CN113202567A - Design method for cooling structure of guide cooling blade edge plate of high-pressure turbine - Google Patents
Design method for cooling structure of guide cooling blade edge plate of high-pressure turbine Download PDFInfo
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- CN113202567A CN113202567A CN202110569194.7A CN202110569194A CN113202567A CN 113202567 A CN113202567 A CN 113202567A CN 202110569194 A CN202110569194 A CN 202110569194A CN 113202567 A CN113202567 A CN 113202567A
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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
- F01D9/00—Stators
- F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
- F01D9/04—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector
- F01D9/041—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector using blades
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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
- F01D9/00—Stators
- F01D9/06—Fluid supply conduits to nozzles or the like
- F01D9/065—Fluid supply or removal conduits traversing the working fluid flow, e.g. for lubrication-, cooling-, or sealing fluids
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- 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
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- Turbine Rotor Nozzle Sealing (AREA)
Abstract
The application belongs to the field of aeroengine blades, and particularly relates to a cooling structure design method for a guide cooling blade flange plate of a high-pressure turbine. The method comprises the following steps: the method comprises the following steps that firstly, a flange plate is partitioned according to the air entraining position of the flange plate and the position of a vertical plate and is divided into a cold air introducing area and a non-cold air introducing area; designing an impact and air film combined cooling structure in a cold air introducing area of the edge plate; designing a flow disturbing column and air film combined cooling structure in a non-cold air introducing area of the edge plate; designing an air film cooling structure on the gas side surface of the edge plate and the blade body of the edge plate; and step five, arranging a variable-radius fillet on the edge plate and the blade body switching section along the blade profile. The cooling structure design method of the high-pressure turbine guide cooling blade flange plate adopts a combined cooling mode to cool the wall surface of the high-pressure turbine guide blade flange plate, can reduce the hot spot temperature of the flange plate, improves the cooling effect of the blade flange plate and prolongs the service life of the blade.
Description
Technical Field
The application belongs to the field of aeroengine blades, and particularly relates to a cooling structure design method for a guide cooling blade flange plate of a high-pressure turbine.
Background
Turbine blade platform cooling issues are increasingly pronounced as turbine front temperatures increase. The existence of cascade end wall effect is caused by the existence of fluid viscosity and the influence of radial, circumferential and axial pressure gradient, the blade flange plate area is the most complicated area of aerodynamic structure in the blade channel, the secondary flow phenomenon is very complicated, the existence of horseshoe vortex, channel vortex and angle vortex, and the adsorption separation of flow directly influence the heat load distribution of the blade, and a plurality of hot spot areas are caused, for example: the vortex separation device comprises a leading edge area, a suction surface convex hull area, a pressure surface and pressure surface angle area, a main channel vortex separation line downstream area, a cascade throat downstream area and a trailing edge wake area.
At present, the traditional cooling of the edge plate of the guide vane of the high-pressure turbine is mainly a film cooling mode, as shown in fig. 1-2, that is, a certain number of film holes are arranged on the flow passage surface of the edge plate for cooling the high-temperature area of the edge plate. The existing high-pressure turbine guide vane platform film cooling has the following disadvantages: the cooling structure is single, the side of the blade basin is affected by the channel vortex, the covering effect is poor, and the cooling effect is low; the cooling effect is unstable under the influence of various factors such as the stamping ratio, the injection angle, the aperture, the blowing ratio and the like; receive riser interference and the restriction of gas film hole processing angle on the flange, some positions are for example difficult trompil in axial afterbody region (as shown in fig. 2, difficult division gas film hole region 5), if the trompil then the same wall angle undersize in hole makes the processing qualification rate low, if do not trompil then with this regional unable cooling of flange, expose for a long time in high temperature gas, can lead to troubles such as ablation, chipping.
Accordingly, a technical solution is desired to overcome or at least alleviate at least one of the above-mentioned drawbacks of the prior art.
Disclosure of Invention
The invention aims to provide a cooling structure design method of a guide cooling blade flange plate of a high-pressure turbine, which aims to solve at least one problem in the prior art.
