CN115701846A - Air film cooling structure for gas turbine blade - Google Patents
Air film cooling structure for gas turbine blade Download PDFInfo
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- CN115701846A CN115701846A CN202110834124.XA CN202110834124A CN115701846A CN 115701846 A CN115701846 A CN 115701846A CN 202110834124 A CN202110834124 A CN 202110834124A CN 115701846 A CN115701846 A CN 115701846A
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Abstract
The invention provides a film cooling structure for a combustion engine blade, comprising: the front edge of the blade top of the gas turbine blade is provided with a plurality of first air film holes in a linear array along the height of the blade; a plurality of air film air outlet areas are arranged on blade top grooves of the gas turbine blades along the middle chord lines; a plurality of third air film holes in a linear array are formed in the full chord length of the outer side of the blade top pressure surface of the combustion engine blade; the axial line of each third air film hole and the tangential direction of the pressure surface of the blade top form an acute angle, and the distance between every two adjacent third air film holes is gradually increased from the front edge to the tail edge; a plurality of linear array fourth air film holes are formed in the outer side of the blade top suction surface of the combustion engine blade from the front edge to the middle chord section; the invention improves the cooling design of the blade top area, and the four areas of the blade top are provided with the air film holes, so that low-temperature cold air immediately covers the surface of the blade top after flowing out of the air film holes, effective cold air protection is formed, and the purpose of improving the overtemperature of the blade top is achieved.
Description
Technical Field
The invention relates to the technical field of cooling of gas turbine blades, in particular to the technical field of an air film cooling structure for gas turbine blades.
Background
With the increasing requirements on efficiency and power of modern combustion engines, the inlet airflow temperature of the turbine blade is higher and higher; the gas turbine blade 1 can be divided into a blade root a, a blade b and a blade top c, and is divided into a leading edge d and a trailing edge e according to the gas feeding direction; during maintenance, the blade top c has a high-temperature phenomenon, and the thermal barrier coating and the base material in the area can be damaged; therefore, the film cooling technology becomes more important as an effective thermal protection measure, the film cooling is sprayed out from the film holes on the wall surface to prevent the gas flow from heating the wall surface, and the film cooling mainly plays two roles, one is to take away a part of heat through the cooling gas flow, and the other is to prevent the wall surface and the gas flow from being blocked through the cooling gas to protect the wall surface.
For example, in patent document CN107143383A, a plurality of air film holes with equal spacing are generally arranged at the blade tip groove of a turbine blade, which easily causes instability of the whole blade structure, and the multiple times of hole opening easily increases the complexity and cost of the process; in addition, the film holes in this document are typically also provided at the pressure surface of the blade, and do not allow for effective cooling of the suction surface of the blade tip.
Disclosure of Invention
In view of the above-mentioned drawbacks of the prior art, an object of the present invention is to provide a film cooling structure for a combustion engine blade, which solves the problems of the prior art in simplifying the film cooling structure while maintaining better cooling performance.
To achieve the above and other related objects, the present invention provides a film cooling structure for a blade of a combustion engine, comprising: .
The gas turbine blade is characterized in that a plurality of first gas film holes in a linear array are formed in the front edge of the blade top of the gas turbine blade along the height of the blade;
four air film air outlet areas are arranged on the blade top groove of the gas turbine blade along a middle chord line; each air film air outlet region is provided with a plurality of second air film holes in a linear array along a middle chord line;
a plurality of third air film holes in a linear array are formed in the full chord length of the outer side of the blade top pressure surface of the combustion engine blade; the axial line of each third air film hole and the tangential direction of the blade top pressure surface form an acute angle, and the distance between every two adjacent third air film holes is gradually increased from the front edge to the tail edge;
a plurality of fourth air film holes in a linear array are formed in the outer side of the suction surface of the blade top of the combustion engine blade from the front edge to the middle chord section;
the first, second, third and fourth film holes are all communicated with a cooling airflow cavity of the combustion engine blade.
