution and insufficient heat transfer enhancement capacity of traditional cylindrical baffles, this invention provides a novel perforated and slotted baffle cooling structure for turbine blades. This structure features a baffle array within a cold air channel, with angled cylindrical impact holes and rectangular slots formed on the curved baffles. After being accelerated out through the impact holes, the cold air impacts the end wall of the baffle's leeward side at a specific angle, enhancing heat transfer on the leeward side. Simultaneously, the cold air from the windward and leeward sides mixes through the rectangular slots, increasing turbulence near the wall and improving localized cooling. The specific technical solution is as follows:

[0032] Reference Figure 1 As shown in the figure, this embodiment discloses a novel perforated slotted spoiler cooling structure for turbine blades, comprising: multiple spoiler bodies 1, an upper end wall 2, and a lower end wall 3. The upper end wall 2 and the lower end wall 3 together form a cold air flow channel, which includes a cold air inlet and a cold air outlet. The spoiler bodies 1 are arranged in the cold air flow channel, and the upper wall surface of the spoiler body 1 is fixedly connected to the upper end wall 2, and the lower wall surface of the spoiler body 1 is fixedly connected to the lower end wall 3.

[0033] Reference Figure 2 As shown, the spoiler body 1 is a thin sheet with a non-closed annular cross-section. The surface of the sheet facing the cold air inlet is the windward side 11, and the surface of the spoiler body 1 facing the cold air outlet is the leeward side 12. The cross-sectional line of the windward side 11 is an arc 13; the cross-sectional lines of the leeward side 12 are an arc 14 and a straight line 15, and the arcs 13 and 14 of the cross-sectional lines of the windward side 11 and the leeward side 12 are concentric arcs. The central angle θ of the arc 13 is 270°, and the central angle Θ of the arc 14 is 180°-190°. The two corners at the tail of the spoiler body 1 are designed as rounded corners 16 (r = 0.5-0.15) to reduce local stress distribution. The spoiler body 1 is provided with a through hole 4 and a through slot 5.

[0034] Furthermore, refer to Figure 3 As shown, the spoiler body 1 is symmetrical about the mid-section of the cold air flow channel.

[0035] Reference Figure 3 As shown, the through hole 4 is a cylindrical through hole with an inclination angle α ranging from 20° to 30°. This design aims to allow the airflow to impact the end wall 1 and end wall 2 at a certain angle after passing through the through hole 4. The center of the through hole 4 on the windward surface 11 is located on the stagnation line of the windward surface 11, and the axis of the through hole 4 intersects the axis of the windward surface 11.

[0036] The perforated holes 4 on the baffle plate 1 form a rectangular array of perforated holes, which are evenly distributed along the axial and circumferential directions. The circumferential angle β between adjacent perforated holes in the circumferential direction ranges from 20° to 30°, and the distance L between adjacent perforated holes in the axial direction ranges from 0.4 mm to 0.7 mm. These parameters limit the formation of a more uniform impact cooling effect.

[0037] The through-hole rows are symmetrically arranged with respect to the mid-section of the cooling channel. The through holes located between the upper end wall and the mid-section constitute the upper through-hole row 41, and the axes of the through holes are inclined from the mid-section on the windward side towards the upper end wall on the leeward side. The through holes located between the lower end wall and the mid-section constitute the lower through-hole row 42, and the axes of the through holes are inclined from the mid-section on the windward side towards the lower end wall on the leeward side. The upper through-hole row 41 and the lower through-hole row 42 are symmetrical about the mid-section of the cooling channel.

[0038] The through-slot 5 is a rectangular wide slot that runs through both the windward and leeward sides, and multiple through-slots 5 are arranged at the junction of the arc 14 and the straight line 15 of the leeward side 12 cross-section. The circumferential angle, width, and height of the through-slot range from γ = 105°-110°, W = 0.6mm-1.2mm, and h = 0.2mm-0.5mm, respectively.

[0039] Furthermore, the plurality of through seams 5 are close to the upper end wall 2 and the lower end wall 3, and the plurality of through seams 5 are simultaneously symmetrical about the plane formed by the mid-section of the cold air channel and the stagnation line of the windward surface 11 and the axis of the windward surface 11.

[0040] The above technical solution will be further explained below with reference to the accompanying drawings.

