CN113006879A - Aeroengine turbine film cooling hole with vortex generator - Google Patents
Aeroengine turbine film cooling hole with vortex generator Download PDFInfo
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- CN113006879A CN113006879A CN202110298612.3A CN202110298612A CN113006879A CN 113006879 A CN113006879 A CN 113006879A CN 202110298612 A CN202110298612 A CN 202110298612A CN 113006879 A CN113006879 A CN 113006879A
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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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- 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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- General Engineering & Computer Science (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
An aeroengine turbine film cooling hole with vortex generators is characterized in that the arc plate-shaped vortex generators are fixed in each film cooling hole, the front edges of the vortex generators face the airflow inlet of the film cooling hole, and the tail edges of the vortex generators face the airflow outlet of the film cooling hole. The vortex generator is additionally arranged on the inner wall of the air film cooling hole, so that the momentum of the cold air outlet is effectively reduced, the pneumatic performance is improved, the diffusion capacity of the cold air is increased, the film coverage rate of the cold air is improved by 18%, the sprayed cold air is not easy to separate from an attachment surface, the adhesion of the cold air is enhanced, the air film cooling efficiency is improved by 20%, the air film cooling mixing loss is reduced, and finally, the technical support is provided for improving the performance of the aero-engine.
Description
Technical Field
The invention relates to the technical field of turbines, in particular to an aeroengine turbine film cooling hole with a vortex generator, which is used for turbine blades and end wall regions.
Background
The thrust of the aircraft engine is greatly increased by the increase of the total temperature before the turbine, and the increase of the total temperature before the turbine becomes an important way for improving the performance of the engine and improving the economy of the engine, but the effective cooling technology is gradually paid attention to by the limitation of the heat resistance of the turbine blade material. The air film cooling technology has obvious advantages and good application prospect as a widely adopted cooling technology. The basic principle is that cooling air flows out through the air film cooling holes near the wall surface and covers the wall surface under the pressure and friction of the main flow to form a cold air film, so that the purposes of isolating the hot main flow and reducing the temperature of the turbine blade are achieved.
The initial study on film cooling technology was by Wiegahardt in Hot-Air Discharge for De-icing on a two-dimensional slot Hot-gas jet designed for wing freeze protection. With the advent of aircraft engines, it has been found that the cooling purpose can be achieved by injecting cool air to the hot parts. Primarily for use in combustion chambers initially. The film cooling technique has been increasingly used for turbine blade cooling until the seventies of the nineteenth century.
The air film cooling effect is mainly influenced by the following two factors: geometric parameters of the gas film cooling hole; and the pneumatic parameters of the air film cooling hole. At present, the most basic hole type is a cylindrical hole, and the cylindrical hole has the characteristics of simple processing and high structural strength, so that the most common application is realized. However, because the air film cooling effect of the cylindrical hole type is not good, a plurality of different hole types are developed continuously at home and abroad in recent years, so as to improve the cooling efficiency. Such as expanding holes, slotted holes, double jet holes, etc. In modern high performance aircraft engine turbine blade designs, the problem of cooling the turbine endwall is increasingly appreciated. The three-dimensional flow near the end wall, such as secondary flow phenomena like channel vortex and horseshoe vortex, makes film cooling of the end wall difficult to implement, so the development of the end wall film cooling technology is particularly important. Friedrichs, Hodsonhp, Daws et al studied the large-scale, low-speed turbine straight blade cascade cooling efficiency Distribution using ammonia-Diazo technology in Journal of turbomachery at Vol.118, No. 4 in 1996 using Distribution of film-coating effect on turbine end wall measured using the ammonia and azo technique, and found that the cooling efficiency increased in the downstream direction of the holes. In the study of ' experimental study on heat transfer of air film cooling of upstream end wall of front edge of turbine blade cascade ' in 3 rd edition of aviation dynamics ' of 2001, start-up and heat transfer experimental study on large-size low-speed turbine guide blade cascade with single-row and double-row hole cooling on upstream end wall of front edge are carried out by Liu Gao text, Liu pine age, Zhuhui people and the like of northwest university of industry, and it is found that the air film cooling of the end wall is influenced by secondary flow to a great extent, and the cooling effect is mainly determined by blowing ratio.
In the invention with publication number CN103244196, a discrete air film cooling hole type is disclosed, in which an expanding boss is arranged between the downstream of the air outlet edge of the cylindrical hole and the air outlet edge of the expanding structure, so as to improve the average air film cooling efficiency in the transverse direction.