Preferably: the fourth air film holes are formed from the front edge to the front of the throat of the suction surface of the blade.
Preferably: the diameters of the first air film hole, the third air film hole and the fourth air film hole are d1, and d1 is more than or equal to 0.5mm and less than or equal to 1.5mm; and the distance between every two adjacent first air film holes or the third air film holes or the fourth air film holes is 5-8 d1, and the normal included angle is 30-60 degrees.
Preferably: the aperture of the second air film hole is d2, and d2 is more than or equal to 1mm and less than or equal to 4mm.
Preferably: a first of said film exit regions from said tip leading edge, said film exit region being positioned at a mid-chord length of between 10% and 20%; the included angle between the axis of the second air film hole of the air film air outlet area and the surface of the blade top groove is theta 1, and theta 1 is more than or equal to 10 degrees and less than or equal to 40 degrees; the included angle between the axis of the second air film hole in the area and the tangent direction of the mean camber line is delta 1, and delta 1 is more than or equal to 0 degree and less than or equal to 40 degrees.
Preferably: a second said film exit region from said tip leading edge, said film exit region being positioned at a mid-chord length of between 35% and 40%; the included angle between the axis of the second air film hole of the air film air outlet area and the surface of the blade top groove is theta 2, theta 2 is more than or equal to 20 degrees and less than or equal to 50 degrees, and the included angle between the axis of the second air film hole in the area and the tangential direction of a mean camber line is delta 2, delta 2 is more than or equal to-10 degrees and less than or equal to 30 degrees.
Preferably: a third said film exit region from said leading edge of said blade tip, said film exit region being positioned at a mid-chord length of between 50% and 55%; the included angle between the axis of the second air film hole of the air film air outlet area and the surface of the blade top groove is theta 3, theta 3 is more than or equal to 20 degrees and less than or equal to 50 degrees, and the included angle between the axis of the second air film hole in the area and the tangential direction of a mean camber line is delta 3, delta 3 is more than or equal to-30 degrees and less than or equal to-10 degrees.
Preferably: a fourth air film air outlet region from the leading edge of the blade tip, the air film air outlet region being positioned at a mid-chord length of 85% to 90%; the included angle between the axis of the second air film hole of the air film air outlet area and the surface of the blade top groove is theta 4, wherein theta 4 is more than or equal to 60 degrees and less than or equal to 90 degrees; an included angle between the axis of the second air film hole in the area and the tangent direction of the mean camber line is delta 4, wherein the included angle is larger than or equal to minus 30 degrees and is larger than or equal to delta 4 and is smaller than or equal to 0 degree.
As mentioned above, the film cooling structure for the top of the high-pressure turbine power blade of the invention has the following beneficial effects:
the invention improves the cooling design of the blade top area, and the four areas of the blade top are provided with the air film holes, so that low-temperature cold air immediately covers the surface of the blade top after flowing out of the air film holes, effective cold air protection is formed, and the purpose of improving the over-temperature of the blade top is achieved.
Drawings
FIG. 1 shows a perspective view of a combustion engine blade of an air film cooling structure for a combustion engine blade according to the present invention;
FIG. 2 shows a combustion engine blade tip of an air film cooling structure for a combustion engine blade according to the present invention;
FIG. 3 is a layout view of a plurality of second film holes of a film cooling structure for a combustion engine blade according to the present invention;
FIG. 4 is a schematic view of the included angles of a plurality of second film holes and tip grooves of a film cooling structure for a gas turbine blade according to the present invention;
FIG. 5 is a schematic view of the included angles of the plurality of second film holes and the camber line of the film cooling structure for a gas turbine blade according to the present invention;
FIG. 6 is a schematic view of a first plurality of film holes and a third plurality of film holes of a film cooling structure for a combustion engine blade according to the present disclosure;
FIG. 7 is a schematic view of a fourth plurality of film holes of a film cooling arrangement for a combustion engine blade according to the present invention;
FIG. 8 is a graph showing a temperature comparison of a film cooling structure for a gas turbine blade according to the present invention and a prior art;
FIG. 9 shows a schematic view of two opposite combustor vane throats of an air film cooling structure for a combustor vane of the present invention;
FIG. 10 is a temperature difference diagram of a gas turbine blade without a film hole of a film cooling structure for a gas turbine blade according to the present invention;
FIG. 11 is a temperature difference diagram illustrating a film cooling structure for a blade of a combustion engine according to the present invention.