The technical scheme of the application is as follows:
a method for designing a cooling structure of a guide cooling blade flange plate of a high-pressure turbine comprises the following steps:
the method comprises the following steps that firstly, a flange plate is partitioned according to the air entraining position of the flange plate and the position of a vertical plate and is divided into a cold air introducing area and a non-cold air introducing area;
designing an impact and air film combined cooling structure in a cold air introducing area of the edge plate;
designing a flow disturbing column and air film combined cooling structure in a non-cold air introducing area of the edge plate;
designing an air film cooling structure on the gas side surface of the edge plate and the blade body of the edge plate;
and step five, arranging a variable-radius fillet on the edge plate and the blade body switching section along the blade profile.
Optionally, in step two, the designing of the impingement and air film combined cooling structure at the cold air introduction area of the platform comprises:
designing a thin plate with millimeter-magnitude thickness as an impact cover plate, wherein the axial distance of the impact cover plate is the axial distance between two vertical plates, a cavity is designed in the middle of the impact cover plate, and the shape of the surrounding edge of the cavity is the same as the outline of the protruding end of the blade body on the non-gas side of the edge plate; partitioning the impact cover plate according to the principle that the static pressure of the gas of the edge plate is similar, wherein all the areas are separated by rib cavities which are strip-shaped cavities and are arranged on the isostatic pressure line of the gas; an inner concave cavity is arranged in each region of the impact cover plate according to the shape and the area of each region; arranging a plurality of impact holes in each area of the impact cover plate;
partitioning the non-gas side surface corresponding to the cold air introduction area of the edge plate, wherein all the areas are separated by rib protrusions which are strip-shaped and are arranged on an isostatic pressure line of gas, and the positions and the shapes of the rib protrusions and rib cavities on the impact cover plate are the same; arranging a plurality of air film holes in each area of the non-gas side surface corresponding to the cold air introducing area of the edge plate, wherein the air film holes are formed in positions different from the bulges of the partition ribs;
installing an impact cover plate on the non-gas side of a cold air introducing area of the edge plate, fixing the front edge and the tail edge of the impact cover plate on the surfaces of the two vertical plates and the non-gas side surface of the edge plate during installation, clamping a cavity of the impact cover plate on the protruding end of the blade body, and enabling partition rib protrusions on the non-flow channel surface of the edge plate to penetrate through partition rib cavities; the cooling gas impacts the corresponding area of the non-gas-side surface of the rim plate through the impact holes of the impact cover plate, then is ejected out of the gas-side surface of the rim plate through the gas film holes, and cools the surface of the rim plate again through a gas film cooling manner.
Optionally, the inner concave cavity is a quadrilateral groove with an inner concave millimeter, and the area of the inner concave cavity is 1/4-1/2 of the corresponding subarea region.
Optionally, bosses are arranged on two circumferential sides of the non-gas side surface of the cold air introduction area of the flange plate and used for being tightly combined with the impact cover plate when being installed.
Optionally, in step three, the designing of the combined cooling structure of the spoiler and the air film at the non-cool air introducing area of the rim plate comprises:
designing a multi-discharge turbulence column in a non-cold air introducing area of the edge plate, wherein the surface of the bottom surface of the turbulence column is lower than the peripheral surface; a plurality of cold air bleed channels are formed in the vertical plate, inlets of the cold air bleed channels are formed in the surface of a cold air introducing area of the vertical plate, and the cold air bleed channels are used for introducing cold air from the cold air introducing area to a spoiler column area of the edge plate;
arranging a plurality of film holes at the non-cold air introducing area of the flange plate, wherein the film holes are opened to the gas side surface of the non-cold air introducing area of the flange plate;
designing a turbulence column cover plate, wherein the positions and the shapes of the turbulence column cover plate and the non-fuel gas side surface of the non-cold air introducing area of the edge plate are the same, and the turbulence column cover plate is arranged on the non-fuel gas side surface of the non-cold air introducing area of the edge plate;
the cooling gas flows through the turbulence column area from the cold air introduction area through the cold air bleed channel, cools the flange plate through turbulence heat exchange, and then is sprayed to the gas side of the flange plate through the gas film holes on the periphery side of the turbulence column, so that the gas side surface of the flange plate is subjected to gas film cooling.
