CN114109518A - Turbine blade leading edge ribbed rotational flow-air film composite cooling structure - Google Patents
Turbine blade leading edge ribbed rotational flow-air film composite cooling structure Download PDFInfo
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- CN114109518A CN114109518A CN202111438860.XA CN202111438860A CN114109518A CN 114109518 A CN114109518 A CN 114109518A CN 202111438860 A CN202111438860 A CN 202111438860A CN 114109518 A CN114109518 A CN 114109518A
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- leading edge
- swirl
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
- F01D5/12—Blades
- F01D5/14—Form or construction
- F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
- F01D5/186—Film cooling
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Abstract
一种涡轮叶片前缘带肋旋流‑气膜复合冷却结构，包括设置在涡轮叶片内部的旋流腔，旋流腔为圆柱形，位于涡轮叶片前缘处，旋流腔内设置有若干螺旋肋片，若干射流孔在涡轮叶片压力面侧与旋流腔切向连通以实现冷气的切向入射，若干前缘气膜孔在涡轮叶片前缘与旋流腔连通，若干叶顶气膜孔在叶顶与旋流腔连通，冷气由射流孔进入旋流腔形成切向旋流，增强附近壁面热交换。螺旋肋片增大周向旋流强度，并增大旋流腔前缘壁面换热面积，部分冷气由前缘气膜孔流出并覆盖在叶片前缘壁面，形成气膜冷却，同时改善腔内流动情况。其余冷气经叶顶气膜孔流出，在叶顶处形成气膜冷却。本发明能改善涡轮叶片前缘换热均匀性，提高叶片前缘换热性能，降低叶片前缘温度。
A ribbed swirl-air-film composite cooling structure on the leading edge of a turbine blade, comprising a swirl cavity arranged inside the turbine blade, the swirl cavity is cylindrical and located at the leading edge of the turbine blade, and a plurality of spirals are arranged in the swirl cavity Ribs, several jet holes are tangentially connected to the swirl cavity on the pressure surface side of the turbine blade to achieve tangential incidence of cold air, several leading edge air film holes are connected to the swirl cavity at the leading edge of the turbine blade, and several tip air film holes At the tip of the blade, it communicates with the swirl cavity, and the cold air enters the swirl cavity through the jet hole to form a tangential swirl, which enhances the heat exchange on the nearby wall. The spiral fins increase the circumferential swirl intensity and increase the heat exchange area of the leading edge wall of the swirl cavity. Part of the cold air flows out from the leading edge air film hole and covers the leading edge wall of the blade, forming an air film cooling and improving the flow in the cavity at the same time. . The rest of the cold air flows out through the air film holes at the tip of the blade and forms an air film cooling at the blade tip. The invention can improve the heat exchange uniformity of the leading edge of the turbine blade, improve the heat exchange performance of the leading edge of the blade, and reduce the temperature of the leading edge of the blade.
Description
Technical Field
The invention belongs to the technical field of cooling and heat exchange of turbine blades, relates to cooling and heat exchange of high-temperature turbine blades of an aero-engine, and particularly relates to a turbine blade leading edge ribbed rotational flow-air film composite cooling structure.
Background
The specific power of a gas turbine is a key measure of its thermal performance, and the thermal efficiency is directly related to the inlet temperature of the turbine. An increase in turbine inlet temperature can result in increased thermal loading of the hot end components (vanes, etc.) and even failure beyond the material tolerance. The leading edge of the blade is one of the areas with the highest thermal load of the blade and is one of the most easily ablated parts. Thus, there is a need for more efficient and reliable blade cooling techniques that allow safe and long lasting operation of the blades.
The rotational flow cooling technique has high heat transfer strength and small flow resistance, and is an important research subject for cooling the inside of the front edge of the blade. The rotational flow cooling enables cooling air flow to be injected into the rotational flow chamber tangentially by changing the air inlet position of the jet hole, and transverse large-scale vortex is formed along the arc target surface, so that the turbulent kinetic energy of the fluid is enhanced, and the internal heat exchange of the blade is enhanced. The air film cooling makes cooling air flow cover the surface of the blade by arranging discrete air on the surface of the blade, so that high-temperature gas is isolated, and the blade is protected.
However, the current research on the swirl cooling structure still remains in the aspects of simply changing geometric structure parameters and the like, and no further work is carried out on the directions of improving the internal swirl, reducing the starting loss and the like.
Disclosure of Invention
In order to overcome the defects of the prior art and solve the problems of the prior rotational flow cooling in the aspects of heat exchange strength, heat exchange uniformity, aerodynamic performance and the like, the invention aims to provide a turbine blade leading edge ribbed rotational flow-air film composite cooling structure.