[0041] Example 1: This example is a novel perforated slotted spoiler cooling structure for turbine blades, including multiple spoiler bodies 1, an upper end wall 2, and a lower end wall 3. The upper end wall 2 and the lower end wall 3 together form a cold air flow channel, which includes a cold air inlet and a cold air outlet. The multiple spoiler bodies 1 are staggered in the cold air flow channel, and the upper wall surface of the spoiler body 1 is connected to the upper end wall 2, and the lower wall surface of the spoiler body 1 is connected to the lower end wall 3.

[0042] The specific dimensions and implementation method of the perforated and slotted spoiler are as follows:

[0043] The height of the spoiler body 1 is H = 4.0 mm. The outer diameter of the arc end 11 is R1 = 2 mm, the inner diameter is R2 = 1.5 mm, the central angle of the outer arc is θ = 270°, the central angle of the inner arc is Θ = 187°, and the radius of the fillet is r = 0.1 mm. The diameter of the through hole 4 is D = 0.2 mm, the inclination angle is α = 30°, the distance from the center of the through hole 4 on the windward side surface to the end wall is L0 = 0.4 mm, the spacing between the through hole rows 41 is L = 0.4 mm, and the circumferential angle between the through holes 4 is β = 27°. The circumferential angle, width, and height of the through slot are γ = 108°, W = 0.9 mm, and h = 0.4 mm, respectively.

[0044] Furthermore, the perforated and slotted spoilers are arranged on the internal channel of the turbine blade trailing edge. For the same row of spoilers, the spanwise distance between each spoiler is P. Y =8.0mm, the flow distance between the front and rear rows of the perforated and slotted baffle is P. X =8mm, and the arrangement of the perforated and slotted spoiler is a cross-arrangement.

[0045] Example 2: This example is a novel perforated slotted spoiler cooling structure for turbine blades, including multiple spoiler bodies 1, an upper end wall 2, and a lower end wall 3. The upper end wall 2 and the lower end wall 3 together form a cold air flow channel, which includes a cold air inlet and a cold air outlet. The multiple spoiler bodies 1 are staggered in the cold air flow channel, and the upper wall surface of the spoiler body 1 is connected to the upper end wall 2, and the lower wall surface of the spoiler body 1 is connected to the lower end wall 3.

[0046] The specific dimensions and implementation method of the perforated and slotted spoiler are as follows:

[0047] The height of the spoiler body 1 is H = 4.0 mm. The outer diameter of the arc end 11 is R1 = 2 mm, and the inner diameter is R2 = 1.5 mm. The central angle of the arc 13 is θ = 270°, the central angle of the arc 14 is Θ = 187°, and the fillet radius is r = 0.1 mm. The diameter of the through hole 4 is D = 0.4 mm, the inclination angle is α = 30°, the distance from the center of the through hole 4 on the windward side surface to the end wall is L0 = 0.5 mm, the spacing between the through hole rows 41 is L = 0.5 mm, and the circumferential angle between the through holes 4 is β = 27°. The circumferential angle, width, and height of the through slot are γ = 108°, W = 0.9 mm, and h = 0.4 mm, respectively.

[0048] Furthermore, the perforated and slotted spoilers are arranged on the internal channel of the turbine blade trailing edge. For the same row of spoilers, the spanwise distance between each spoiler is P. Y =8.0mm, the flow distance between the front and rear rows of the perforated and slotted baffle is P. X=8.0mm, and the arrangement of the perforated and slotted spoiler is a cross-arrangement.

[0049] like Figure 5 The comparison shows the surface Nusselt number distribution of Embodiments 1 and 2 of the novel perforated slit-type baffle cooling structure of the present invention with that of a cylindrical baffle cooling structure with the same hydraulic diameter when the dimensionless cold gas mass flow rate Re = 10000. It can be found that the novel perforated slit-type baffle cooling structure of the present invention has better heat transfer performance on the windward side than the cylindrical baffle element structure. Simultaneously, due to the arrangement of the impact holes, a local high heat transfer zone appears on the leeward side of the novel perforated slit-type baffle cooling structure of the present invention, and the heat transfer effect in this region is also better than that of the traditional cylindrical baffle cooling structure.