Optionally, the turbulence columns are in the shape of cylinders, prisms or drop-shaped arranged in a row or a row.
Optionally, in step four, the designing of the film cooling structure on the gas-side surface of the edge plate and the blade body of the edge plate includes:
arranging a plurality of film holes on the gas side surface of the flange plate;
and a plurality of inclined air film holes are arranged at the front edge of the blade body of the near edge plate and the blade basin, and cold air is impacted to the front edge of the edge plate and the side of the blade basin from the cold air cavity of the blade body through the inclined air film holes to carry out air film cooling on the cold air.
Optionally, the gas film hole comprises, in a region where the gas-side surface of the rim plate is arranged: the vortex separation device comprises a leading edge area, a suction surface convex hull area, a pressure surface and pressure surface angle area, a main channel vortex separation line downstream area, a cascade throat downstream area and a trailing edge wake area.
Alternatively, the shape of the film holes arranged on the gas-side surface of the rim plate is circular or elliptical.
Optionally, in the fifth step, the radius-variable fillets arranged along the blade profile at the flange plate and the blade body transition section are 2-4mm, wherein the fillets at the front edge and the tail edge are large, and the middle position is in smooth transition.
The invention has at least the following beneficial technical effects:
the cooling structure design method of the high-pressure turbine guide cooling blade flange plate adopts a combined cooling mode to cool the wall surface of the high-pressure turbine guide blade flange plate, can reduce the hot spot temperature of the flange plate, improves the cooling effect of the blade flange plate and prolongs the service life of the blade.
Drawings
FIG. 1 is an angled schematic view of a prior art high pressure turbine guide cooling vane platform;
FIG. 2 is another angle schematic view of a prior art high pressure turbine guide cooling bucket platform;
FIG. 3 is a high pressure turbine guide vane platform area schematic view of an embodiment of the present application;
FIG. 4 is a schematic view of a high pressure turbine guide vane impingement plate configuration according to an embodiment of the present application;
FIG. 5 is a non-gas side surface schematic view of a high pressure turbine guide vane platform according to an embodiment of the present application;
FIG. 6 is a schematic illustration of a high pressure turbine guide vane platform cold gas introduction according to an embodiment of the present application;
FIG. 7 is a schematic illustration of high pressure turbine guide vane platform spoiler cooling according to an embodiment of the present application;
FIG. 8 is a schematic view of a high pressure turbine guide vane platform spoiler flow post configuration in accordance with an embodiment of the present application;
FIG. 9 is a schematic illustration of a high pressure turbine guide vane spoiler cover plate configuration according to an embodiment of the present application;
FIG. 10 is a schematic view of a high pressure turbine guide vane platform gas side surface film hole arrangement in accordance with an embodiment of the present application;
FIG. 11 is a schematic view of a high pressure turbine guide vane proximal plate airfoil canted film hole arrangement according to one embodiment of the present application.
Wherein:
1-gas inflow; 2-a flange plate; 3-leading edge side; 4-trailing edge side; 5-a region with difficult to open the air film hole; 6-standing the plate; 7-gas side; 8-non-gas side; 9-a cold gas introduction zone; 10-a non-cold gas introduction zone; 11-cold air inflow; 12-lobe ends; 13-outline of the blade body overhang; 14-impact cover plate; 15-surrounding edge; 16-rib cavities; 17-an internal concave cavity; 18-impingement holes; 19-air film hole; 20-rib protrusions; 21-a boss; 22-a cold air bleed air channel; 23-a turbulence column; 24-the surface of the bottom surface of the turbulence column; 25-a spoiler column cover plate; 26-a first gas film hole arrangement zone; 27-a second gas film hole arrangement zone; 28-body cooling air cavity; 29-near edge plate blade air film hole impact area; 30-the intersection line of the margin plate and the blade body.
Detailed Description
In order to make the implementation objects, technical solutions and advantages of the present application clearer, the technical solutions in the embodiments of the present application will be described in more detail below with reference to the drawings in the embodiments of the present application. In the drawings, the same or similar reference numerals denote the same or similar elements or elements having the same or similar functions throughout. The described embodiments are a subset of the embodiments in the present application and not all embodiments in the present application. The embodiments described below with reference to the drawings are exemplary and intended to be used for explaining the present application and should not be construed as limiting the present application. All other embodiments, which can be derived by a person skilled in the art from the embodiments given herein without making any creative effort, shall fall within the protection scope of the present application. Embodiments of the present application will be described in detail below with reference to the accompanying drawings.