In order to achieve the purpose, the invention adopts the technical scheme that:
the utility model provides a turbine blade leading edge ribbed whirl-air film composite cooling structure, is including setting up the whirl chamber in turbine blade inside, and the whirl chamber is cylindrical, is located turbine blade leading edge department, be provided with a plurality of spiral fins in the whirl chamber, a plurality of jet holes communicate with whirl chamber tangential in order to realize the tangential incidence of air conditioning at turbine blade pressure face side, and a plurality of leading edge air film holes communicate with the whirl chamber at turbine blade leading edge, and a plurality of blade top air film holes communicate with the whirl chamber at the blade top, and the blade top of accomplishing air conditioning is flowed out.
In one embodiment, the jet holes are uniformly distributed along the axial direction of the swirl cavity (namely the radial direction of the turbine blade), the jet holes are controlled independently, the hole sections of the jet holes are rectangular, the ratio P/D of the hole pitch P of the adjacent jet holes to the diameter D of the swirl cavity is 1.5-6, the ratio H/D of the section height H to D of the jet holes is 0.6-1, the ratio R/D of the hole length R to D of the jet holes is 0.3-0.8, and the ratio delta/D of the width delta to D of the jet holes is 0.02-0.1.
In one embodiment, the plurality of helical fins are disposed at the leading edge of the turbine blade at an angle θ of 90-180 °.
In one embodiment, the plurality of spiral fins are divided into a plurality of groups, and each group of spiral fins is arranged between adjacent jet holes along the axial direction of the swirl chamber.
In one embodiment, the ratio phi/D of the pitch phi of each adjacent group of spiral fins to the diameter D of the swirl cavity is 0.5-1, the ratio h/D of the height h of the fins to the diameter D of the swirl cavity is 0.05-0.2, the ratio w/D of the thickness w of the fins to the diameter D of the swirl cavity is 0.05-0.2, the number of the fins of each group of spiral fins is single rib, double rib or three ribs, and when the number of the fins of each group of spiral fins is double rib or three rib, the ratio delta/D of the distance delta of the fins in the group to the diameter D of the swirl cavity is 0.5-2.
In one embodiment, the plurality of leading edge air film holes are divided into a plurality of groups, and each group of leading edge air film holes are arranged between adjacent jet holes along the axial direction of the swirl cavity and are communicated with the downstream position of the swirl fins.
In one embodiment, each group of leading edge air film holes between the adjacent jet holes is divided into 1-3 subgroups along the axial direction of the vortex chamber, each subgroup is composed of three air film holes, wherein the horizontal direction projection direction of the middle air film hole is opposite to the incoming flow direction of the blade, and the other two air film holes are symmetrically arranged on two sides of the middle air film hole and are respectively biased to the pressure surface side and the suction surface side wall surface, so that the outgoing flow flows to the pressure surface side and the suction surface side respectively.
In one embodiment, theThe leading edge air film holes are cylindrical holes, the hole diameter d is 0.3mm-1mm, the included angle formed by the center line of the air film holes and the cross section perpendicular to the turbine blade is defined as an air film hole composite angle alpha, the composite angle of each leading edge air film hole is alpha, and the alpha is more than 0 degree and less than or equal to 60 degrees; the other two air film holes are positioned at the same axial position of the rotational flow cavity, and the interval length between the other two air film holes and the middle air film hole is djA included angle beta with the horizontal direction of the middle air film hole (namely the direction vertical to the cross section of the turbine blade), wherein the included angle beta is more than or equal to 0 degree and less than or equal to 40 degrees, and d isjThe length of the long axis of the intersected elliptic sideline of the front edge air film hole and the rotational flow cavity.
In one embodiment, the tip film holes are cylindrical holes, are 1 in number, and are coaxial with the swirl chamber.
In one embodiment, the diameter of the gas film hole at the tip of the blade
Compared with the prior art, the invention has the beneficial effects that:
compared with the traditional rotational flow cooling structure, the cooling jet flow forms large-scale circumferential rotational flow in the rotational flow cavity through the uniformly arranged tangential jet holes, and the circumferential rotational flow strength is enhanced and the heat exchange area is increased through the circumferential rotation promoting effect generated by the spiral fins arranged among the jet holes. Meanwhile, the plurality of front edge air film holes arranged at the downstream of the spiral fins are used for leading part of cooling jet flow to flow out from the front edge air film holes under the suction effect generated by the internal and external pressure of the rotational flow cavity, so that the heat exchange capability of the wall surface near the front edge air film holes is enhanced. In addition, the flow state of the surrounding swirling flow is improved due to the reduction of the mass flow of the cold air near the front edge air film hole, and the temperature distribution uniformity of the front edge wall surface of the blade is further improved. In addition, partial cold air is discharged from the air film hole at the top of the blade to form blade top air film cooling, so that the cooling effect near the front edge of the blade top is improved, and the local thermal uniformity is improved.