Description of the element reference
Prior Art
a blade root
b blade
c leaf top
d leading edge
e trailing edge
This application is a
1. Combustion engine blade
111. First air film hole
12. Blade tip groove
120. Air film air outlet area
1201 Region A
1202 Region B
1203 C region
1204 D region
121. Second air film hole
1211 Second air film hole of A area
1212 Second air film hole of B area
1213 Second gas film hole of C area
1214 Second air film hole of D area
13. Pressure surface of blade top
131. Third air film hole
14. Suction surface of blade top
141. Fourth air film hole
Detailed Description
The following description of the embodiments of the present invention is provided for illustrative purposes, and other advantages and effects of the present invention will become apparent to those skilled in the art from the present disclosure.
Please refer to fig. 1 to 9. It should be understood that the structures, ratios, sizes, and the like shown in the drawings are only used for matching the disclosure of the present disclosure, and are not used for limiting the conditions that the present disclosure can be implemented, so that the present disclosure is not limited to the technical essence, and any structural modifications, ratio changes, or size adjustments should still fall within the scope of the present disclosure without affecting the efficacy and the achievable purpose of the present disclosure. In addition, the terms "upper", "lower", "left", "right", "middle" and "one" used in the present specification are for clarity of description, and are not intended to limit the scope of the present invention, and the relative relationship between the terms and the terms is not to be construed as a scope of the present invention.
As shown in fig. 1, the present invention provides a film cooling structure for a blade of a combustion engine, comprising:
the gas turbine blade comprises a gas turbine blade 1, wherein a plurality of first air film holes 111 in a linear array are formed in the front edge of the blade top of the gas turbine blade 1 along the height of the blade;
four air film air outlet areas 120 are arranged along a middle chord line in the blade top groove 12 of the combustion engine blade 1; each air film air outlet region 121 is provided with a plurality of second air film holes 121 in a linear array along a middle chord line; the four gas film gas outlet areas 120 are an area A1201, an area B1202, an area C1203 and an area D1204 from the front edge to the tail edge respectively; the second film holes 121 arranged in the area a 1201 face the blade top suction surface 14, so that the cold air blown out from the area a can cool the area 1201 and the blade top suction surface 14;
a plurality of third air film holes 131 in a linear array are arranged on the outer side of the blade top pressure surface 13 of the combustion engine blade 1 in the full chord length;
a plurality of fourth film holes 141 in a linear array are formed in the outer side of the blade top suction surface 14 of the combustion engine blade 1 from the front edge to the middle chord section;
the first, second, third, and fourth film holes 111, 121, 131, and 141 are all in communication with the cooling airflow chamber of the combustion engine blade 1.
According to the invention, the air film holes (comprising the first air film hole 111, the second air film hole 121, the third air film hole 131 and the fourth air film hole 141) are arranged in four areas (the blade top leading edge, the blade top groove 12, the blade top pressure surface 13 and the blade top suction surface 14) of the blade top for cooling, and compared with a blade top design without air film holes (see figure 10) through a CFD numerical method, see figure 11, the design of the invention can reduce the average temperature of the blade top by about 70K, and the cold air increment percentage is only 0.8%; the design of the blade top gas film can effectively reduce the heat load of the blade top area under the condition of avoiding obviously increasing the cold air quantity, and plays a role in cooling and protecting the blade top; in addition, region a 1201 of the present invention, in which tip cavity 12 is provided, may spray cool air toward tip suction surface 14, thereby further cooling tip suction surface 14.