In the description of the present application, it is to be understood that the terms "center", "longitudinal", "lateral", "front", "back", "left", "right", "vertical", "horizontal", "top", "bottom", "inner", "outer", and the like indicate orientations or positional relationships based on those shown in the drawings, and are used merely for convenience in describing the present application and for simplifying the description, and do not indicate or imply that the referenced device or element must have a particular orientation, be constructed in a particular orientation, and be operated, and therefore should not be construed as limiting the scope of the present application.
The present application will be described in further detail with reference to fig. 3 to 9.
The application provides a cooling structure design method of a guide cooling blade flange plate of a high-pressure turbine, which comprises the following steps:
dividing the edge plate 2 into a cold air introducing area 9 and a non-cold air introducing area 10 according to the air entraining position of the edge plate 2 and the position of the vertical plate 6;
designing an impact and air film combined cooling structure in a cold air introducing area 9 of the edge plate 2;
designing a flow disturbing column and air film combined cooling structure in a non-cold air introducing area 10 of the edge plate 2;
designing an air film cooling structure on the surface of the gas side 7 of the edge plate 2 and the blade body of the edge plate;
and step five, arranging a variable-radius fillet on the flange plate 2 and the blade body switching section along the blade profile.
The design method of the cooling structure of the high-pressure turbine guide cooling blade edge plate comprises the steps that firstly, as shown in fig. 3, the edge plate 2 is partitioned according to the air entraining position of the edge plate 2 and the positions of the vertical plates 6, the directions of the gas incoming flow 1 and the cold air incoming flow 11 are shown in the figure, the non-gas side 8 of the edge plate 2 is axially partitioned into three areas by the two vertical plates 6, the blades respectively correspond to the axial front part, the middle part and the tail part of the edge plate 2, the middle part area is a cold air introducing area 9, and the tail part area is a non-cold air introducing area 10.
The cooling structure design method of the high-pressure turbine guide cooling blade flange plate of the application designs the impact and air film combined cooling structure in the cold air introducing area 9 of the flange plate 2 and comprises the following steps:
the design of the impingement cooling structure is shown in fig. 4, firstly, a thin plate with millimeter-scale thickness for impingement cooling is designed as an impingement cover plate 14, the axial distance of the impingement cover plate 14 is the axial distance between two vertical plates 6, a cavity is designed in the middle of the impingement cover plate 14, and the shape of a surrounding edge 15 of the cavity is the same as that of an outer molded line 13 of a blade body protruding end of a non-gas side 8 of a flange plate 2; partitioning the impact cover plate 14 according to the principle that the static pressure of the gas of the edge plate is similar, wherein all the areas are separated by rib separating cavities 16, and the rib separating cavities 16 are strip-shaped cavities and are arranged on the isostatic pressure line of the gas; inner concave cavities 17 are arranged in each region of the impact cover plate 14 according to the shape and the area of each region; a plurality of impingement holes 18 are arranged in each region of the impingement cover plate 14;
in the preferred embodiment of the present application, the thickness of the impingement cover plate 14 is preferably 0.3mm, and after the impingement cover plate 14 is partitioned according to the principle that the static gas pressure of the edge plate is similar, an inner concave cavity 17 is provided in each region, and if the region is small in area and long and thin in shape, the inner concave cavity may not be provided. In this embodiment, the inner cavity 17 is a quadrilateral groove recessed by about several millimeters (0.5 mm in the example), and the area of the quadrilateral groove is about 1/4-1/2 of the corresponding subarea region, and the design of the inner cavity 17 can play a role in reducing the impact distance and increasing the strength of the cover plate.
