CN113236370A - Cooling structure of high-pressure moving blade of turbine of gas turbine - Google Patents
Cooling structure of high-pressure moving blade of turbine of gas turbine Download PDFInfo
- Publication number
- CN113236370A CN113236370A CN202110571303.9A CN202110571303A CN113236370A CN 113236370 A CN113236370 A CN 113236370A CN 202110571303 A CN202110571303 A CN 202110571303A CN 113236370 A CN113236370 A CN 113236370A
- Authority
- CN
- China
- Prior art keywords
- channel
- blade
- equal
- holes
- cooling
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Granted
Links
Images
Classifications
-
- 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
-
- 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
- F01D5/187—Convection cooling
-
- Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
- Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
- Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
- Y02T50/00—Aeronautics or air transport
- Y02T50/60—Efficient propulsion technologies, e.g. for aircraft
Landscapes
- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
The invention provides a cooling structure of a high-pressure moving blade of a turbine of a gas turbine, belonging to the technical field of gas turbines. According to the invention, the interior of the blade is sequentially provided with a first, a second, a third, a fourth, a fifth and a sixth channel from the front edge of the blade to the tail edge of the blade; the first channel, the second channel, the third channel, the fourth channel, the fifth channel and the sixth channel are provided with turbulence ribs, the third channel, the second channel and the first channel are sequentially connected into a snake-shaped channel, the joint of adjacent channels forms a U-shaped space, the fourth channel, the fifth channel and the sixth channel are sequentially connected into a snake-shaped channel, and the joint of adjacent channels forms a U-shaped space; the third, fourth and seventh channels are provided with a first, a second and a third air inlet at the bottom; the tail edge of the blade is provided with a plurality of cleft seams through which air is exhausted; the blade top is provided with a blade top groove, and a cooling hole and a process hole are arranged in the blade top groove; the blade top pressure surface is provided with a blade top pressure surface cooling hole. The invention can implement sufficient cooling to the positions of the front edge of the blade, the suction surface of the blade, the top of the blade and the like, ensure that the temperature field of the blade root is more uniform, the temperature distribution of the blade body is reasonable, and the flow of the used cold air is less.
Description
Technical Field
The invention relates to the technical field of gas turbines, in particular to a cooling structure of a high-pressure moving blade of a turbine of a gas turbine.
Background
The most effective method for improving the overall efficiency and the working capacity of the gas turbine is to increase the gas temperature, and the gas inlet temperature of the current advanced heavy-duty gas turbine reaches the 1600 ℃ grade and is far higher than the use temperature of blade materials, so that a high-efficiency turbine blade cooling structure must be designed, and the blades are cooled to an acceptable temperature level by using the cooling air flow as little as possible. The high-pressure moving blade of the gas turbine has high heat load, small diameter of the front edge and high heat exchange coefficient, and can reach 2-4 times of the average heat exchange coefficient of the outer surface of the blade. This makes the leading edge region of the turbine high pressure moving blades one of the difficulties in cooling design. The middle position of the suction surface of the blade is provided with airflow flowing transition, the heat exchange coefficient of the gas side is increased, and the cooling requirement is higher. In addition, the blade top heat exchange coefficient of the movable blade is high, and the difficulty of cold air supply is high, so that the cooling of the blade top is difficult.
Chinese patent application CN103052765A discloses a gas turbine rotor blade, which is composed of the following components: a platform; a blade body having a cooling flow path including a meandering serpentine cooling flow path; a fillet portion disposed on a connection surface between the blade body and the platform; and a base portion including a cooling flow path communicating with the serpentine cooling flow path, wherein the cooling flow path includes a bypass flow path that branches from a high-pressure portion of the cooling flow path, is arranged along the rounded portion inside the rounded portion, and is connected to a low-pressure portion of the cooling flow path. The bypass flow path is led out from a high-pressure portion of the cooling flow path where the pressure of the cooling air is high, is arranged inside the rounded portion along the rounded portion, is connected to a low-pressure portion of the cooling flow path where the pressure is low, and circulates the cooling air by causing a part of the cooling air to flow in the bypass flow path by a pressure difference of the cooling air between the high-pressure portion and the low-pressure portion, thereby reducing the amount of the cooling air. However, this solution has the following drawbacks: too much cooling chamber division and too small chamber area result in larger cold air flow resistance and too high pressure requirement of cooling air supply; the cooling air at the trailing edge of the blade is heated up upstream, the cooling capacity of the cooling air is greatly reduced, and the cooling performance of the trailing edge is reduced; the structure is too complex; the manufacturing cost and difficulty are high.
The prior art has at least the following disadvantages:
1. the manufacturing difficulty is high, and the cost is high;
2. the cooling effect on certain positions, such as the tail edge of the blade, is not sufficient;
3. the pressure requirement for the cold air supply is too high, and the difficulty of air supply of the cold air is high.
