CN116950723B - Low-stress double-wall turbine guide vane cooling structure and design method thereof - Google Patents
Low-stress double-wall turbine guide vane cooling structure and design method thereof Download PDFInfo
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- 238000001816 cooling Methods 0.000 title claims abstract description 100
- 238000000034 method Methods 0.000 title claims abstract description 14
- 238000005192 partition Methods 0.000 claims description 13
- 238000009413 insulation Methods 0.000 claims description 4
- 230000007423 decrease Effects 0.000 claims description 2
- 230000000694 effects Effects 0.000 description 8
- 230000035882 stress Effects 0.000 description 8
- 230000008646 thermal stress Effects 0.000 description 6
- 230000002035 prolonged effect Effects 0.000 description 3
- 238000010586 diagram Methods 0.000 description 2
- 238000005516 engineering process Methods 0.000 description 2
- 238000004378 air conditioning Methods 0.000 description 1
- 238000013142 basic testing Methods 0.000 description 1
- 230000009286 beneficial effect Effects 0.000 description 1
- 238000007664 blowing Methods 0.000 description 1
- 238000005266 casting Methods 0.000 description 1
- 238000012938 design process Methods 0.000 description 1
- 230000002045 lasting effect Effects 0.000 description 1
- 238000003754 machining Methods 0.000 description 1
- 239000000463 material Substances 0.000 description 1
- 238000012986 modification Methods 0.000 description 1
- 230000004048 modification Effects 0.000 description 1
- 239000007921 spray Substances 0.000 description 1
- 239000012720 thermal barrier coating Substances 0.000 description 1
- 238000011144 upstream manufacturing Methods 0.000 description 1
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Classifications
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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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- 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
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- G—PHYSICS
- G06—COMPUTING; CALCULATING OR COUNTING
- G06F—ELECTRIC DIGITAL DATA PROCESSING
- G06F30/00—Computer-aided design [CAD]
- G06F30/10—Geometric CAD
- G06F30/17—Mechanical parametric or variational design
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- G—PHYSICS
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- G06F—ELECTRIC DIGITAL DATA PROCESSING
- G06F30/00—Computer-aided design [CAD]
- G06F30/20—Design optimisation, verification or simulation
- G06F30/28—Design optimisation, verification or simulation using fluid dynamics, e.g. using Navier-Stokes equations or computational fluid dynamics [CFD]
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- G—PHYSICS
- G06—COMPUTING; CALCULATING OR COUNTING
- G06F—ELECTRIC DIGITAL DATA PROCESSING
- G06F2113/00—Details relating to the application field
- G06F2113/08—Fluids
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- G—PHYSICS
- G06—COMPUTING; CALCULATING OR COUNTING
- G06F—ELECTRIC DIGITAL DATA PROCESSING
- G06F2119/00—Details relating to the type or aim of the analysis or the optimisation
- G06F2119/08—Thermal analysis or thermal optimisation
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- G—PHYSICS
- G06—COMPUTING; CALCULATING OR COUNTING
- G06F—ELECTRIC DIGITAL DATA PROCESSING
- G06F2119/00—Details relating to the type or aim of the analysis or the optimisation
- G06F2119/14—Force analysis or force optimisation, e.g. static or dynamic forces
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Abstract
The invention relates to the technical field of turbine blade structure design, and discloses a low-stress double-wall turbine guide blade cooling structure and a design method thereof.
Description
Technical Field
The invention relates to the technical field of turbine blade structure design, and discloses a low-stress double-wall turbine guide blade cooling structure and a design method thereof.
Background
In high-performance aeroengines, because the temperature of the gas at the inlet of the turbine exceeds the bearing limit of the turbine blade material, and the effect of the thermal barrier coating on the improvement of the temperature bearing capacity of the blade is limited, effective cooling measures are required to ensure the reliable operation of the turbine blade in the high-temperature high-pressure high-rotation-speed environment.
Conventional blade cooling often employs conventional scale impingement, convection, bypass intensified internal cooling and film outflow based external cooling schemes. However, cooling technologies at conventional scales have matured over decades, and the speed of increase has been difficult to keep pace with the speed of increase in turbine inlet temperatures, and new cooling technologies must be developed.