The design of the air film cooling structure is as shown in fig. 5, the surface of the non-gas side 8 corresponding to the cold air introducing area 9 of the edge plate 2 is divided, all areas are separated by rib protrusions 20, the rib protrusions 20 are strip-shaped and are arranged on the isostatic pressure line of the gas, wherein the positions and the shapes of the rib protrusions 20 and the rib cavities 16 on the impact cover plate 14 are the same; a plurality of air film holes 19 are arranged in each area of the surface of the non-gas side 8 corresponding to the cold air introducing area 9 of the flange plate 2, wherein the air film holes 19 are arranged at positions different from the rib protrusions 20;
installing an impact cover plate 14 on a non-gas side 8 of a cold air introducing area 9 of the edge plate 2, fixing the front edge and the tail edge of the impact cover plate 14 on the surfaces of the two vertical plates 6 and the non-gas side 8 of the edge plate 2 during installation, clamping a cavity of the impact cover plate 14 on a blade body protruding end 12, and enabling a rib partition bulge 20 on a non-flow passage surface of the edge plate 2 to penetrate through a rib partition cavity 16; the cooling gas impinges on the corresponding area of the surface of the non-combustion side 8 of the platform 2 through impingement holes 18 of impingement cover plate 14 and then exits the surface of the combustion side 7 of the platform 2 through film holes 19, again cooling the platform 2 surface by film cooling.
In a preferred embodiment of the present application, the cold air introduction zone 9 of the rim plate 2 is provided with bosses 21 on both circumferential sides of the surface of the non-gas side 8 for tight engagement with the impingement cover plate 14 during installation.
The cooling structure design method of the high-pressure turbine guide cooling blade flange plate designs a flow disturbing column and air film combined cooling structure in a non-cold air introducing area 10 of the flange plate 2, and comprises the following steps:
as shown in fig. 6 to 7, the drain pins 23 are formed in the non-cold air introduction area 10 of the rim plate 2 such that the bottom surfaces of the drain pins 23 are located at a lower level than the peripheral surfaces; a plurality of cold air bleed air channels 22 are formed in the vertical plate 6, inlets of the cold air bleed air channels 22 are arranged on the surface of the cold air introducing area 9 of the vertical plate 6, and the cold air bleed air channels 22 are used for introducing cold air in the direction of the cold air incoming flow 11 from the cold air introducing area 9 to the area of the flow disturbing columns 23 of the edge plate 2;
in this embodiment, it is preferable to design a plurality of turbulence columns 23 in the non-cold air introduction region 10 having a longer axial length of the rim plate 2, the shape of the turbulence columns 23 may be any suitable form such as a cylinder, a prism, a water drop, etc., and the turbulence columns 23 may also be designed with a variable diameter and arranged in a forward or differential manner. Advantageously, cold air bleed air passage 22 should transition smoothly from the inlet to the outlet, and should not have sharp discontinuities to avoid excessive drag losses.
As shown in fig. 8, a plurality of film holes 19 are arranged in the non-cold air introduction area 10 of the rim plate 2, and the film holes 19 are opened to the surface of the gas side 7 of the non-cold air introduction area 10 of the rim plate 2;
advantageously, in this embodiment, the surface 24 of the bottom surface of the turbulator column is preferably designed to be lowered with respect to the surrounding surface, and a plurality of film holes 19 are arranged therein, so that the film holes 19 avoid interference with the turbulator column 23.
As shown in fig. 9, a spoiler cover 25 is designed, the spoiler cover 25 is identical in position and shape to the surface of the non-gas side 8 of the non-cold gas introduction area 10 of the rim plate 2, and is installed on the surface of the non-gas side 8 of the non-cold gas introduction area 10 of the rim plate 2; in this embodiment, the spoiler pillar cover plate 25 may be integrally formed with the rim plate 2, or may be separately formed and welded to the surface of the non-gas side 8 of the non-cold air introduction area 10 of the rim plate 2.
The cooling gas flows from the cold air introducing area 9 through the cold air bleed air channel 22 to the area of the turbulence column 23, cools the flange plate 2 through turbulence heat exchange, and then is sprayed to the gas side 7 of the flange plate 2 through the gas film holes 19 on the peripheral side of the turbulence column 23 to perform gas film cooling on the surface of the gas side 7 of the flange plate 2.