Disclosure of Invention
In order to solve the technical problems in the prior art, the invention provides a cooling structure of a high-pressure moving blade of a turbine of a gas turbine, which divides the interior of the blade into seven channels, and sequentially and respectively comprises a first channel, a second channel, a third channel, a fourth channel, a fifth channel, a sixth channel and a seventh channel from the front edge of the blade to the tail edge of the blade; the first channel, the second channel, the third channel, the fourth channel, the fifth channel and the sixth channel are provided with turbulence ribs, the third channel, the second channel and the first channel are sequentially connected into a snake-shaped channel, the joint of adjacent channels forms a U-shaped space, the fourth channel, the fifth channel and the sixth channel are sequentially connected into a snake-shaped channel, and the joint of adjacent channels forms a U-shaped space; three rows of air film holes are arranged on the front edge of the blade of the first channel; the fourth channel is provided with a row of air film holes on the pressure surface and the suction surface of the blade respectively, the fifth channel is provided with a row of air film holes on the pressure surface of the blade, the bottom of the third channel is provided with a first air inlet, the bottom of the fourth channel is provided with a second air inlet, and the bottom of the seventh channel is provided with a third air inlet; the seventh channel is provided with a flow disturbing column, the tail edge of the blade is provided with a plurality of cleft seams, and the gas is exhausted through the cleft seams; the blade top is provided with a blade top groove, and a cooling hole and a process hole are arranged in the blade top groove; the blade top pressure surface is provided with a blade top pressure surface cooling hole.
The invention provides a cooling structure of a high-pressure moving blade of a turbine of a gas turbine, which comprises:
the interior of the blade is divided into seven channels, and a first channel, a second channel, a third channel, a fourth channel, a fifth channel, a sixth channel and a seventh channel are sequentially arranged from the front edge of the blade to the tail edge of the blade;
turbulence ribs are arranged in the first channel, the second channel, the third channel, the fourth channel, the fifth channel and the sixth channel;
arranging a film hole at the front edge of the blade of the first channel;
the third channel, the second channel and the first channel are sequentially connected to form a snake-shaped channel, and a U-shaped space is formed at the joint of adjacent channels;
the bottom of the third channel is provided with a first air inlet;
the fourth channel, the fifth channel and the sixth channel are sequentially connected to form a snake-shaped channel, and a U-shaped space is formed at the joint of the adjacent channels;
a second air inlet is formed at the bottom of the fourth channel;
the fourth channel is provided with air film holes on the pressure surface and the suction surface of the blade respectively;
the fifth channel is provided with a gas film hole on the pressure surface of the blade;
the seventh channel is a straight channel, and a flow disturbing column is arranged in the seventh channel;
a third air inlet is formed at the bottom of the seventh channel;
the tail edge of the blade is provided with a plurality of cleft seams through which air is exhausted;
the blade top is provided with a blade top groove, a cooling hole and a process hole are arranged in the blade top groove, and a blade top pressure surface is provided with a blade top pressure surface cooling hole.
Preferably, the cooling holes in the tip pocket are plural, and the tip pressure surface cooling holes of the tip pressure surface are plural; three rows of air film holes, namely a first row of air film holes, a second row of air film holes and a third row of air film holes, are arranged at the front edge of the blade of the first channel; the fourth channel is provided with a row of air film holes respectively on the pressure surface and the suction surface of the blade, namely a fourth row of air film holes and a fifth row of air film holes respectively, and the fifth channel is provided with a row of air film holes on the pressure surface of the blade and is a sixth row of air film holes.
Preferably, the number of the process holes of the blade tip is four, the first process hole, the second process hole, the third process hole and the fourth process hole are respectively arranged in the blade tip grooves of the first channel, the third channel, the sixth channel and the seventh channel, the cold air entering from the first air inlet is discharged from the first process hole, the second process hole and the first row, the second row and the third row of the air film holes of the blade front edge of the first channel, the cold air entering from the second air inlet is discharged from the third process hole, the fourth row and the fifth row and the sixth row of the air film holes, and the cold air entering from the third air inlet is discharged from the fourth process hole and the cleft seam of the blade tail edge.
Preferably, the cooling holes have a compound angle, the film holes have a compound angle, and the compound angle of the film holes is 30 degrees.
Preferably, the gas film hole is a circular straight hole, and the pore diameter DM is in the range: DM is more than or equal to 0.02D1 and less than or equal to 0.04D1, and the pitch T range of the holes is as follows: t is more than or equal to 0.15D1 and less than or equal to 0.25D1, wherein D1 is the maximum thickness of the leaf profile.
Preferably, the blade tip groove is an incomplete groove, and a part of a blade tip side wall on the side of a blade tip pressure surface of the blade tip groove is cut off; the value range of the depth XT of the blade top groove is as follows: XT is more than or equal to 0.042D1 and less than or equal to 0.083D1, and the value range of the wall surface width YT at the two sides of the blade top groove is as follows: 0.050D1 is less than or equal to YT is less than or equal to 0.080D 1; the cooling holes provided in the tip pocket are located where the tip pocket sidewalls are cut away, where D1 is the maximum airfoil thickness.
Preferably, the spoiler ribs are arranged on the blade pressure surfaces and the blade suction surfaces of the first channel, the second channel, the third channel, the fourth channel, the fifth channel and the sixth channel of the blade; the included angle between the turbulence ribs and the airflow flowing direction is 45 degrees; the three rows of the flow disturbing columns are arranged in a staggered mode.
Preferably, the height S of the turbulator ribs ranges: s is more than or equal to 0.030D1 and less than or equal to 0.040D 1; rib width W range: w is more than or equal to 0.030D1 and less than or equal to 0.040D 1; rib pitch P range: p is more than or equal to 0.30D1 and less than or equal to 0.40D1, wherein D1 is the maximum thickness of the leaf profile.
Preferably, the cross section of the spoiler pillar 19 is circular, the spoiler pillar diameter YF range: YF is more than or equal to 0.090D1 and less than or equal to 0.180D1, and the height WF of the turbulence column is within the range: WF is more than or equal to 0.11D1 and less than or equal to 0.33D1, and the radial pitch ZF range of the turbulence column is as follows: ZF is more than or equal to 0.25D1 and less than or equal to 0.42D1, and the axial intercept XF range of the turbulence column is as follows: XF is more than or equal to 0.25D1 and less than or equal to 0.42D1, wherein D1 is the maximum thickness of the blade profile.
