阅读说明：本技术 一种端壁气膜孔布置结构 (End wall air film hole arrangement structure ) 是由 刘钊 张韦馨 刘战胜 杨星 丰镇平 于 2020-06-29 设计创作，主要内容包括：本发明公开了一种端壁气膜孔布置结构。该结构利用端壁的流动特性对冷却单元进行布置,其气膜孔构型可随意替换,适用于多种吹风比及上游泄漏流存在的情况下的燃机透平第一级静叶端壁以及动叶端壁。通过引入冷却空气,降低端壁表面整体的温度,强化压力面根部的冷却效率,保护叶片防止烧蚀,同时最大程度上减少对端壁流场边界层的破坏,节约冷却空气用量,提升了透平的工作温度极限。整套布置方案结构简单,简化了工艺流程,可以依据上游泄漏流的情况灵活调整,可以通过调整复合角的大小,满足对不同叶片压力面区域的冷却需求,使得整体的设计能够更加具有普适的应用价值,提升多类型透平的工作效率,满足工业生产的需求。(The invention discloses an end wall air film hole arrangement structure. The structure utilizes the flow characteristic of the end wall to arrange the cooling unit, the configuration of the air film hole can be replaced at will, and the structure is suitable for the first-stage stationary blade end wall and the movable blade end wall of the combustion engine turbine under the conditions of various blowing ratios and the existence of upstream leakage flow. By introducing cooling air, the temperature of the whole surface of the end wall is reduced, the cooling efficiency of the root of the pressure surface is enhanced, the blades are protected from ablation, meanwhile, the damage to the boundary layer of the end wall flow field is reduced to the maximum extent, the using amount of the cooling air is saved, and the working temperature limit of the turbine is improved. The whole set of arrangement scheme simple structure has simplified process flow, can let out the nimble adjustment of the condition of leakage flow according to the upper reaches, can satisfy the cooling demand to different blade pressure face regions through the size of adjusting compound angle for holistic design can have universal using value more, promotes the work efficiency of polymorphic type turbine, satisfies industrial production's demand.)

1. The end wall air film hole arrangement structure is characterized in that the end wall is a first-stage stationary blade end wall of an aeroengine, and the air film holes are arranged in two rows, wherein the projection vector of the cold air momentum of the first air film hole on the plane of the end wall is parallel to the main flow direction, and covers a front edge point (0.21-0.53) C away from the lower pressure surface side_axA region (2) forming a pressure surface side (0.5 to 1.0) C with respect to the end wall after being synthesized with the secondary flow amount_axThe air film covering of the position cools the pressure surface area of the end wall; the second exhaust film hole is arranged in an air film gap formed by the first exhaust film hole and the leakage flow, and the gaps are respectively deviated by 0.06 on the two sides of the central line of the upstream leakage flow close to the side boundary of the pressure surface and the molded line of the pressure surfaceC_axWidth 0.12C_axTo form a cooling film to achieve maximum film coverage of the entire end wall, wherein C_axIs the axial chord length.

2. The endwall film hole arrangement of claim 1, wherein C_ax75.31 mm.

3. The endwall film hole arrangement of claim 1, wherein the cooling area of the endwall is divided into four parts, namely a leakage flow delta cooling area, a secondary flow cooling area, a pressure side cooling area and a leading edge cooling area, wherein the leakage flow delta cooling area is an area covered by an upstream leakage flow on the endwall surface and shows a boundary with a cooling efficiency higher than 0.3 on an adiabatic film cooling efficiency cloud chart, and the pressure side cooling area is an area in which the upstream leakage flow is shifted to the pressure side by 0.06C from the centerline of the pressure side profile close to the boundary of the pressure side to the pressure side_axThe region between the pressure surface profile and the leading edge cooling region refers to the outward deviation of the leading edge of the blade by 0.2C_axThe secondary flow cooling area refers to the position of the upstream leakage flow close to the side boundary of the pressure surface and the two sides of the central line of the pressure surface molded line are respectively deviated by 0.06C_axWidth 0.12C_axThe area of (a).

4. The endwall film hole arrangement of claim 3, wherein the hole pitch P/d between adjacent film holes of the first and second film holes is greater than 2, and the pressure side cooling zone entrance width is 0.24C_axThe intersection point of the central line and the front edge point of the blade is the central hole position of the first exhaust film hole, a perpendicular line is made at the position, and the positive and negative error between the projection width of the first exhaust film hole on the perpendicular line and the width of the pressure surface inlet area is 0.04C_axThe hole row is parallel to the main flow inlet, and the width of the second exhaust membrane hole is 0.12C_axPositive and negative error 0.02C_axAnd the hole rows are vertical to the position of the local polar limiting line, and the number of the air film holes is obtained by downward rounding according to the diameters of the air film holes, wherein P is the distance between the air film holes, and d is the diameter of the cylindrical section of the air film hole.