At present, various super-strong cooling structures, such as a Lamiloy laminate cooling structure in the United kingdom, have been proposed, and a multi-layer channel is designed in the wall surface of a turbine blade, so that on one hand, heat exchange is enhanced through a turbulence structure, and on the other hand, the cooling air flow is deflected through the channel to realize the maximization of the flow heat exchange wetting surface area, so that the heat exchange effect is remarkably improved. However, the design process of the super-strong cooling structure is complex, a large amount of basic test data is needed for supporting, and the engineering realization difficulty is extremely high; particularly, in the process of machining and casting, the dimensional deviation of the cooling structure can bring great deviation of cooling effect, and the laminate cooling structure cannot be widely used in a large scale.
Disclosure of Invention
The invention aims to provide a low-stress double-wall turbine guide vane cooling structure and a design method thereof, which can realize efficient cooling under the condition of limited cold air quantity, avoid the problem of overlarge thermal stress caused by high temperature difference between inner and outer walls, and greatly prolong the service life of turbine vanes due to low vane stress.
In order to achieve the technical effects, the technical scheme adopted by the invention is as follows:
the utility model provides a low stress double-walled turbine guide vane cooling structure, includes first air film hole and the first impact hole of setting on double-walled turbine guide vane, first impact hole is rectangular shape structure, the length direction of first impact hole is the leaf height direction, first impact hole sets up in leading edge inner wall, first air film hole sets up in double-walled turbine guide vane's leading edge outer wall.
Further, the aspect ratio of the first impingement holes is (2-5): 1.
further, four corners of the first impact hole are rounded.
Further, a first partition plate is arranged between the inner wall and the outer wall corresponding to the front edge residence point position of the double-wall turbine guide vane in the radial direction, the front edge of the vane is divided into two mutually independent impact cavities by the first partition plate, each impact cavity is communicated with a first impact hole, and a first air film hole close to the first partition plate is crossed and penetrates through the first partition plate to form an air film hole structure with an inclined angle on the surface of the outer wall of the front edge.
Further, at least one second impact hole is formed in the inner wall, corresponding to the chord zone, of the double-wall turbine guide blade, the second impact holes are of strip-shaped structures, and each second impact hole is provided with a corresponding second air film hole at the downstream of the gas flow on the outer wall of the blade.
Further, the second air film holes are distributed in a variable density in the middle chord area of the blade, and the distribution density is gradually reduced along the direction of the gas flow.
In order to achieve the above technical effects, the present invention further provides a method for designing a low-stress double-wall turbine guide vane cooling structure, which is used for obtaining the low-stress double-wall turbine guide vane cooling structure, and includes:
obtaining the Reynolds number and the Knudsen number of a leading edge cooling structure of the double-wall turbine guide vane according to the leading edge cooling requirement of the double-wall turbine guide vane and cooling airflow parameters, wherein the cooling airflow parameters comprise the inlet pressure of the leading edge cooling airflow, the temperature of the cooling airflow and the mass flow of the leading edge cooling airflow;
design flow coefficient functional relation of limited impact channels under different impact distances according to leading edge cooling structures of double-layer turbine guide bladesCalculating to obtain the equivalent diameter of the first impact hole under the corresponding impact distance; wherein->Design flow coefficient for blade leading edge film impingement cooling structure +.>Reynolds number for blade leading edge film impingement cooling structure,/->For the number of knudsen->For the impact distance>Is the equivalent diameter of the first impingement hole;
analyzing and obtaining the flow area of the strip-shaped first impact hole according to the calculated equivalent diameter of the first impact hole;
the length and width of the opening of the first impingement hole are calculated based on the flow area of the first impingement hole and the aspect ratio of the first impingement hole.
Further, the method further comprises the following steps:
calculating the actual flow of the first impingement hole according to the inlet pressure of cooling air flow at the front edge of the double-layer turbine guide vane, the temperature of the cooling air flow, the outlet pressure of the first air film hole and the design flow coefficient of the air film impingement cooling structure at the front edge of the double-layer turbine guide vaneWherein->For the flow area of the first impingement hole +.>Air inlet pressure for front edge cooling air flow, +.>For the first film hole outlet pressure, +.>Is the thermal insulation coefficient of air, and the heat insulation coefficient of air,is the gas constant of air->Is the cooling air flow temperature;
judging the actual flow of the first impact holeWhether the relative deviation from the mass flow of the front cooling air flow is within a predetermined range, if the actual flow of the first impingement holes +.>If the relative deviation between the calculated first impingement hole and the mass flow of the front edge cooling air flow is within a preset range, calculating that the obtained open length and width data of the first impingement hole meet the design requirement, otherwise, adjusting the impingement interval or the cooling air flow parameters to enable the actual flow of the first impingement hole to be +.>The relative deviation from the mass flow of the leading edge cooling air flow is within a preset range.