The design method of the cooling structure of the guide cooling blade edge plate of the high-pressure turbine comprises the following steps of designing the film cooling structure on the surface of the gas side 7 of the edge plate 2 and the blade body of the edge plate, wherein the film cooling structure comprises the following steps:
a plurality of gas film holes are arranged on the surface of the gas side 7 of the flange plate 2;
in this embodiment, as shown in fig. 10, the gas film hole emphasis arrangement region includes: the vortex separation device comprises a leading edge area, a suction surface convex hull area, a pressure surface and pressure surface angle area, a main channel vortex separation line downstream area, a cascade throat downstream area and a trailing edge wake area. Wherein, the openings of the film holes of the first film hole arrangement area 26 are arranged on the surface of the non-gas side 8 of the cold air introducing area 9, the openings of the film holes of the second film hole arrangement area 27 are arranged on the surface 24 of the bottom surface of the turbulence column, and the shapes of the film holes can be various appropriate shapes such as a circle, an ellipse and the like.
Further, as shown in fig. 11, a plurality of inclined film holes are arranged at the leading edge of the blade body of the near edge plate and the blade basin, and this area can be used as an impact area 29 of the blade body film holes of the near edge plate, and cold air is impacted to the leading edge of the edge plate and the side of the blade basin through the inclined film holes by a blade body cold air cavity 28, so that the cold air can carry out film cooling on the cold air.
In addition, according to the design method of the cooling structure of the guide cooling blade edge plate of the high-pressure turbine, the radius-variable fillet arranged on the edge plate and blade body switching section along the 30 blade profiles of the intersection line of the edge plate and the blade body is 2-4mm, wherein the fillet of the front edge and the fillet of the tail edge are larger, and the middle position of the fillet is in smooth transition, so that the loss of the end wall is reduced, and the strength is improved.
The cooling structure design method of the high-pressure turbine guide cooling blade flange plate divides the flange plate, and designs different combined cooling structures in each division, and adopts a combined cooling mode to cool the wall surface of the high-pressure turbine guide blade flange plate, so that the hot point temperature of the flange plate can be reduced, the cooling effect of the blade flange plate is improved, and the service life of the blade is prolonged. The cooling structure solves the defects that the cooling effect of the single air film cooling structure of the guide vane of the existing high-pressure turbine is poor, and the wall temperature is too high due to the fact that holes cannot be formed in part of the positions of the guide vane of the existing high-pressure turbine due to the reasons of processing and the like.
The above description is only for the specific embodiments of the present application, but the scope of the present application is not limited thereto, and any changes or substitutions that can be easily conceived by those skilled in the art within the technical scope of the present application should be covered within the scope of the present application. Therefore, the protection scope of the present application shall be subject to the protection scope of the claims.
Claims (10)
1. A method for designing a cooling structure of a guide cooling blade flange plate of a high-pressure turbine is characterized by comprising the following steps:
the method comprises the following steps that firstly, a flange plate is partitioned according to the air entraining position of the flange plate and the position of a vertical plate and is divided into a cold air introducing area and a non-cold air introducing area;
designing an impact and air film combined cooling structure in a cold air introducing area of the edge plate;
designing a flow disturbing column and air film combined cooling structure in a non-cold air introducing area of the edge plate;
designing an air film cooling structure on the gas side surface of the edge plate and the blade body of the edge plate;
and step five, arranging a variable-radius fillet on the edge plate and the blade body switching section along the blade profile.