5. The endwall film hole arrangement of claim 1, wherein the row of film holes is oriented perpendicular to the local surface streamline of the endwall to maximize film coverage, increase cooling efficiency downstream of the film hole exit, and reduce the arrangement of film holes in the middle and aft portion of the endwall.

6. The endwall gas film hole arrangement of claim 1, wherein the gas film hole compound angle is 0 °, jet angle is 30 °, length to diameter ratio L/d is 4.56, where L is the gas film hole centerline length.

7. The end wall film hole arrangement of claim 1, wherein the gas sources for all of said film holes are uniform.

8. An end wall film hole arrangement according to any of claims 1 to 7, wherein the first film hole is at a characteristic distance of 0.15C from the leakage flow outlet_axThe center of the lower end hole is 0.21C away from the leading edge point of the stator blade_axThe arrangement is such that the row direction is parallel to the inlet boundary; the characteristic distance between the second exhaust membrane hole and the leakage flow outlet is 0.40C_axThe center of the lower end hole is 0.66C away from the leading edge point of the stator blade_axThe arrangement mode is that the included angle between the row direction and the inlet boundary is 44.9 degrees; the hole pitch P/d of the first exhaust film is 2.67, the hole pitch P/d of the second exhaust film is 2.0, wherein P is the hole pitch of the film, and d is the diameter of the cylindrical section of the film.

Technical Field

The invention belongs to the technical field of turbine blade air film cooling, and particularly relates to an end wall air film hole arrangement structure.

Background

In the pursuit of high output and efficiency, the inlet temperature of gas turbines is increasing and the thermal load to which the turbine blades are subjected is also increasing. The inlet temperature of gas turbines currently exceeds the temperature that the blades can withstand, and therefore cooling techniques are required to ensure the operational safety and longevity of the combustion engine. Common cooling techniques include blade external cooling and internal cooling techniques.

Among the cooling solutions, the blade film cooling technology starts at the end of the fifties of the last century, and has become an indispensable and important part in the cooling solutions after half a century of development, and in the arrangement solution of the blade surface discrete film air, the turbine end wall is affected by complex secondary flow due to its special position, such as channel vortex, horseshoe vortex, etc., and the cooling is difficult. In the current design, a relatively common cooling method is to form a slot or a continuous film hole at the inlet of the end wall of the turbine, then arrange discrete film holes at the end wall in the middle of the flow channel, and perform film cooling by adjusting the flow rate and the jet angle of the cooling film. This method suffers from the complex flow of the endwalls, and it is difficult to obtain effective cooling gas coverage in the area near the leading edge root and pressure face of the blade endwall. Therefore, the cooling technology of the end wall needs to be established on the premise of fully researching the flow heat exchange of the end wall, in the past research, researchers try to rearrange the end wall air film holes through the heat transfer coefficient distribution of the surface of the end wall, the Mach number lines of the main flow and the like and the secondary flow of the end wall, although the whole end wall surface is covered by good cold air, the air used for cooling is obviously used too much, meanwhile, the arrangement of the air film holes is too single, only the conditions of the end wall are considered, the positions of the blade grids in the turbine are not considered, and in the practical design application, the functions of some cooling holes need to be considered.

Disclosure of Invention

In order to overcome the defects of the prior art, the invention aims to provide an end wall air film hole arrangement structure, the arrangement of a cooling unit of a turbine end wall of a combustion engine under the condition of upstream leakage flow is redesigned, on the basis of not changing the structure, the number and the air supply mode of the air film holes, the whole air film cooling efficiency of the end wall is improved by utilizing the flow characteristic of the turbine end wall and adjusting the arrangement mode, the distance between the air film holes and other conditions, the aim is to save cooling air to the maximum extent, the original cooling condition of the end wall is not greatly influenced, the influence of the air film holes on the flow heat exchange condition of the end wall is ensured to be minimum, the flow loss of working media in a cascade is reduced, and the whole working efficiency of the turbine is improved.