In the invention and creation with the publication number of CN111042872A, a transverse expansion meridian contraction groove-shaped air film hole is disclosed, and the flow speed in the hole is improved and the far-downstream cooling effect is improved by combining a groove-shaped cross section with a meridian contraction structure.
In the existing hole type structure design method, the air film cooling efficiency is improved to different degrees, but the average air film coverage rate of most hole types is not greatly improved, and the improvement of the air film cooling effect is limited.
Disclosure of Invention
In order to overcome the defects that the average gas film coverage rate is low and the improvement of the gas film cooling effect is limited in the prior art, the invention provides a turbine gas film cooling hole of an aeroengine with a vortex generator.
In the invention, a plurality of blade air film cooling holes are arranged on each turbine blade and are respectively distributed on each turbine blade and the turbine end wall. The blade film cooling holes on the turbine blades are arranged along the spanwise direction of the turbine blades, and the blade film cooling holes on the turbine end wall are arranged along the circumference of the turbine end wall. The air film cooling holes are expanded air film cooling holes and cylindrical air film cooling holes.
The invention is characterized in that the blade film cooling holes on the turbine end wall are divided into a plurality of film cooling hole groups, and each film cooling hole group is respectively positioned between adjacent turbine blades; vortex generators are fixed in the air film cooling holes. The vortex generator is a circular arc-shaped plate. The radius of the arc is 20D-25D, and D is the diameter of the airflow inlet of the film cooling hole. The central angle a between the two ends of the arc length of the vortex generator is 3-5 degrees. The lower end face of the vortex generator is a cambered surface, and the radian of the lower end face is the same as that of the inner surface of the matched air film cooling hole.
One side edge of the vortex generator is the front edge of the vortex generator; the height of the front edge is 0.3D, and the other side edge of the vortex generator is the tail edge of the vortex generator; the height of the trailing edge was 0.5D. The thickness of the vortex generator is 0.5 mm.
When the vortex generator is arranged in the expanding type film cooling hole, the vortex generator is arranged on the inner lower surface of the expanding section, and the distance between the midpoint between the front edge and the tail edge of the vortex generator and the geometric center of the outlet of the film cooling hole is 0.25L1(ii) a The leading edge of the vortex generator faces the airflow inlet of the film cooling hole, and the trailing edge of the vortex generator faces the airflow outlet of the film cooling hole. And a tangent line of a midpoint between the front edge and the tail edge of the vortex generator and a longitudinal section of the geometric center of the film cooling hole of the expansion section form a line angle theta, wherein the theta is 45-65 degrees. The concave surface of the vortex generator corresponds to the outlet of the air film cooling hole.
When the vortex generator is arranged in the cylindrical air film cooling hole, the vortex generator is positioned on the inner lower surface of the cylindrical air film cooling hole, and the distance between the midpoint between the front edge and the tail edge of the vortex generator and the center of the outlet of the air film cooling hole is 0.25L2And directing a leading edge of the vortex generator toward an inlet end of the film-cooling hole and directing a trailing edge of the vortex generator toward an outlet end of the film-cooling hole. And a linear surface angle gamma formed between the tangent line of the arc midpoint of the vortex generator and the longitudinal section of the expansion section passing through the central line is 45-65 degrees. The concave surface of the vortex generator corresponds to the outlet of the air film cooling hole.
The expanding type film cooling hole is composed of an equal-diameter section and an expanding section, the orifice of the equal-diameter section is an airflow inlet, and the orifice of the expanding section is an airflow outlet. The equal-diameter section: the expansion segment is 1: 2. The expanding section is obtained by stretching the expanding point M of the equal diameter section along the chord direction of the turbine blade. The airflow inlet is circular, and the airflow outlet is strip-shaped. The expansion point M is far from the air filmThe distance between the outlets of the cooling holes is 2/3L1；L1The length of the connecting line from the inlet center to the geometric center of the outlet of the expanding hole is in mm. The included angle alpha between the inner surface of the expansion section elongated along the chord direction of the turbine blade and the geometric center line of the expansion section is 10-18 degrees.
The length of the cylindrical gas film cooling hole is L2The aperture is D; l is2: d is more than or equal to 3.0. And an included angle beta between the central line of the cylindrical gas film cooling hole and the inner surface and the lower surface of the cylindrical gas film cooling hole is an injection angle of the cylindrical gas film cooling hole, and the injection angle beta is 30-60 degrees.