Since the portion near the leading edge of the tip is more susceptible to gas erosion, a linear array of a plurality of first film holes 111 is provided along the height of the blade at the leading edge of the tip to protect the leading edge of the combustion engine blade 1 from gas erosion.
Because the cold air has a superposition effect, the cold air superposition effect can be better played by distributing the hole spacing from dense to sparse, so that the high temperature of the pressure surface 13 and the trailing edge of the blade top is improved; therefore, a linear array of the third film holes 131 is provided at the full chord length outside the tip pressure surface 13, and the distance between every two adjacent third film holes 131 gradually increases from the leading edge to the trailing edge.
In order to make the third film hole 131 cooling jet flow follow the direction of the external main flow, the cooling jet flow is attached to the downstream wall surface; the third film holes 131 are arranged such that the center line direction thereof is at an acute angle to the tangential direction of the tip pressure surface 13.
For the rear half section of the blade top suction surface 14, due to the phenomenon of blade top airflow crossing, cooling protection can be performed to a certain extent without arranging an air film hole; and in order to avoid flow separation and complicated flow pattern at the rear side of the blade top suction surface 14 of the combustion engine blade, where the blade top suction surface 14 is located at the throat (see fig. 9, the combustion engine blade cascade channel, which has the smallest cross section, and one side of the throat is located at the trailing edge of a combustion engine blade, and the other side is located at the blade top suction surface 14 opposite to the combustion engine blade), a film hole is now provided in front of the throat on the outer side of the blade top suction surface 14, and specifically, a plurality of fourth film holes 141 are provided in front of the front edge to the throat of the blade suction surface 14.
In the embodiment, the diameters of the first air film hole 111, the third air film hole 131 and the fourth air film hole 141 are d1, and d1 is greater than or equal to 0.5mm and less than or equal to 1.5mm; and the distance between every two adjacent first air film holes 111, or third air film holes 131 or fourth air film holes 141 is 5-8 d1, and the normal included angle is 30-60 degrees.
In this embodiment, the aperture of the second gas film hole 121 is d2, and the range of d2 is 1mm or more and 4mm or less; the effect of increasing the cold air amount is realized by increasing the aperture of the d2, so that the cooling amplitude of the blade top can be increased by blown cold air. In addition, the axial direction of the second air film hole 121 forms an acute angle with the groove surface of the blade top groove 12 and forms an acute angle with the tangent direction of the mean camber line; in addition, the arrangement of the second film holes 121 with the best effect is obtained through a CFD simulation mode, and when the number of the second film holes 121 is five, the following settings are specifically set:
in order to be able to achieve different effects for the individual holes of the tip pocket 12; specifically, the four gas film outlet regions 120 are respectively an area a 1201, an area B1202, an area C1203 and an area D1204 from the leading edge to the trailing edge; region A enables focused cooling of 30-40% chord length of tip suction surface 14 and forward portion of tip cavity 12, while second film holes 121 of regions B and C enable cooling of aft portion of tip cavity 12, and second film holes 1214 of region D1204 enable cooling of aft portion of tip cavity 12. Specifically, the method comprises the following steps:
first film exit region 120-A from leading edge of blade tip region 1201:the A area 1201 is positioned at 10% -20% of the length of the middle chord line; the included angle between the axis of the second air film hole 1211 of the area A1201 and the surface of the leaf top groove 12 is theta 1, and the theta 1 is more than or equal to 10 degrees and less than or equal to 40 degrees; the included angle between the axis of the second air film hole 1211 of the area A1201 and the tangent direction of the mean camber line is delta 1, and delta 1 is more than or equal to 0 degrees and less than or equal to 40 degrees. By setting δ 1 to a positive number, the cold air ejected from the a-region 1201 can be directed toward the tip suction surface 14.