Compared with the prior art, the invention has the following beneficial effects: according to the invention, an air inlet mode of a strip-shaped first impact hole structure is adopted, air flow enters an impact cavity channel between the inner wall and the outer wall of the double-layer wall turbine guide vane through the strip-shaped first impact hole/seam, is discharged through an air film hole on the outer wall of the front edge, forms an air film cover on the surface behind the vane, and realizes efficient cooling under the condition of limited cold air quantity.
Drawings
FIG. 1 is a schematic illustration of a low stress double wall turbine guide vane cooling structure in accordance with an embodiment;
FIG. 2 is a schematic illustration of the positional relationship of a first impingement hole, a first film hole, and a first diaphragm of a double-walled turbine guide vane in an embodiment;
FIG. 3 is a schematic illustration of a relationship between the first impingement holes and the second impingement holes of a double-walled turbine guide vane according to an embodiment;
wherein, 1, double-walled turbine guide vanes; 2. a first air film hole; 3. a first impingement hole; 4. a first separator; 5. an impingement cavity; 6. a second impingement hole; 7. and a second air film hole.
Detailed Description
The present invention will be described in further detail with reference to the following examples and drawings. It should not be construed that the scope of the above subject matter of the present invention is limited to the following embodiments, and all techniques realized based on the present invention are within the scope of the present invention.
Examples
Referring to fig. 1-3, a low-stress double-wall turbine guide vane cooling structure comprises a first air film hole 2 and a first impact hole 3 which are arranged on a double-wall turbine guide vane 1, wherein the first impact hole 3 is of a strip-shaped structure, the length direction of the first impact hole 3 is the vane height direction, the first impact hole 3 is arranged on the inner wall of the front edge, and the first air film hole 2 is arranged on the outer wall of the front edge of the double-wall turbine guide vane 1.
In this embodiment, adopt the air inlet mode of rectangular shape first impact hole 3 structure, the air current gets into in the impact chamber 5 ways between double-walled turbine guide vane 1 inner wall and the outer wall through rectangular shape first impact hole 3, by the air film hole discharge on the leading edge outer wall, and form the air film to the surface behind the blade and cover, realize high-efficient cooling under limited air conditioning volume condition, rectangular shape first impact hole 3 structure can release the inner wall simultaneously, the thermal stress between the outer wall surface, avoid because the high difference in temperature between the inner and outer wall brings the too big problem of thermal stress, the blade stress is low, thereby promote turbine vane's life-span by a wide margin.
To achieve a satisfactory leading edge cooling effect, the present embodiment also provides a method for designing a low stress double-walled turbine guide vane cooling structure, comprising:
obtaining the Reynolds number and the Knudsen number of the front edge cooling structure of the double-wall turbine guide vane 1 according to the front edge cooling requirement of the double-wall turbine guide vane 1 and cooling air flow parameters, wherein the cooling air flow parameters comprise the inlet pressure of the front edge cooling air flow, the temperature of the cooling air flow and the mass flow of the front edge cooling air flow;
design flow coefficient function relationship of limited impact channels under different impact distances according to front edge cooling structure of double-wall turbine guide vane 1Calculating to obtain the equivalent diameter of the first impact hole 3 under the corresponding impact distance; wherein->Design flow coefficient for blade leading edge film impingement cooling structure +.>Reynolds number for blade leading edge film impingement cooling structure,/->For the number of knudsen->For the impact distance>Is the equivalent diameter of the first impingement hole 3;
analyzing and obtaining the flow area of the strip-shaped first impact hole 3 according to the calculated equivalent diameter of the first impact hole 3;
the opening length and width of the first impingement holes 3 are calculated from the flow area of the first impingement holes 3 and the aspect ratio of the first impingement holes 3.
Further, the method further comprises the following steps:
according to the front of the double-wall turbine guide vane 1The actual flow of the first impingement holes 3 is calculated by the design flow coefficient of the edge cooling air flow inlet pressure, the cooling air flow temperature, the outlet pressure of the first air film holes 2 and the front edge air film impingement cooling structure of the double-layer turbine guide vane 1Wherein->For the flow area of the first impingement hole 3, < > and>air inlet pressure for front edge cooling air flow, +.>For the outlet pressure of the first air film hole 2 +.>For the air insulation factor, +.>=1.4；/>The gas constant is that of air, in this embodiment +.>=286.7，/>For cooling the air flow temperature.