In order to achieve the purpose, the invention adopts the technical scheme that:

the end wall is a first-stage stationary blade end wall of an aeroengine, and the film holes are arranged in two rows, wherein the projection vector of the cold air momentum of the first film hole in the end wall plane is parallel to the main flow direction, and covers the front edge point (0.21-0.53) C on the pressure surface side below the distance_axA region (2) forming a pressure surface side (0.5 to 1.0) C with respect to the end wall after being synthesized with the secondary flow amount_axThe air film covering of the position cools the pressure surface area of the end wall; the second exhaust film hole is arranged in an air film gap formed by the first exhaust film hole and the leakage flow, and the gaps are respectively deviated by 0.06C from the two sides of the central line of the upstream leakage flow close to the side boundary of the pressure surface and the molded line of the pressure surface_axWidth 0.12C_axTo form a cooling film to achieve maximum film coverage of the entire end wall, wherein C_axIs the axial chord length.

The cooling area of the end wall is divided into a leakage flow triangular cooling area, a secondary flow cooling area, a pressure surface cooling area and a front edge cooling area, wherein the leakage flow triangular cooling area refers to an air film covering area formed by upstream leakage flow on the surface of the end wall, the boundary with the cooling efficiency higher than 0.3 is shown on an adiabatic air film cooling efficiency cloud chart, and the pressure surface cooling area refers to the condition that the upstream leakage flow is close to the boundary of the pressure surface side and deviates from the center line of the pressure surface molded line to the pressure surface side by 0.06C_axThe region between the pressure surface profile and the leading edge cooling region refers to the outward deviation of the leading edge of the blade by 0.2C_axThe secondary cooling zone refers to the upstream leakage flow near the pressure surfaceThe two sides of the central line of the side boundary and the pressure surface molded line are respectively deviated by 0.06C_axWidth 0.12C_axThe area of (a).

In the first exhaust film hole and the second exhaust film hole, the hole pitch P/d between adjacent film holes is more than 2, and the inlet width of the pressure surface cooling area is 0.24C_axThe intersection point of the central line and the front edge point of the blade is the central hole position of the first exhaust film hole, a perpendicular line is made at the position, and the positive and negative error between the projection width of the first exhaust film hole on the perpendicular line and the width of the pressure surface inlet area is 0.04C_axThe hole row is parallel to the main flow inlet, and the width of the second exhaust membrane hole is 0.12C_axPositive and negative error 0.02C_axThe hole rows are perpendicular to the position of the local polar limiting line, and the number of the air film holes is obtained by downward rounding according to the diameters of the air film holes.

The row direction of the film holes is adjusted to be perpendicular to the local surface limit streamline of the end wall, so that the coverage area of the film is increased as much as possible, the cooling efficiency of the downstream of the outlet of the film holes is increased, and the arrangement of the film holes in the middle and the rear of the end wall is reduced.

The compound angle of the air film hole is 0 degree, the jet angle is 30 degrees, the length-diameter ratio L/d is 4.56, wherein L is the length of the central line of the air film hole.

The air sources of all the air film holes are uniform.

The characteristic distance between the first exhaust film hole and the leakage flow outlet is 0.15C_axThe center of the lower end hole is 0.21C away from the leading edge point of the stator blade_axThe arrangement is such that the row direction is parallel to the inlet boundary; the characteristic distance between the second exhaust membrane hole and the leakage flow outlet is 0.40C_axThe center of the lower end hole is 0.66C away from the leading edge point of the stator blade_axThe arrangement mode is that the included angle between the row direction and the inlet boundary is 44.9 degrees; the first exhaust film hole pitch P/d was 2.67 and the second exhaust film hole pitch P/d was 2.0.

Compared with the prior art, the invention realizes the improvement of the air film cooling efficiency of the whole end wall under the condition of leakage flow, simultaneously reduces the damage to the whole structure of the end wall to the maximum extent, enhances the structural strength of the end wall, improves the working temperature limit of the whole turbine, reduces the cooling air introduced from the air compressor, reduces the power consumption of the air compressor and improves the operation stability of the air compressor, and in addition, when in application, the air film hole compound angle can be flexibly adjusted according to the actual condition, and the whole air film cooling effect is optimized.

Drawings

FIG. 1 shows the principal locations of two vent holes in the arrangement of the present invention.

FIG. 2 shows the specific structural parameters of the gas film holes in the arrangement of the present invention.

Fig. 3 is a schematic diagram of the overall layout of the layout structure of the present invention.

FIG. 4 is a schematic view of the endwall limit streamlines in the inventive arrangement.

FIG. 5 illustrates the cooling effect of the original upstream leakage flow on the endwall.

FIG. 6 is a view showing the distribution of the gas film in the end wall of the arrangement according to the present invention.

Detailed Description

The embodiments of the present invention will be described in detail below with reference to the drawings and examples.