Preferably, the split width WL range: WL is more than or equal to 0.045D1 and less than or equal to 0.054D1, and the height UL range of the split seam: WL is more than or equal to 0.045D1 and less than or equal to 0.09D1, and the range of the splitting pitch ZL: 0.167D1 ZL ≤ 0.333D1, wherein D1 is the maximum thickness of the leaf profile.
Compared with the prior art, the invention has the following beneficial effects:
(1) the invention adopts three rows of front edge air film holes with compound angles, which can implement sufficient cooling to the front edge of the blade; in the invention, a row of air film holes are arranged on the suction surface of the blade, so that the suction surface of the blade can be cooled sufficiently; the blade top of the blade is provided with the non-complete groove, the blade top groove process hole, the cooling hole, the blade top pressure surface cooling hole and other accurate cooling structures, so that the blade top can be cooled sufficiently;
(2) the invention can uniformly intake air at the bottom of the blade, and can ensure that the temperature field of the blade root is relatively uniform;
(3) the invention uses a simpler snake-shaped channel structure, has relatively simple manufacture and uses less cold air flow.
Drawings
FIG. 1 is a longitudinal cross-sectional view of one embodiment of the present invention, wherein
Indicates the cold air flow direction;
FIG. 2 is a transverse cross-sectional view of one embodiment of the present invention taken along the line E-E in FIG. 1;
FIG. 3 is a longitudinal cross-sectional view of a first channel of one embodiment of the present invention;
FIG. 4 is an enlarged partial cross-sectional view taken along line U-U of FIG. 3;
FIG. 5 is a longitudinal cross-sectional view of a seventh channel according to an embodiment of the present invention;
FIG. 6 is an enlarged partial cross-sectional view taken along line V-V of FIG. 5;
FIG. 7 is a schematic view of tip cooling features of an embodiment of the present invention;
in the figure, 1-gas turbine high-pressure moving blade, 2A-first air inlet, 2B-second air inlet, 2C-third air inlet, 3-first channel, 4-second channel, 5-third channel, 6-fourth channel, 7-fifth channel, 8-sixth channel, 9-seventh channel, 10-blade leading edge, 11-turbulence rib, 12D-first row of air film holes, 12E-second row of air film holes, 12F-third row of air film holes, 12G-fourth row of air film holes, 12H-fifth row of air film holes, 12J-sixth row of air film holes, 13-blade tip, 14-blade tip groove, 15L-first process hole, 15M-second process hole, 15N-third process hole, 15P-fourth process hole, 16-blade tip groove inner cooling hole, 17Q-U-shaped space between the third channel and the second channel, 17R-U-shaped space between the second channel and the first channel, 17S-U-shaped space between the fourth channel and the fifth channel, 17T-U-shaped space between the fifth channel and the sixth channel, 18-blade trailing edge, 19-turbulence column, 20-cleft seam, 21-blade root, 22-blade root, 23-blade suction surface, 24-blade pressure surface, 25-blade tip pressure surface cooling hole, P-rib pitch, W-rib width, A1-turbulence rib and airflow direction included angle, A2-air film hole composite angle, S-rib height, ZF-turbulence column radial pitch, UL-cleft seam height, ZL-cleft seam pitch, XF-turbulence column axial pitch, YF-turbulence column diameter, WF-turbulence column height, WL-split width, YT-wall surface width at two sides of a blade top groove, XT-blade top groove depth, Cax-axial chord length and D1-maximum blade profile thickness.
Detailed Description
The technical scheme of the invention is further described according to a specific embodiment of the invention with reference to the attached drawings.
The present invention provides a cooling structure of a gas turbine high-pressure moving blade 1, in which structure,
the interior of the blade is divided into seven channels, namely a first channel 3, a second channel 4, a third channel 5, a fourth channel 6, a fifth channel 7, a sixth channel 8 and a seventh channel 9 from the front edge 10 of the blade to the tail edge 18 of the blade;
arranging a film hole at the blade leading edge 10 of the first channel 3;
the third channel 5, the second channel 4 and the first channel 3 are sequentially connected to form a snake-shaped channel, and a U-shaped space is formed at the joint of adjacent channels;
the bottom of the third channel 5 is provided with a first air inlet 2A;
the fourth channel 6, the fifth channel 7 and the sixth channel 8 are sequentially connected to form a snake-shaped channel, and the joint of adjacent channels forms a U-shaped space;
the bottom of the fourth channel 6 is provided with a second air inlet 2B;
the fourth channel 6 is provided with air film holes on the pressure surface 24 and the suction surface 23 of the blade respectively;
the fifth channel 7 is provided with a film hole on the pressure surface 24 of the blade;
the seventh channel 9 is a straight channel, and a flow disturbing column 19 is arranged in the seventh channel 9;
a third air inlet 2C is formed at the bottom of the seventh channel 9;
the blade tail edge 18 is provided with a plurality of cleft seams 20, and air is exhausted through the cleft seams 20;
the tip 13 has a tip pocket 14, in which tip pocket 14 cooling holes and process holes are provided, and the tip pressure face is provided with tip pressure face cooling holes 25.