2. The method as claimed in claim 1, wherein the step two of designing the impingement and film combined cooling structure at the cool air introduction area of the platform comprises:
designing a thin plate with millimeter-magnitude thickness as an impact cover plate, wherein the axial distance of the impact cover plate is the axial distance between two vertical plates, a cavity is designed in the middle of the impact cover plate, and the shape of the surrounding edge of the cavity is the same as the outline of the protruding end of the blade body on the non-gas side of the edge plate; partitioning the impact cover plate according to the principle that the static pressure of the gas of the edge plate is similar, wherein all the areas are separated by rib cavities which are strip-shaped cavities and are arranged on the isostatic pressure line of the gas; an inner concave cavity is arranged in each region of the impact cover plate according to the shape and the area of each region; arranging a plurality of impact holes in each area of the impact cover plate;
partitioning the non-gas side surface corresponding to the cold air introduction area of the edge plate, wherein all the areas are separated by rib protrusions which are strip-shaped and are arranged on an isostatic pressure line of gas, and the positions and the shapes of the rib protrusions and rib cavities on the impact cover plate are the same; arranging a plurality of air film holes in each area of the non-gas side surface corresponding to the cold air introducing area of the edge plate, wherein the air film holes are formed in positions different from the bulges of the partition ribs;
installing an impact cover plate on the non-gas side of a cold air introducing area of the edge plate, fixing the front edge and the tail edge of the impact cover plate on the surfaces of the two vertical plates and the non-gas side surface of the edge plate during installation, clamping a cavity of the impact cover plate on the protruding end of the blade body, and enabling partition rib protrusions on the non-flow channel surface of the edge plate to penetrate through partition rib cavities; the cooling gas impacts the corresponding area of the non-gas-side surface of the rim plate through the impact holes of the impact cover plate, then is ejected out of the gas-side surface of the rim plate through the gas film holes, and cools the surface of the rim plate again through a gas film cooling manner.
3. The method as claimed in claim 2, wherein the internal cavities are quadrilateral grooves with internal concavity in millimeter order, and the area of the internal cavities is 1/4-1/2 corresponding to the subarea region.
4. The method as claimed in claim 2, wherein the cold air introduction area of the platform is provided with bosses on both sides of the non-combustion side surface in the circumferential direction for tightly engaging with the impingement cover plate when installed.
5. The method as claimed in claim 2, wherein the step three is a step of designing a combined cooling structure of a spoiler and a film in the non-cool air introducing region of the platform, comprising:
designing a multi-discharge turbulence column in a non-cold air introducing area of the edge plate, wherein the surface of the bottom surface of the turbulence column is lower than the peripheral surface; a plurality of cold air bleed channels are formed in the vertical plate, inlets of the cold air bleed channels are formed in the surface of a cold air introducing area of the vertical plate, and the cold air bleed channels are used for introducing cold air from the cold air introducing area to a spoiler column area of the edge plate;
arranging a plurality of film holes at the non-cold air introducing area of the flange plate, wherein the film holes are opened to the gas side surface of the non-cold air introducing area of the flange plate;
designing a turbulence column cover plate, wherein the positions and the shapes of the turbulence column cover plate and the non-fuel gas side surface of the non-cold air introducing area of the edge plate are the same, and the turbulence column cover plate is arranged on the non-fuel gas side surface of the non-cold air introducing area of the edge plate;
the cooling gas flows through the turbulence column area from the cold air introduction area through the cold air bleed channel, cools the flange plate through turbulence heat exchange, and then is sprayed to the gas side of the flange plate through the gas film holes on the periphery side of the turbulence column, so that the gas side surface of the flange plate is subjected to gas film cooling.
6. The method as claimed in claim 5, wherein the turbulence column is in the form of a cylinder, a prism or a drop, and is arranged in a row or a row.
7. The method as claimed in claim 5, wherein the step four of designing the film cooling structure on the gas-side surface of the platform and the blade body of the near-platform includes:
arranging a plurality of film holes on the gas side surface of the flange plate;
and a plurality of inclined air film holes are arranged at the front edge of the blade body of the near edge plate and the blade basin, and cold air is impacted to the front edge of the edge plate and the side of the blade basin from the cold air cavity of the blade body through the inclined air film holes to carry out air film cooling on the cold air.
8. The method of designing a cooling structure of a platform for a guide cooling blade of a high pressure turbine as recited in claim 7, wherein said film holes are provided in a gas-side surface of the platform in a region including: the vortex separation device comprises a leading edge area, a suction surface convex hull area, a pressure surface and pressure surface angle area, a main channel vortex separation line downstream area, a cascade throat downstream area and a trailing edge wake area.
9. The method for designing a cooling structure of a platform for a guide cooling blade of a high pressure turbine as recited in claim 7, wherein the shape of the film hole provided in the gas-side surface of the platform is circular or elliptical.
10. The method as claimed in claim 1, wherein in step five, the radius of the fillet between the platform and the blade body transition section along the blade profile is 2-4mm, wherein the leading edge and the trailing edge have larger fillets and smooth transition at the middle.
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