Second gas film exit region 120-B from the leading edge of the blade tip region 1202:the B region 1202 is located at 35% to 40% of the mid-chord length; the included angle between the axis of the second air film hole 1212 of the B area 1202 and the plane of the blade tip groove 12 is theta 2, 20 degrees is larger than or equal to theta 2 and is smaller than or equal to 50 degrees, and the included angle between the axis of the second air film hole 1212 of the B area 1202 and the tangent direction of the mean camber line is delta 2, 10 degrees is larger than or equal to delta 2 and is smaller than or equal to 30 degrees.
Third gas film out-gassing zone 120-C from the leading edge of the leaf tip, zone 1203:the position of the C area 1203 is 50% -55% of the length of the middle chord line; the included angle between the axis of the second air film hole 1213 of the C region 1203 and the plane of the tip groove 12 is theta 3, 20-50 deg. theta 3, and the included angle between the axis of the second air film hole 1213 of the C region 1203 and the tangent direction of the mean camber line is delta 3, -30 deg. -delta 3-10 deg..
Fourth gas film exit region 120-D region 1204 from the leading edge of the blade tip:the D region 1204 is positioned at 85% -90% of the middle chord length; the included angle between the axis of the second air film hole 1214 of the D area 1204 and the surface of the blade top groove 12 is theta 4, and the theta 4 is more than or equal to 60 degrees and less than or equal to 90 degrees; the included angle between the axis of the second air film hole 1214 in the D area 1204 and the tangent direction of the mean camber line is delta 4, delta 4 is more than or equal to minus 30 degrees and less than or equal to 0 degree.
In summary, in the cooling design of the blade tip region, the air film holes (including the first air film hole 111, the second air film hole 121, the third air film hole 131 and the fourth air film hole 141) are arranged in the four regions (the blade tip leading edge, the blade tip groove 12, the blade tip pressure surface 13 and the blade tip suction surface 14) of the blade tip, so that low-temperature cold air immediately covers the surface of the blade tip after flowing out of the air film holes, effective cold air protection is formed, and the purpose of improving the blade tip overtemperature is achieved; in addition, compared with the existing arrangement without the air film holes, the three-dimensional temperature distribution of the blade top provided with the air film holes is only 0.8 percent of the cold air increment percentage. The effect of reducing the average temperature by about 70K is realized, the heat load of the top area of the blade is effectively reduced, and the effect of cooling and protecting the blade top is realized.
Therefore, the invention effectively overcomes various defects in the prior art and has high industrial utilization value.
The foregoing embodiments are merely illustrative of the principles and utilities of the present invention and are not intended to limit the invention. Any person skilled in the art can modify or change the above-mentioned embodiments without departing from the spirit and scope of the present invention. Accordingly, it is intended that all equivalent modifications or changes which may be made by those skilled in the art without departing from the spirit and scope of the present invention as defined in the appended claims.
Claims (8)
1. A film cooling structure for a combustion engine blade, comprising: .
The gas turbine blade is characterized in that a plurality of first gas film holes in a linear array are formed in the front edge of the blade top of the gas turbine blade along the height of the blade;
four air film air outlet areas are arranged on the blade top groove of the gas turbine blade along a middle chord line; each air film air outlet region is provided with a plurality of second air film holes in a linear array along a middle chord line; the second air film hole arranged in the first air film air outlet area from the front edge faces to a blade top suction surface;
a plurality of third film holes which are full in chord length and are in linear array are arranged on the outer side of the blade top pressure surface of the gas turbine blade; the axial line of each third film hole and the tangential direction of the blade top pressure surface form an acute angle, and the distance between every two adjacent third film holes is gradually increased from the front edge to the tail edge;
a plurality of fourth film holes in a linear array are formed in the outer side of the blade top suction surface of the combustion engine blade from the front edge to the middle chord section;
the first film hole, the second film hole, the third film hole and the fourth film hole are all communicated with a cooling airflow cavity of the combustion engine blade.