Therefore, the invention can improve the cooling effect of the front edge wall surface under the same cooling jet flow, improve the heat transfer uniformity of the front edge target surface and reduce the thermal stress of the front edge of the blade.
Drawings
FIG. 1 is a schematic layout of a leading edge swirl-film composite cooling structure in a blade. In the figure, M denotes the suction surface side and N denotes the pressure surface side.
Fig. 2 is a partial sectional view taken along line a-a in fig. 1.
FIG. 3 is a front view of the fluid domain of the swirl-film composite cooling structure.
FIG. 4 is a top view of a fluid domain of a swirl-film composite cooling structure.
FIG. 5 is a left side view of the fluid domain of the swirl-film composite cooling structure. And J in the figure is the tail edge side of the rotational flow cooling cavity, and K is the front edge side of the rotational flow cooling cavity.
Fig. 6 is an enlarged partial view of the front view of fig. 3 at the air film hole, i.e., in the area B.
FIG. 7 is a schematic view of a series of leading edge film holes disposed downstream of two spiral fins.
FIG. 8 is a schematic view of a series of leading edge film holes disposed between two spiral fins.
FIG. 9 is a schematic structural diagram of two groups of leading edge film holes respectively arranged between and at the downstream of two spiral fins.
FIG. 10 is a schematic view of the structure of the spiral fins, where the pitch of a complete turn of the spiral fins is phi.
FIG. 11 is a schematic structural view of helical fins arranged in a helical swirl chamber, wherein the arrangement angle of the helical fins in the front edge of the swirl chamber is theta.
Detailed Description
The embodiments of the present invention will be described in detail below with reference to the drawings and examples.
As shown in FIGS. 1, 2, 3, 4, 5 and 6, the invention provides a turbine blade leading edge ribbed swirl-film composite cooling structure, which comprises a swirl chamber 2 arranged inside a turbine blade, wherein the swirl chamber 2 is cylindrical and is positioned at the leading edge of the turbine blade. On the pressure surface side of the turbine blade, a plurality of jet holes 1 are tangentially communicated with a rotational flow cavity 2 so as to realize tangential incidence of cold air; at the front edge of the turbine blade, a plurality of front edge air film holes 4 are communicated with the vortex cavity 2; at the blade top of the turbine blade, a plurality of blade top air film holes 5 are communicated with the rotational flow cavity 2; inside the cyclone chamber 2, a plurality of spiral fins 3 are also arranged.
Cold air is jetted into the cyclone cavity 2 from the jet hole 1 in a tangential direction, a strong circumferential cyclone is formed in the cyclone cavity 2, a transverse large-scale vortex is formed under the action of the wall surface of the arc-shaped cavity, heat exchange of the wall surface near the vortex is enhanced, and strong convective heat transfer is carried out. Along with the radial flow of the air flow, the circumferential rotational flow strength is gradually reduced, at the moment, the circumferential rotational flow strength is increased through the circumferential flow guiding effect of the spiral fins 3, the rotation promotion is formed, and the convection heat exchange strength nearby is enhanced. On the other hand, the spiral fins 3 can also increase the heat exchange area of the front edge wall surface of the vortex cavity 2, so that the heat exchange is enhanced. And then part of cold air flows out through the front edge air film holes 4 and covers the front edge wall surface of the blade to isolate the wall surface from main flow gas, so that air film cooling is formed, meanwhile, cross flow in the blade is reduced, the internal rotational flow flowing state is improved, the internal heat exchange uniformity is improved, and the rest part of cold air flows out from the blade top air film holes 5 and forms air film cooling at the blade top.
In one embodiment, the jet holes 1 are uniformly distributed at equal intervals along the axial direction of the swirl chamber 2 (i.e. the radial direction of the turbine blade), and the jet holes 1 are independently controlled so as to control the flow rate of each jet hole 1 to meet the required requirement. Illustratively, the hole cross section of the jet holes 1 is rectangular, along the axial direction of the swirl cavity 2, the ratio P/D of the hole pitch P of the adjacent jet holes 1 to the diameter D of the swirl cavity 2 is 1.5-6, the ratio H/D of the section height H to D of the jet holes 1 is 0.6-1, the ratio R/D of the hole length R to D of the jet holes 1 is 0.3-0.8, and the ratio delta/D of the width delta to D of the jet holes 1 is 0.02-0.1. The length L of the swirl chamber 2 and the row spacing epsilon of the jet holes 1 are also shown.