According to a specific embodiment of the present invention, the cooling holes in the tip pocket 14 are plural, i.e., a plurality of tip pocket cooling holes 16, and the tip pressure surface cooling holes 25 of the tip pressure surface are plural; three rows of air film holes, namely a first row of air film holes 12D, a second row of air film holes 12E and a third row of air film holes 12F, are arranged at the blade front edge 10 of the first channel 3; the fourth channel 6 is provided with a row of air film holes respectively on the pressure surface 24 and the suction surface 23 of the blade, namely a fourth row of air film holes 12G and a fifth row of air film holes 12H, and the fifth channel 7 is provided with a row of air film holes respectively on the pressure surface 24 of the blade, namely a sixth row of air film holes 12J.
According to a specific embodiment of the present invention, the four process holes of the blade tip are arranged, the first process hole 15L, the second process hole 15M, the third process hole 15N and the fourth process hole 15P are respectively arranged in the blade tip groove 14 of the first channel 3, the third channel 5, the sixth channel 8 and the seventh channel 9, the cold air entering from the first air inlet 2A is discharged from the first process hole 15L, the second process hole 15M and the first row of air film holes 12D, the second row of air film holes 12E and the third row of air film holes 12F of the blade leading edge 10 of the first channel 3, the cold air entering from the second air inlet 2B is discharged from the third process hole 15N, the fourth row of air film holes 12G, the fifth row of air film holes 12H and the sixth row of air film holes, and the cold air entering from the third air inlet 2C is discharged from the fourth process hole 15P and the cleft 20 of the blade trailing edge 18.
According to a specific embodiment of the present invention, the cooling holes have a compound angle, the film holes have a compound angle, and the compound angle a2 of the film holes is 30 °.
According to a specific embodiment of the invention, the gas film holes are circular straight holes, and the pore diameter DM is in the range of: DM is more than or equal to 0.02D1 and less than or equal to 0.04D1, and the pitch T range of the holes is as follows: t is more than or equal to 0.15D1 and less than or equal to 0.25D1, wherein D1 is the maximum thickness of the leaf profile.
According to a specific embodiment of the present invention, the tip groove 14 is an incomplete groove, and a part of the tip sidewall on the tip pressure surface side of the tip groove 14 is cut off; the depth XT of the tip pocket 14 ranges from: XT is more than or equal to 0.042D1 and less than or equal to 0.083D1, and the value range of the wall surface width YT at the two sides of the blade top groove 14 is as follows: 0.050D1 is less than or equal to YT is less than or equal to 0.080D 1; the cooling holes provided in the tip pocket 14 are located where the sidewalls of the tip pocket 14 are cut away, where D1 is the maximum thickness of the airfoil.
According to a specific embodiment of the present invention, the turbulator ribs 11 are arranged on the blade pressure surface 24 and the blade suction surface 23 of the first channel 3, the second channel 4, the third channel 5, the fourth channel 6, the fifth channel 7 and the sixth channel 8 of the blade; an included angle A1 between the turbulence ribs 11 and the airflow flowing direction is 45 degrees; the turbulence columns 19 are three rows and are arranged in a staggered manner.
According to a specific embodiment of the present invention, the height S of the turbulator ribs 11 ranges: s is more than or equal to 0.030D1 and less than or equal to 0.040D 1; rib width W range: w is more than or equal to 0.030D1 and less than or equal to 0.040D 1; rib pitch P range: p is more than or equal to 0.30D1 and less than or equal to 0.40D 1.
According to a specific embodiment of the present invention, the cross section of the spoiler pillar 19 is circular, and the spoiler pillar diameter YF ranges: YF is more than or equal to 0.090D1 and less than or equal to 0.180D1, and the height WF of the turbulence column is within the range: WF is more than or equal to 0.11D1 and less than or equal to 0.33D1, and the radial pitch ZF range of the turbulence column is as follows: ZF is more than or equal to 0.25D1 and less than or equal to 0.42D1, and the axial intercept XF range of the turbulence column is as follows: XF is more than or equal to 0.25D1 and less than or equal to 0.42D1, wherein D1 is the maximum thickness of the blade profile.
According to a particular embodiment of the invention, the width of the cleft WL ranges: WL is more than or equal to 0.045D1 and less than or equal to 0.054D1, and the height UL range of the split seam: WL is more than or equal to 0.045D1 and less than or equal to 0.09D1, and the range of the splitting pitch ZL: 0.167D1 ZL ≤ 0.333D1, wherein D1 is the maximum thickness of the leaf profile.
Example 1
The embodiment is a cooling structure of a first-stage movable blade of a heavy-duty gas turbine.
Referring to fig. 1 and 2, a front view and a left side view of a cooling structure for a first stage bucket of a gas turbine are shown. In this example, the blade profile line section height is 5.50D1, chord length 4.05D1, axial chord length 3.11D1, where D1 is the maximum thickness of the blade profile.
Referring to fig. 1, the blade base has three inlets 2A, 2B and 2C, which are uniformly arranged in the axial direction at the blade root. The third channel, the second channel and the first channel are sequentially connected to form a snake-shaped channel, and a U-shaped space is formed at the joint of adjacent channels; the fourth channel, the fifth channel and the sixth channel are sequentially connected to form a serpentine channel, and the joint of adjacent channels forms a U-shaped space.
The bottom of the third channel 5 is provided with a first air inlet 2A; the bottom of the fourth channel 6 is provided with a second air inlet 2B; a third air inlet 2C is formed at the bottom of the seventh channel 7; the first air inlet 2A supplies cold air to the third channel 5, the second channel 4, the first channel 3 and the blade top, the second air inlet 2B supplies cold air to the fourth channel 6, the fifth channel 7, the sixth channel 8 and the blade top, and the third air inlet 2C supplies cold air to the seventh channel 9 and the blade top.