2. The film cooling structure for a blade of a combustion engine according to claim 1, wherein: the fourth air film holes are formed from the front edge to the front of the throat of the suction surface of the blade.
3. The film cooling structure for a blade of a combustion engine according to claim 1, wherein: the diameters of the first air film hole, the third air film hole and the fourth air film hole are d1, and d1 is more than or equal to 0.5mm and less than or equal to 1.5mm; and the distance between every two adjacent first air film holes or the third air film holes or the fourth air film holes is 5-8 d1, and the normal included angle is 30-60 degrees.
4. The film cooling structure for a blade of a combustion engine according to claim 1, wherein: the aperture of the second air film hole is d2, and d2 is more than or equal to 1mm and less than or equal to 4mm.
5. The film cooling structure for a blade of a combustion engine according to claim 1, wherein: a first of said film exit regions from said tip leading edge, said film exit region being positioned at a mid-chord length of between 10% and 20%; the included angle between the axis of the second air film hole of the air film air outlet area and the surface of the blade top groove is theta 1, and theta 1 is more than or equal to 10 degrees and less than or equal to 40 degrees; the included angle between the axis of the second air film hole in the area and the tangent direction of the mean camber line is delta 1, and delta 1 is more than or equal to 0 degree and less than or equal to 40 degrees.
6. The film cooling structure for a blade of a combustion engine according to claim 1, wherein: a second said film exit region from said leading edge of said blade tip, said film exit region being positioned at a mid-chord length of from 35% to 40%; the included angle between the axis of the second air film hole of the air film air outlet area and the surface of the blade top groove is theta 2, theta 2 is more than or equal to 20 degrees and less than or equal to 50 degrees, and the included angle between the axis of the second air film hole in the area and the tangential direction of a mean camber line is delta 2, delta 2 is more than or equal to-10 degrees and less than or equal to 30 degrees.
7. The film cooling structure for a blade of a combustion engine according to claim 1, wherein: a third said film exit region from said leading edge of said blade tip, said film exit region being positioned at a mid-chord length of between 50% and 55%; the included angle between the axis of the second air film hole of the air film air outlet area and the surface of the blade top groove is theta 3, the included angle between the axis of the second air film hole and the blade top groove is more than or equal to 20 degrees and less than or equal to 50 degrees, and the included angle between the axis of the second air film hole in the area and the tangential direction of a mean camber line is delta 3, and the included angle is more than or equal to-30 degrees and less than or equal to delta 3 and less than or equal to-10 degrees.
8. The film cooling structure for a blade of a combustion engine according to claim 1, wherein: a fourth said film exit region from said tip leading edge, said film exit region being positioned at a mid-chord length of between 85% and 90%; the included angle between the axis of the second air film hole of the air film air outlet area and the surface of the blade top groove is theta 4, and the theta 4 is more than or equal to 60 degrees and less than or equal to 90 degrees; the included angle between the axis of the second air film hole in the area and the tangent direction of the mean camber line is delta 4, delta 4 is more than or equal to minus 30 degrees and less than or equal to 0 degree.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202110834124.XA CN115701846A (en) | 2021-07-20 | 2021-07-20 | Air film cooling structure for gas turbine blade |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202110834124.XA CN115701846A (en) | 2021-07-20 | 2021-07-20 | Air film cooling structure for gas turbine blade |
Publications (1)
| Publication Number | Publication Date |
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| CN115701846A true CN115701846A (en) | 2023-02-14 |
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ID=85162769
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| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN202110834124.XA Pending CN115701846A (en) | 2021-07-20 | 2021-07-20 | Air film cooling structure for gas turbine blade |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN115701846A (en) |
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2021
- 2021-07-20 CN CN202110834124.XA patent/CN115701846A/en active Pending
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