The invention relates to an end wall film hole arrangement structure which comprises the arrangement positions of a first row and a second row of cooling holes, an included angle between the first row and a main flow direction, hole spacing and other parameters. The arrangement mode of the first exhaust film holes is mainly based on the consideration of pressure surface side cooling, streamline is converged from the pressure surface side to the suction surface side due to the strong secondary flow influence on the turbine end wall, therefore, under the action of a limited cooling unit, the cooling efficiency of the end part area close to the pressure surface is difficult to improve, the cooling capacity of the first exhaust film holes is parallel to the main flow direction, after the secondary flow quantity is synthesized, the gas film covering attached to the side close to the pressure surface is formed, so the pressure surface area of the end wall is cooled, the upstream of the first exhaust film holes has reached enough cooling effect due to the action of upstream leakage flow, so the arrangement of the gas film holes is reduced, and the second exhaust film holes are arranged due to the main flow re-attaching action in the gas film clearance formed by the first exhaust film holes and the leakage flow, so the cooling gas film is formed by the arrangement of a small amount of the gas film holes, therefore, the gas film coverage rate of the whole end wall is highest, the using effect is not repeated, and the gas for cooling is saved to the maximum extent. This arrangement, which is based on the flow characteristics of the end wall itself, therefore has the effect of cooling the turbine end with such flow phenomena.

As shown in FIG. 1, the present invention is described in detail with respect to an aircraft engine first stage vane endwall, with an axial chord length C_ax75.31mm, which is characteristic for all the following positional parameters. The film holes are arranged in two rows, wherein the projection vector of the cold air amount of the first exhaust film hole on the end wall plane is parallel to the main flow direction, and covers the front edge point (0.21-0.53) C on the lower pressure surface side_axA region (2) forming a pressure surface side (0.5 to 1.0) C with respect to the end wall after being synthesized with the secondary flow amount_axThe air film covering of the position cools the pressure surface area of the end wall; the second exhaust film hole is arranged in an air film gap formed by the first exhaust film hole and the leakage flow, and the gaps are respectively deviated by 0.06C from the two sides of the central line of the upstream leakage flow close to the side boundary of the pressure surface and the molded line of the pressure surface_axWidth 0.12C_axSo as to form a cooling air film and achieve the highest air film coverage rate on the whole end wall.

The specific arrangement of the two rows of gas film holes can be seen from fig. 1, wherein the gas film hole at the lowest end is taken as the positioning hole, and the characteristic distance from the first gas film hole to the leakage flow outlet is 0.15C_axThe center of the lower end hole is 0.21C away from the leading edge point of the stator blade_axThe pitch P/d of the holes is 2.67, and the row direction is parallel to the inlet boundary; the characteristic distance between the second exhaust membrane hole and the leakage flow outlet is 0.40C_axThe center of the lower end hole is 0.66C away from the leading edge point of the stator blade_axThe hole pitch is P/d of 2.0, the included angle between the row direction and the inlet boundary is 44.9 degrees, the compound angle of the two rows of air film holes is 0 degree, wherein the compound angle of the air film holes refers to the included angle between the projection line of the central line of the air film holes on the end wall plane and the inlet main flow direction.

FIG. 2 shows the specific structural parameters of the film holes, the jet angles of the two rows of film holes are both 30 degrees, the thickness of the end wall is 2.28d, and the length-diameter ratio L/d is 4.56, wherein the jet angle is the included angle between the center line of the film hole and the plane of the end wall and is parallel to the cooling air momentum direction, and the length-diameter ratio represents the relative wall thickness of the end wall.

The design principle of the present invention needs to be explained with reference to fig. 3. Fig. 3 shows several cooling zones into which the end wall is divided according to the invention, in modern combustion turbine engines the leading edge of the end wall is usually given a large radius, and the angle of the leading edge film holes is adjusted according to the condition of the end wall, so that the leading edge horseshoe area already has sufficient cooling effect, and the upstream leakage flow forms a triangular film covering area in the presence of the secondary flow, which area also has sufficient cooling effect, and finally, the part to be cooled is mainly the area near the pressure surface and the strip-like area between the upstream leakage flow and the pressure surface.

Specifically, the cooling area of the end wall is divided into a leakage flow triangular cooling area 1, a secondary flow cooling area 2, a leading edge cooling area 3 and a pressure surface cooling area 4, wherein the leakage flow triangular cooling area 1 refers to an air film covering area formed by upstream leakage flow on the surface of the end wall, and is shown as a boundary with cooling efficiency higher than 0.3 on an adiabatic air film cooling efficiency cloud chart, and the cold air sprayed out from the leading edge of the cascade and the sealing of the wheel rim is partially migrated to the end wall, so that the cooling effect of the cold air is sufficient. The pressure surface cooling zone 4 means that the upstream leakage flow is deviated by 0.06C from the pressure surface side toward the pressure surface side boundary and the center line of the pressure surface molded line_axThe region between the pressure profile and the leading edge cooling zone 3 is the blade leading edge which is shifted outwards by 0.2C_axIn the leading edge cooling zone 3, film migration of the leading edge of the blade towards the endwall results in a film coverage effect. The secondary flow cooling area 2 refers to that the upstream leakage flow is close to the side boundary of the pressure surface and is respectively deviated by 0.06C from the two sides of the central line of the pressure surface molded line_axWidth 0.12C_axThe area of (a).