The blade top is provided with a blade top groove, a plurality of cooling holes and four process holes are arranged in the blade top groove, and a blade top pressure surface is provided with a blade top pressure surface cooling hole. Only one cooling hole within the tip pocket is shown in the cross-sectional view of FIG. 1, with the remaining plurality of cooling holes being visible in FIG. 7.
The cold air introduced from the first air inlet 2A is discharged from the first through holes 15L, the second through holes 15M and the first through holes 12D, the second through holes 12E and the third through holes 12F of the blade leading edge 10 of the first passage 3, the cold air introduced from the second air inlet 2B is discharged from the third through holes 15N, the fourth through holes 12G, the fifth through holes 12H and the sixth through holes, and the cold air introduced from the third air inlet 2C is discharged from the fourth through holes 15P and the slit 20 of the blade trailing edge 18.
The cooling air entering from the first air inlet 2A flows through the third channel, the U-shaped space between the third channel and the second channel, the U-shaped space between the second channel and the first channel, and the first channel in sequence.
The cooling air entering from the second air inlet 2B flows through the fourth channel, the U-shaped space between the fourth channel and the fifth channel, the U-shaped space between the fifth channel and the sixth channel, and the sixth channel in this order.
The cooling air entering from the third air intake port 2C passes through the straight passage seventh passage.
Referring to fig. 3 and 4, a longitudinal cross-sectional view of the first channel and a partial enlarged view thereof, respectively. The rib height and the rib width of the first channel 3 are respectively 0.033D1 and 0.033D1, the rib pitch is 0.33D1, and the turbulence ribs and the flow direction of the cooling air form an included angle of 45 degrees. The first channel 3 has 3 rows of film holes near the leading edge 10 of the blade. The air film holes are circular holes, the diameter of each air film hole is 0.02D1, and the compound angle of the air film holes is 30 degrees. The distance between the centers of the three rows of holes on the wall surface of the blade is more than 0.03D 1. The top of the first channel 3 is provided with 1 fabrication hole, namely a first fabrication hole.
The heights of the turbulence ribs, the widths of the turbulence ribs, the included angles between the turbulence ribs and the flow direction of the cooling air and the arrangement of the turbulence ribs of the second channel, the third channel, the fourth channel, the fifth channel and the sixth channel are the same as those of the first channel.
1 row of air film holes are respectively arranged on the pressure surface and the suction surface of the blade of the fourth channel; and the fifth passage is provided with 1 row of film holes on the pressure surface of the blade, the diameter, the angle and the pitch of the film holes are the same as those of the first passage, and D1 is the maximum thickness of the blade profile.
Referring to fig. 5 and 6, the seventh channel is a turbulator cooling channel that exhausts air through the split at the trailing edge of the blade and the blade tip. The turbulence columns 19 are 3 rows in total and are arranged in a staggered manner, the cross section of each turbulence column is circular, the diameter of each turbulence column is 0.12D1, the radial intercept of each turbulence column is 0.24D1, the axial intercept of each turbulence column is 0.29D1, and the average heights of the three rows of turbulence columns are 0.14D1, 0.25D1 and 0.33D1 respectively. The detailed sizes of the cleft of the blade tail edge are as follows: the width of the cleft seam is 0.045D1, the height is 0.07D1, and the intercept is 0.20D1, wherein D1 is the maximum thickness of the leaf profile.
Referring to FIG. 7, a recessed tip was used, the depth of the tip recess being 0.075D1, where D1 is the maximum profile thickness and the wall thickness outside the tip recess, pressure and suction sides of the blade is 0.05D 1. And (3) cutting off the blade top at the position of the pressure surface with the relative axial length of 20-100% to form an incomplete blade top structure, wherein D1 is the maximum thickness of the blade profile.
Referring to FIG. 7, the tip is arranged with cooling holes 16 and process holes. The pressure surface of the blade top is provided with 10 cooling holes, the hole diameter is 0.04D1, and the compound angle of the ten cooling holes changes along with the curvature of the local blade profile. The blade top groove is internally designed with 6 cooling holes 16 with the diameter of 0.03D 1. The first channel, the third channel, the sixth channel and the seventh channel are respectively provided with a casting process hole, namely a first process hole 15L, a second process hole 15M, a third process hole 15N and a fourth process hole 15P. The first fabrication hole 15L of the first channel, the second fabrication hole 15M of the third channel, and the third fabrication hole 15N of the sixth channel are circular holes, and the hole diameters are 0.14D1, 0.08D1, and 0.10D1, respectively. The fourth process hole 15P of the seventh passage has a rectangular shape with dimensions of 0.05D1 × 0.05D1, wherein D1 is the maximum thickness of the leaf profile.
The working principle is as follows:
when the blade is in work, cooling air enters the interior of the blade from the first air inlet 2A positioned at the bottom of the third channel 5 of the blade to generate a convection cooling effect on the blade, a small part of the cooling air is discharged from the second technical hole at the top of the blade of the third channel, and most of the cooling air flows into the second channel through the U-shaped space to continuously generate a cooling effect on the blade. The cooling air then flows into the first channel through a U-shaped space to continue cooling the blade, and the cooling air flows upward in the radial direction and is discharged from the first process holes located in the tip groove of the first channel, while the cooling air continuously flows out from the three rows of film holes located in the first channel 3, i.e., the leading edge of the blade. The cold air discharged from the film hole forms a film on the outer surface of the blade, and has a cooling effect. The flow and cooling characteristics of the cold air entering from the second air inlet 2B at the bottom of the fourth channel 6 of the blade are similar to those of the first inlet, and the cold air is discharged from the third process hole in the tip groove of the sixth channel 8.