In one embodiment, several helical fins 3 are arranged at the position of the leading edge of the turbine blade, the leading edge arrangement angle θ being 90-180 °. For example, the spiral fins 3 may be divided into a plurality of groups, and each group of the spiral fins 3 may be arranged between the adjacent jet holes 1 in the axial direction of the swirling chamber 2.
Referring to fig. 7, 8, 9, 10, 11, in one embodiment, the number of ribs of each set of spiral ribs 3 is single rib, double rib, or triple rib, and the ribs are uniformly arranged between adjacent jet holes 1 according to the number of ribs. Example (c): for the single-rib structure, the ribs are arranged in the middle of the adjacent jet holes 1; for the double-rib structure, double ribs are arranged between adjacent jet holes 1 at equal intervals; so as to be regularly arranged. The ratio phi/D of the pitch phi of each group of adjacent spiral fins 3 to the diameter D of the swirling flow cavity 2 is 0.5-1, the pitch of the spiral fins 3 arranged between the two jet holes 1 can be different, and the arrangement radial position of the spiral fins changes from the middle of the two jet holes 1 to the vicinity of the upstream of the jet holes 1. The ratio h/D of the height h to D of the ribs is 0.05-0.2, and the ratio w/D of the thickness w to D of the ribs is 0.05-0.2. When the double ribs or the triple ribs are adopted, the ratio delta/D of the distance delta of the ribs in the group to the distance D is 0.5-2.
In one embodiment, the plurality of leading edge film holes 4 can be divided into a plurality of groups, and each group of leading edge film holes 4 is arranged between the adjacent jet holes 1 along the axial direction of the swirling chamber 2 and is communicated with the downstream position of the swirling rib 3.
Along the axial direction of the rotational flow cavity 2, each group of front edge air film holes 4 between the adjacent jet holes 1 is divided into 1-3 groups, each group consists of three air film holes, wherein the horizontal direction projection direction of the middle air film hole is just opposite to the incoming flow stagnation line direction of the blade, and the other two air film holes are symmetrically arranged at two sides of the middle air film hole and are respectively deviated to the side wall surfaces of the pressure surface and the suction surface, so that the outflow flows to the side of the pressure surface and the side of the suction surface respectively.
Illustratively, the leading edge film holes 4 are cylindrical holes, the hole diameter d is 0.3mm-1mm, an included angle formed by the center line of the film hole and the cross section perpendicular to the turbine blade is defined as a film hole composite angle alpha, the composite angle of each leading edge film hole 4 is alpha, and alpha is more than 0 degree and less than or equal to 60 degrees. Two air film holes arranged at two sides of the middle air film hole are positioned at the same axial position of the swirling flow cavity 2, and the interval length between the two air film holes and the middle air film hole is djA included angle beta with the horizontal direction of the middle air film hole (namely the direction vertical to the cross section of the turbine blade), wherein the included angle beta is more than or equal to 0 degree and less than or equal to 40 degrees, and d isjThe length of the long axis of the intersected elliptic sideline of the front edge air film hole 4 and the swirling cavity 2.
In one embodiment, the gas film holes 5 of the blade top are cylindrical holes, the number of the holes is 1, the holes are coaxial with the swirling cavity 2, and the hole diameter of the holes
The hole length b is also shown.
The specific embodiments of the present invention are merely exemplary and should not be construed as limiting the scope of the invention, which is intended to include other equivalents within the scope of the invention.
Claims (10)
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| CN202111438860.XA CN114109518A (en) | 2021-11-29 | 2021-11-29 | Turbine blade leading edge ribbed rotational flow-air film composite cooling structure |
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| CN202111438860.XA CN114109518A (en) | 2021-11-29 | 2021-11-29 | Turbine blade leading edge ribbed rotational flow-air film composite cooling structure |
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Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN115013076A (en) * | 2022-08-10 | 2022-09-06 | 中国航发四川燃气涡轮研究院 | Gondola water faucet form turbine blade cooling unit and turbine blade |
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| CN115013076A (en) * | 2022-08-10 | 2022-09-06 | 中国航发四川燃气涡轮研究院 | Gondola water faucet form turbine blade cooling unit and turbine blade |
| CN115013076B (en) * | 2022-08-10 | 2022-10-25 | 中国航发四川燃气涡轮研究院 | Gondola water faucet form turbine blade cooling unit and turbine blade |
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