Judging the actual flow rate of the first impingement holes 3Whether the relative deviation from the mass flow of the front cooling air flow is within a preset range, such as + -10% in the present embodiment, if the actual flow of the first impingement holes 3 +.>If the relative deviation between the calculated first impingement hole 3 and the mass flow of the front edge cooling air flow is within a preset range of +/-10%, the calculated open length and width data of the first impingement hole 3 meet the design requirements, otherwise, the impingement interval or the cooling air flow parameters are adjusted to ensure that the actual flow of the first impingement hole 3 is ±>The relative deviation from the mass flow of the leading edge cooling airflow is within a preset range of + -10%.
To ensure the structural strength of the double-walled turbine guide vane 1, the four corners of the first impingement holes 3 are rounded in this embodiment. The number of the first impact holes 3 may be one or more, and it is sufficient to ensure that the flow area of the first impact holes 3 satisfies the requirement.
On the basis of the above structural design, in this embodiment, a first partition plate 4 is radially disposed between the inner wall and the outer wall corresponding to the residence point position of the front edge of the double-wall turbine blade, the front edge of the blade is divided into two mutually independent impact cavities 5 by the first partition plate 4, each impact cavity 5 is communicated with a first impact hole 3, and the first air film hole 2 near the first partition plate 4 crosses through the first partition plate 4 to form an air film hole structure with an inclined angle on the outer wall surface of the front edge. Two impact cavities 5 are formed on the front edge, which is close to the two sides of the leaf basin and the leaf back according to the standing point position, through the first partition plate 4, and air flows to the standing point area along the channel of the impact cavities 5, so that heat of the wall surface is absorbed along the path. The air flow in the impact cavity 5 at the back side of the blade flows out to the blade basin side through the inclined air film holes in the area close to the stagnation point to form air film coverage. The air flow in the impact cavity 5 at the side of the leaf basin is close to the standing point area, flows out to the side of the leaf back through the inclined air film hole to form air film coverage, so that the air film coverage effect of the front edge cooling air flow can be improved, the problem that the air blowing ratio of the front edge area is higher due to the fact that the wall surface air film spray cooling structure and other vertical wall surface air vents are adopted conventionally is avoided, and the problem that the cooling effect is poor due to the fact that cold air is blown off the wall surface is further avoided.
By analyzing and comparing the front edge stress distribution diagram of the conventional small hole impact plus film cooling structure with the front edge stress distribution diagram of the cooling structure of the first film holes 2 crossed by the first long-strip impact holes 3, the first long-strip impact holes 3 can release the problem of overlarge thermal stress between the inner wall surface and the outer wall surface caused by high temperature difference, and the maximum stress is reduced by about 15 percent, so that the lasting life of the blade is prolonged by about 10 times.
In this embodiment, at least one second impact hole 6 is provided on the inner wall corresponding to the chord zone in the double-walled turbine guiding vane 1, the second impact holes 6 are in a strip-shaped structure, and each second impact hole 6 is provided with a corresponding second air film hole 7 at the downstream of the gas flow on the outer wall of the vane. The long strip-shaped second impact holes 6 are arranged at the upstream position of the inner wall of the double-layer wall cooling structure, the air flow enters the inner cavity of the chord zone in the double-layer wall turbine guide vane 1 through the long strip-shaped second impact holes 6, flows out from the downstream microscale air film hole row after passing through the inner turbulence structure, and forms an air film cover on the rear surface. The strip-shaped second impact holes 6 can uniformly release thermal stress between the inner wall and the outer wall, and the service life of the turbine blade is greatly prolonged.
In addition, the second air film holes 7 in the embodiment have variable density distribution in the chord zone of the blade, and the distribution density gradually decreases along the direction of the gas flow. By adopting the design of the arrangement of the second air film holes 7 of the variable density array (front dense and rear sparse), compared with the arrangement of the equidistant air film holes, the method can realize higher air film coverage efficiency and longer air film cooling coverage length under the same opening ratio of the second air film holes 7 and the same cold air flow condition, and meanwhile, the surface temperature distribution of the blade is more uniform, so that the surface temperature and the temperature gradient of the blade are further obviously reduced, the thermal stress of the blade is further reduced, and the service life of the turbine blade is prolonged.