The turbulence ribs arranged in the first channel, the second channel, the third channel, the fourth channel, the fifth channel and the sixth channel disturb the flow of cooling air, so that the heat exchange capacity is enhanced, meanwhile, the existence of the turbulence ribs is also equivalent to the increase of the heat exchange area of the cold air side, and the two functions enable the cooling effect of the blade to be better. Some experimental researches show that the heat exchange coefficient of the cooling air can be improved by 1-2 times by the turbulence ribs.
The cold air entering from the third air inlet 2C at the bottom of the seventh channel 9 of the blade flows upwards to generate a convection cooling effect on the interior of the blade, and simultaneously, the cold air continuously flows through the turbulence column 19 and is discharged from the cleft slit 20 at the trailing edge 18 of the blade, and finally, the residual cold air reaching the top of the blade is discharged from the process hole and the cooling hole at the top of the blade.
The three rows of flow disturbing columns of the seventh channel can generate strong disturbance to cooling air in the blade near the limited tail edge area of the blade, so that the heat exchange coefficient of cold air is improved, and meanwhile, the heat exchange area is increased by the flow disturbing columns, so that the cooling effect of the tail edge of the blade is improved. In addition, the turbulence column also connects the pressure surface and the suction surface of the blade, so that the metal temperature of the blade is more uniform, and the service life of the blade is prolonged.
The cleft seam of the blade tail edge of the seventh channel plays a role in exhausting, and meanwhile, the narrow and long cleft seam structure enables the flow speed of cold air to be high, the heat exchange coefficient to be high, and the blade tail edge can be effectively cooled.
The blade cooling structure of the invention realizes the full cooling of each part of the blade on the premise of less cold air consumption and simpler blade structure, so that the temperature distribution of the blade is reasonable, the stress is reasonable, the service life is qualified, and the requirements of all aspects are fully met.
The above description is only a preferred embodiment of the present invention and is not intended to limit the present invention, and various modifications and changes may be made by those skilled in the art. Any modification, equivalent replacement, or improvement made within the spirit and principle of the present invention shall fall within the protection scope of the present invention.
Claims (10)
1. A cooling structure of a high-pressure moving blade of a turbine of a gas turbine, characterized in that:
the interior of the blade is divided into seven channels, and a first channel, a second channel, a third channel, a fourth channel, a fifth channel, a sixth channel and a seventh channel are sequentially arranged from the front edge of the blade to the tail edge of the blade;
turbulence ribs are arranged in the first channel, the second channel, the third channel, the fourth channel, the fifth channel and the sixth channel;
arranging a film hole at the front edge of the blade of the first channel;
the third channel, the second channel and the first channel are sequentially connected to form a snake-shaped channel, and a U-shaped space is formed at the joint of adjacent channels;
the bottom of the third channel is provided with a first air inlet;
the fourth channel, the fifth channel and the sixth channel are sequentially connected to form a snake-shaped channel, and a U-shaped space is formed at the joint of the adjacent channels;
a second air inlet is formed at the bottom of the fourth channel;
the fourth channel is provided with air film holes on the pressure surface and the suction surface of the blade respectively;
the fifth channel is provided with a gas film hole on the pressure surface of the blade;
the seventh channel is a straight channel, and a flow disturbing column is arranged in the seventh channel;
a third air inlet is formed at the bottom of the seventh channel;
the tail edge of the blade is provided with a plurality of cleft seams through which air is exhausted;
the blade top is provided with a blade top groove, a cooling hole and a process hole are arranged in the blade top groove, and a blade top pressure surface is provided with a blade top pressure surface cooling hole.
2. The cooling structure for a high-pressure moving blade of a gas turbine according to claim 1, wherein a plurality of cooling holes are provided in the tip groove, and a plurality of tip pressure surface cooling holes are provided in the tip pressure surface; three rows of air film holes, namely a first row of air film holes, a second row of air film holes and a third row of air film holes, are arranged at the front edge of the blade of the first channel; the fourth channel is provided with a row of air film holes respectively on the pressure surface and the suction surface of the blade, namely a fourth row of air film holes and a fifth row of air film holes respectively, and the fifth channel is provided with a row of air film holes on the pressure surface of the blade and is a sixth row of air film holes.
3. The cooling structure for high-pressure moving blades of a turbine of a gas turbine as set forth in claim 1, wherein the number of the tooling holes of the tip of the blade is four, and the first, third, sixth and seventh passages have first, second, third and fourth tooling holes respectively formed in the tip groove thereof, and the cooling air introduced from the first air inlet is discharged from the first, second and third rows of the tooling holes of the first, second and first passages at the leading edge of the blade, and the cooling air introduced from the second air inlet is discharged from the third, fourth and fifth rows of the tooling holes and the sixth row of the tooling holes, and the cooling air introduced from the third air inlet is discharged from the fourth tooling holes and the slit of the trailing edge of the blade.
4. The cooling structure for high-pressure moving blades of a turbine of a gas turbine as claimed in claim 1, wherein said cooling holes have a compound angle, and said film holes have a compound angle, and the compound angle of the film holes is 30 °.
5. The cooling structure for high-pressure moving blades of a turbine of a gas turbine as claimed in claim 1, wherein the film holes are circular straight holes, and the diameter DM of the holes is in the range of: DM is more than or equal to 0.02D1 and less than or equal to 0.04D1, and the pitch T range of the holes is as follows: t is more than or equal to 0.15D1 and less than or equal to 0.25D1, wherein D1 is the maximum thickness of the leaf profile.