The invention carries out air film hole arrangement in a secondary flow cooling area 2 and a pressure surface cooling area 4, which are mainly arranged according to the area size and the flowing condition of boundary layer fluid, and a proper number of air film holes are uniformly arranged between the widths of the two areas by adjusting the hole spacing, and meanwhile, the included angle between the hole row and the flowing direction is mainly adjusted, and the included angle is basically ensured to be vertical to the flowing direction (the local surface limit streamline of the end wall), thereby increasing the coverage area of the air film as much as possible, increasing the cooling efficiency of the downstream of the air film hole outlet, and reducing the condition that the air film holes are arranged in the middle and the rear part of the end wall. In practical application, the air sources of all the air film holes are uniform.

Wherein, in order to avoid the mutual mixing of the air films between the two air film holes, the P/d is more than 2, and on the basis, the inlet width of the pressure surface cooling area 4 is 0.24C_axThe intersection point of the line and the front edge point of the blade is the central hole position of the first exhaust film hole, a perpendicular line is made at the position, the projection width of the first exhaust film hole on the perpendicular line is close to the width of the inlet area of the pressure surface, and the positive and negative error is 0.04C_axThe hole row is parallel to the main flow inlet, and the width of the second exhaust membrane hole is 0.12C_axPositive and negative error 0.02C_axThe hole row is vertical to the position of the local polar limiting line, and the number of the air film holes is obtained by downward integration according to the diameter of the air film holes after the relation is clear.

As can be seen from fig. 4, there is a strong secondary flow in the turbine end wall, which is caused by the horseshoe vortex generated by the leading edge and the pressure gradient in the boundary layer, under the action of the secondary flow, the fluid flows from the pressure surface side to the suction surface side of the end wall, so that the cooling film formed by the upstream leakage flow on the end wall is in a triangular distribution, as shown in fig. 5, and the covered area is sufficiently cooled, so that only the area where the leakage flow cannot be cooled needs to be arranged with the cooling unit, the first film hole is arranged parallel to the inlet, mainly aiming at the cooling effect of the end wall close to the pressure surface side, the resultant of the cooling air momentum and the secondary flow amount deflects the film, adapts to the pressure surface side profile, forms the covered cooling film, and due to the main flow re-attaching effect existing between the first film hole and the leakage flow, a region where the cooling cannot be formed in the middle, the second film discharge hole is thus oriented substantially perpendicular to the area, so that the film covers substantially the entire end wall surface.

As shown in FIG. 6, the whole scheme only uses two rows of 11 circular film holes, and achieves sufficient cooling effect, which shows that the arrangement scheme has great reference value, and the end wall of the first stage static blade of the turbine of the combustion engine is sufficiently cooled under the condition that upstream leakage flow exists, and meanwhile, cooling air for cooling is saved to the maximum extent, and the working temperature limit of the turbine is improved. The whole set of arrangement scheme simple structure has simplified process flow, has increased blade end wall structural strength simultaneously, can let out the nimble adjustment of the condition of leakage flow according to the upper reaches, can satisfy the cooling effect to different blade pressure face regions through the size of adjusting compound angle for holistic design can have more general application value, promotes the availability factor of polymorphic type turbine, satisfies industrial production's demand.

In conclusion, the invention realizes the improvement of the film cooling efficiency of the whole end wall under the condition of the existence of leakage flow, reduces the temperature of the whole end wall surface by introducing cooling air, strengthens the cooling efficiency of the root part of a pressure surface, protects blades from ablation, simultaneously reduces the damage to the boundary layer of an end wall flow field to the maximum extent, enhances the structural strength of the end wall, improves the working temperature limit of the whole turbine, reduces the cooling air introduced from a gas compressor, reduces the power consumption of the gas compressor and improves the operation stability of the gas compressor, and is suitable for the first-stage stationary blade end wall and the movable blade end wall of the gas turbine under the conditions of various blowing ratios and the existence of upstream leakage flow. In addition, when the air film cooling device is applied, the air film hole composite angle and the air film hole structure can be flexibly adjusted according to actual conditions, and the overall air film cooling effect is optimized.