The foregoing description of the preferred embodiments of the invention is not intended to be limiting, but rather is intended to cover all modifications, equivalents, and alternatives falling within the spirit and principles of the invention.
Claims (7)
1. The low-stress double-wall turbine guide vane cooling structure is characterized by comprising a first air film hole and a first impact hole, wherein the first air film hole and the first impact hole are arranged on a double-wall turbine guide vane, the first impact hole is of a strip-shaped structure, the length direction of the first impact hole is in the vane height direction, the first impact hole is arranged on the inner wall of the front edge, and the first air film hole is arranged on the outer wall of the front edge of the double-wall turbine guide vane; the double-wall turbine guide vane comprises a double-wall turbine guide vane, wherein a first partition plate is arranged between an inner wall and an outer wall corresponding to the front edge residence point of the double-wall turbine guide vane along the radial direction, the front edge of the double-wall turbine guide vane is divided into two mutually independent impact cavities by the first partition plate, each impact cavity is communicated with a first impact hole, and a first air film hole close to the first partition plate is crossed and penetrates through the first partition plate to form an air film hole structure with an inclined angle on the surface of the outer wall of the front edge.
2. The low stress double-walled turbine guide vane cooling structure of claim 1 wherein the aspect ratio of the first impingement holes is (2-5): 1.
3. the low stress double-walled turbine guide vane cooling structure of claim 1 wherein the four corners of the first impingement holes are rounded.
4. The low stress double-walled turbine guide vane cooling structure of claim 1 wherein at least one second impingement hole is provided in the inner wall of the double-walled turbine guide vane corresponding to the chord zone, the second impingement holes being of elongated configuration, each second impingement hole being provided with a corresponding second film hole downstream of the outer wall of the vane in the gas flow stream.
5. The low stress double-walled turbine guide vane cooling structure of claim 4 wherein said second film holes are distributed in a variable density in the mid-chord region of the vane with a distribution density that gradually decreases in the direction of the gas flow.
6. A method of designing a low stress double wall turbine guide vane cooling structure for obtaining a low stress double wall turbine guide vane cooling structure according to any one of claims 1 to 5, comprising:
obtaining the Reynolds number and the Knudsen number of a leading edge cooling structure of the double-wall turbine guide vane according to the leading edge cooling requirement of the double-wall turbine guide vane and cooling airflow parameters, wherein the cooling airflow parameters comprise the inlet pressure of the leading edge cooling airflow, the temperature of the cooling airflow and the mass flow of the leading edge cooling airflow;
design flow coefficient functional relation of limited impact channels under different impact distances according to leading edge cooling structures of double-layer turbine guide bladesCalculating to obtain the equivalent diameter of the first impact hole under the corresponding impact distance; wherein->Design flow coefficient for blade leading edge film impingement cooling structure +.>Reynolds number for blade leading edge film impingement cooling structure,/->For the number of knudsen->For the impact distance>Is the equivalent diameter of the first impingement hole;
analyzing and obtaining the flow area of the strip-shaped first impact hole according to the calculated equivalent diameter of the first impact hole;
the length and width of the opening of the first impingement hole are calculated based on the flow area of the first impingement hole and the aspect ratio of the first impingement hole.
7. The method of designing a low stress double wall turbine guide vane cooling structure of claim 6 further comprising:
calculating the actual flow of the first impingement hole according to the inlet pressure of cooling air flow at the front edge of the double-layer turbine guide vane, the temperature of the cooling air flow, the outlet pressure of the first air film hole and the design flow coefficient of the air film impingement cooling structure at the front edge of the double-layer turbine guide vaneWherein->For the flow area of the first impingement hole +.>Air inlet pressure for front edge cooling air flow, +.>For the first film hole outlet pressure, +.>Is the air insulation coefficient>Is the gas constant of air->Is the cooling air flow temperature;
judging the actual flow of the first impact holeWhether the relative deviation from the mass flow of the front cooling air flow is within a predetermined range, if the actual flow of the first impingement holes +.>The relative deviation between the mass flow rate of the front edge cooling airflow and the calculated open length and width data of the first impingement holes meet the designThe requirement is otherwise to adjust the impingement distance or the cooling air flow parameters such that the actual flow of the first impingement holes +.>The relative deviation from the mass flow of the leading edge cooling air flow is within a preset range.
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