6. The cooling structure for high-pressure moving blades of a gas turbine according to claim 1, wherein the tip groove is an incomplete groove, and a portion of a tip sidewall on a tip pressure surface side of the tip groove is cut off; the value range of the depth XT of the blade top groove is as follows: XT is more than or equal to 0.042D1 and less than or equal to 0.083D1, and the value range of the wall surface width YT at the two sides of the blade top groove is as follows: 0.050D1 is less than or equal to YT is less than or equal to 0.080D 1; the cooling holes provided in the tip pocket are located where the tip pocket sidewalls are cut away, where D1 is the maximum airfoil thickness.
7. The cooling structure of high-pressure moving blades of a gas turbine according to claim 1, wherein the turbulator ribs are arranged on the blade pressure surface and the blade suction surface of the first passage, the second passage, the third passage, the fourth passage, the fifth passage and the sixth passage of the blade; the included angle between the turbulence ribs and the airflow flowing direction is 45 degrees; the three rows of the flow disturbing columns are arranged in a staggered mode.
8. The cooling structure of high-pressure moving blades of a gas turbine according to claim 1, wherein the height S of the turbulator rib is in the range of: s is more than or equal to 0.030D1 and less than or equal to 0.040D 1; rib width W range: w is more than or equal to 0.030D1 and less than or equal to 0.040D 1; rib pitch P range: p is more than or equal to 0.30D1 and less than or equal to 0.40D1, wherein D1 is the maximum thickness of the leaf profile.
9. The cooling structure for high-pressure moving blades of a turbine of a gas turbine as claimed in claim 1, wherein the cross section of the spoiler pillar is circular, and the diameter YF of the spoiler pillar is in the range of: YF is more than or equal to 0.090D1 and less than or equal to 0.180D1, and the height WF of the turbulence column is within the range: WF is more than or equal to 0.11D1 and less than or equal to 0.33D1, and the radial pitch ZF range of the turbulence column is as follows: ZF is more than or equal to 0.25D1 and less than or equal to 0.42D1, and the axial intercept XF range of the turbulence column is as follows: XF is more than or equal to 0.25D1 and less than or equal to 0.42D1, wherein D1 is the maximum thickness of the blade profile.
10. The cooling structure for a high-pressure moving blade of a gas turbine according to claim 1, wherein the width of the cleavage slit WL ranges from: WL is more than or equal to 0.045D1 and less than or equal to 0.054D1, and the height UL range of the split seam: WL is more than or equal to 0.045D1 and less than or equal to 0.09D1, and the range of the splitting pitch ZL: 0.167D1 ZL ≤ 0.333D1, wherein D1 is the maximum thickness of the leaf profile.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202110571303.9A CN113236370B (en) | 2021-05-25 | 2021-05-25 | Cooling structure of high-pressure moving blade of gas turbine |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202110571303.9A CN113236370B (en) | 2021-05-25 | 2021-05-25 | Cooling structure of high-pressure moving blade of gas turbine |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| CN113236370A true CN113236370A (en) | 2021-08-10 |
| CN113236370B CN113236370B (en) | 2023-04-25 |
Family
ID=77138691
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN202110571303.9A Active CN113236370B (en) | 2021-05-25 | 2021-05-25 | Cooling structure of high-pressure moving blade of gas turbine |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN113236370B (en) |
Cited By (7)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN113605992A (en) * | 2021-08-26 | 2021-11-05 | 华能国际电力股份有限公司 | Gas turbine cooling blade with internal micro-channels |
| CN114109518A (en) * | 2021-11-29 | 2022-03-01 | 西安交通大学 | Turbine blade leading edge ribbed rotational flow-air film composite cooling structure |
| CN114215607A (en) * | 2021-11-29 | 2022-03-22 | 西安交通大学 | Turbine blade leading edge rotational flow cooling structure |
| CN114922734A (en) * | 2022-06-10 | 2022-08-19 | 南京航空航天大学 | Uniform temperature nature rectification extension board steam anti-icing structure based on rib post subregion vortex |
| CN114991879A (en) * | 2022-07-18 | 2022-09-02 | 陕西联信材料科技有限公司 | High-precision turbine blade assembly of aircraft engine and manufacturing method thereof |
| CN115093231A (en) * | 2022-06-23 | 2022-09-23 | 西安鑫垚陶瓷复合材料有限公司 | Ceramic matrix composite guide vane with tail edge cleft and preparation method thereof |
| CN117418906A (en) * | 2023-12-19 | 2024-01-19 | 哈尔滨工业大学 | Turbine internal cold air structure based on fractal theory |
Citations (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| EP1094199A1 (en) * | 1999-10-18 | 2001-04-25 | ABB (Schweiz) AG | Rotor for a gas turbine |
| DE602004003331T2 (en) * | 2004-01-14 | 2007-06-21 | Snecma | Cooling air outlet slots of turbine blades |
| CN103052765A (en) * | 2011-03-11 | 2013-04-17 | 三菱重工业株式会社 | Turbine blade and gas turbine |
| CN109538304A (en) * | 2018-11-14 | 2019-03-29 | 哈尔滨工程大学 | A kind of miniature turbo blade combination cooling structure that staggeredly rib is combined with air film hole |