[0050] like Figure 6 The diagram compares the spanwise average Nusselt number distribution of the novel perforated slit-type baffle cooling structure of this invention (Examples 1 and 2) with that of a cylindrical baffle cooling structure of the same hydraulic diameter at a dimensionless cold gas mass flow rate Re = 10000. It can be observed that the overall heat transfer intensity of the novel perforated slit-type baffle cooling structure is superior to that of the cylindrical baffle cooling structure. The average Nusselt number increase in Example 1 is the largest, reaching 70.7%, and in Example 2, the increase is the largest, reaching 45.2%.

[0051] Although embodiments of the present invention have been shown and described above, it is understood that the above embodiments are exemplary and should not be construed as limiting the present invention. Those skilled in the art can make changes, modifications, substitutions and variations to the above embodiments within the scope of the present invention without departing from the principles and spirit of the present invention.

Claims

1. A novel perforated-slot combined spoiler cooling structure for turbine blades, characterized in that: It includes an upper end wall and a lower end wall that constitute a cooling channel, and an array of baffles disposed in the cooling channel. The upper end face of the baffle array is fixedly connected to the upper end wall of the cooling channel, and its lower end face is fixedly connected to the lower end wall of the cooling channel. The baffle is a non-closed annular structure with its annular side facing the entrance of the cooling channel, the outer annular side being the windward side, and the inner annular side being the leeward side; its open side facing the cold air outlet of the cooling channel and opposite to the leeward side; its annular surface has several through holes and through slots, through which the cooling airflow is directed to the end wall on the leeward side, and through the rectangular slots to mix the cold air on the windward side and the cold air on the leeward side. The cross-section of the windward side of the spoiler is an arc; the cross-section of the leeward side is an arc in the middle and straight at both ends near the opening; and the arcs of the windward side and the arcs of the leeward side are concentric arcs. The angle between the through hole and the upper / lower end wall is 20° to 30°. The center of the through hole on the windward surface is located on the stagnation line of the windward surface, and the axis of each through hole intersects with the axis of the windward surface. The through-slot is a rectangular channel that runs through both the windward and leeward sides. Multiple through-slots are distributed at the junction of the arc segment and the straight segment on the leeward side. The circumferential width of the through-slot is W=0.6mm-1.2mm, the axial height is h=0.2mm-0.5mm, and the circumferential angle from the axis to the vertical symmetrical section of the spoiler is γ=105°-110°.

2. The novel perforated slotted spoiler cooling structure for turbine blades according to claim 1, characterized in that: The central angle of the arc of the windward side θ =270°, the central angle of the leeward arc. Θ =180°~190°.

3. The novel perforated slotted spoiler cooling structure for turbine blades according to claim 2, characterized in that: The windward and leeward sides of the spoiler intersect with rounded corners, with a radius of curvature of [missing information]. r =0.5-0.15mm.

4. The novel perforated slotted spoiler cooling structure for turbine blades according to claim 3, characterized in that: The perforated plate has a rectangular array of through holes, which are evenly distributed along the axial and circumferential directions, respectively. The circumferential angle between adjacent through holes in the circumferential direction is... β The value ranges from 20° to 30°, representing the spacing between adjacent through holes in the axial direction. L The value range is 0.4mm-0.7mm.

5. The novel perforated-slot combined spoiler cooling structure for turbine blades according to claim 4, characterized in that: The through-holes are symmetrically arranged with respect to the mid-section of the cooling channel. The through-holes located between the upper end wall and the mid-section constitute the upper through-holes, and the axes of the through-holes are inclined from the mid-section on the windward side to the upper end wall on the leeward side. The through-holes located between the lower end wall and the mid-section constitute the lower through-holes, and the axes of the through-holes are inclined from the mid-section on the windward side to the lower end wall on the leeward side.

6. The novel perforated slotted spoiler cooling structure for turbine blades according to claim 5, characterized in that: The number of through seams is four. The four through seams are symmetrical about the mid-section of the cold air passage and about the plane containing the stagnation line of the windward side of the baffle and the windward side axis. The two through seams on the baffle above the mid-section of the cold air passage are adjacent to the upper end wall, and the two through seams on the baffle below the mid-section of the cold air passage are adjacent to the lower end wall.

7. A novel perforated slotted spoiler cooling structure for turbine blades, as described in any one of claims 1-5, is applied to the internal channel of the trailing edge of a turbine blade, characterized in that: In the spoiler array, the distance between adjacent spoilers in the spanwise direction is equal to the distance between adjacent spoilers in the flow direction.

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