| CN111810245A (en) * | 2020-07-20 | 2020-10-23 | 浙江燃创透平机械股份有限公司 | Cooling structure of turbine rotor of gas turbine |
-
2021
- 2021-05-25 CN CN202110571303.9A patent/CN113236370B/en active Active
Patent Citations (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| EP1094199A1 (en) * | 1999-10-18 | 2001-04-25 | ABB (Schweiz) AG | Rotor for a gas turbine |
| DE602004003331T2 (en) * | 2004-01-14 | 2007-06-21 | Snecma | Cooling air outlet slots of turbine blades |
| CN103052765A (en) * | 2011-03-11 | 2013-04-17 | 三菱重工业株式会社 | Turbine blade and gas turbine |
| CN109538304A (en) * | 2018-11-14 | 2019-03-29 | 哈尔滨工程大学 | A kind of miniature turbo blade combination cooling structure that staggeredly rib is combined with air film hole |
| CN111810245A (en) * | 2020-07-20 | 2020-10-23 | 浙江燃创透平机械股份有限公司 | Cooling structure of turbine rotor of gas turbine |
Cited By (10)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN113605992A (en) * | 2021-08-26 | 2021-11-05 | 华能国际电力股份有限公司 | Gas turbine cooling blade with internal micro-channels |
| CN114109518A (en) * | 2021-11-29 | 2022-03-01 | 西安交通大学 | Turbine blade leading edge ribbed rotational flow-air film composite cooling structure |
| CN114215607A (en) * | 2021-11-29 | 2022-03-22 | 西安交通大学 | Turbine blade leading edge rotational flow cooling structure |
| CN114922734A (en) * | 2022-06-10 | 2022-08-19 | 南京航空航天大学 | Uniform temperature nature rectification extension board steam anti-icing structure based on rib post subregion vortex |
| CN115093231A (en) * | 2022-06-23 | 2022-09-23 | 西安鑫垚陶瓷复合材料有限公司 | Ceramic matrix composite guide vane with tail edge cleft and preparation method thereof |
| CN115093231B (en) * | 2022-06-23 | 2023-09-01 | 西安鑫垚陶瓷复合材料有限公司 | Ceramic matrix composite guide vane with trailing edge split joint and preparation method thereof |
| CN114991879A (en) * | 2022-07-18 | 2022-09-02 | 陕西联信材料科技有限公司 | High-precision turbine blade assembly of aircraft engine and manufacturing method thereof |
| CN114991879B (en) * | 2022-07-18 | 2022-11-15 | 陕西联信材料科技有限公司 | High-precision turbine blade assembly of aircraft engine and manufacturing method thereof |
| CN117418906A (en) * | 2023-12-19 | 2024-01-19 | 哈尔滨工业大学 | Turbine internal cold air structure based on fractal theory |
| CN117418906B (en) * | 2023-12-19 | 2024-03-22 | 哈尔滨工业大学 | Turbine internal cold air structure based on fractal theory |
Also Published As
| Publication number | Publication date |
|---|---|
| CN113236370B (en) | 2023-04-25 |
Similar Documents
| Publication | Publication Date | Title |
|---|---|---|
| CN113236370B (en) | Cooling structure of high-pressure moving blade of gas turbine | |
| US8753083B2 (en) | Curved cooling passages for a turbine component | |
| US10920597B2 (en) | Turbine blade cooling system with channel transition | |
| US8523527B2 (en) | Apparatus for cooling a platform of a turbine component | |
| US6554571B1 (en) | Curved turbulator configuration for airfoils and method and electrode for machining the configuration | |
| EP2547871B1 (en) | Gas turbine airfoil with shaped trailing edge coolant ejection holes and corresponding turbine | |
| US8052395B2 (en) | Air cooled bucket for a turbine | |
| CN111927562A (en) | Turbine rotor blade and aircraft engine | |
| CN110410158B (en) | Turbine rotor blade of gas turbine | |
| EP2383436B1 (en) | Airfoil trailing edge and method of manufacturing the same | |
| CN111810245B (en) | Cooling structure of gas turbine rotor plate | |
| CN210118169U (en) | Low-flow low-aspect-ratio high-pressure turbine cooling guide vane | |
| CA2953594A1 (en) | Turbine airfoil trailing edge cooling passage | |
| EP3669054B1 (en) | Turbine blade and corresponding method of servicing | |
| CN112922674B (en) | Turbine blade with air film cooling groove | |
| CN113550794B (en) | Multi-cavity efficient cooling structure and cooling method for turbine rotor blade | |
| CN112282860A (en) | Turbine rotor blade platform cooling structure | |
| CN111425261A (en) | Turbine blade cooling structure suitable for integrated printing and forming and engine | |
| CN112922676B (en) | Internal back basin rotary cooling channel of turbine blade | |
| CN114876583B (en) | Cooling structure of turbine movable blade trailing edge | |
| CN114412577B (en) | Turbine moving blade | |
| CN112943380A (en) | Rotary cooling channel turbine blade adopting T-shaped partition wall | |
| CN116537886A (en) | Blade top structure and blade with improve cooling efficiency | |
| CN116085055A (en) | Component with cooling channels for a turbine engine | |
| CN113404548A (en) | Blade, turbine and gas turbine |
Legal Events
| Date | Code | Title | Description |
|---|---|---|---|
| PB01 | Publication | ||
| PB01 | Publication | ||
| SE01 | Entry into force of request for substantive examination | ||
| SE01 | Entry into force of request for substantive examination | ||
| GR01 | Patent grant | ||
| GR01 | Patent grant | ||
| CP01 | Change in the name or title of a patent holder |
Address after: No. 1188, Dongxin Road, Hangzhou, Zhejiang, 310022 Patentee after: Hangzhou Steam Turbine Holding Co.,Ltd. Address before: No. 1188, Dongxin Road, Hangzhou, Zhejiang, 310022 Patentee before: HANGZHOU TURBINE POWER GROUP CO.,LTD. |
|
| CP01 | Change in the name or title of a patent holder |