The film cooling holes of the first end wall of the two groups of film cooling holes between the two turbine blades are arranged along the circumferential direction of the end wall; the center-to-center distance between adjacent endwall film cooling holes was 2.3D, and the center of the first endwall film cooling hole in the first group was 1.5D from the root of the suction side of the blade. The center of each endwall film cooling hole of the first group is located at-10% C of the endwallX～15％CXAnd at-10% CX～15％CXAny arrangement of (a); said C isXIs the axial chord length of the turbine blade. Second set of endwall film cooling hole distribution locations: when the second set of endwall film cooling holes are arranged, each of the second set of endwall film cooling holes is arranged along the pressure face of the turbine blade. The circle center of the first end wall air film cooling hole in the group is far from the air inlet end of the end wall by the axial chord length 25C of the turbine bladeX% and the center distance of each other adjacent end wall film cooling hole in the group is 10CX% of the total weight of the composition. And the distance from the circle center of each end wall film cooling hole in the second group to the pressure surface is randomly distributed within the range of 1.5-2D.
The number of the blade film cooling holes on the turbine blade is 15, and the number of the blade film cooling holes in each row is 18-20. The 1 st row to the 6 th row are distributed on a suction surface, the 7 th row to the 15 th row are distributed on a pressure surface, the 1 st row to the 6 th row are distributed on the suction surface, the 7 th row to the 15 th row are distributed on the pressure surface, and the airflow outlets of the blade air film cooling holes are respectively communicated with a cavity between the suction surface and the pressure surface of the turbine blade.
In the blade film cooling holes in each row of the suction surface, the centers of the 1 st row of the blade film cooling holes to the 6 th row of the blade film cooling holes are respectively positioned at 20 percent C of the axial chord length of the turbine bladeX，16％CX，12％CX，4％CX，2％CXAnd 1% of CXTo (3). The centers of the 7 th row blade film cooling holes and the 15 th row blade film cooling holes are respectively positioned at 1% C of the axial chord length of the turbine bladeX，3％CX，6％CX，10％CX，20％CX，30％CX，46％CX，56％CXAnd 75% of CXTo (3). The injection angles beta of the air film cooling holes in the 1 st row to the 15 th row on the turbine blade are 50.0 degrees, 43.5 degrees, 41.0 degrees, 90.0 degrees, 34.0 degrees, 30.0 degrees, 40.0 degrees, 37.0 degrees, 30.0 degrees and 34.0 degrees in sequence.
The centers of the blade film cooling holes in the rows from 1 to 4 and the centers of the blade film cooling holes in the rows from 10 to 15 are the same in the spanwise position of the turbine blade, wherein the center of the blade film cooling hole in the row from 1 is located at 4% of the spanwise position of the turbine blade, and the center distance of the rest adjacent blade film cooling holes is 4.7% of the spanwise position of the turbine blade.
The blade film cooling holes of the rows 5 to 9 are staggered, and the staggered distance is 2.4% of the spanwise direction of the turbine blade. The distances from the circle centers of the film cooling holes of the 1 st row of the blades in the 5 th row to the circle centers of the film cooling holes of the 1 st row of the 9 th row to the end wall are 6.4%, 4%, 6.4%, 4% and 6.4% of the spanwise direction of the turbine blade in sequence.
In the design of modern high-performance aircraft engines, an important way for improving the performance of the aircraft engine by further increasing the temperature of a turbine inlet is that an air film cooling technology is taken as a solution with remarkable effect, and the aim of the invention is to improve the cooling efficiency as much as possible. According to the invention, through the designed air film cooling hole, the aerodynamic parameters are improved, the air film coverage rate is increased, the air film cooling mixing loss is reduced, the cooling efficiency is improved, and finally, technical support is provided for improving the performance of the aircraft engine.
The invention is improved by adding vortex generator on the inner wall of hole on the basis of common cylindrical gas film cooling hole. The vortex generator can effectively reduce the momentum of the cold air outlet, increase the diffusion capacity of the cold air and improve the coverage rate of the cold air film. The evolution mode is to change the cylindrical hole into an expansion hole, namely, the outlet part of the cylindrical hole is expanded, so that the gas film coverage rate is improved, the sprayed cold air is not easy to separate from an attachment surface, the adhesion of the cold air is strengthened, and the gas film cooling efficiency is further improved.
The film cooling hole of the invention is composed of the film cooling hole and a built-in vortex generator. When cold air is sprayed from the film cooling hole, the momentum of the sprayed cold air is reduced by nearly 30 percent due to the existence of the vortex generators, and the loss of the cold air caused by the fact that the cold air is not attached to the surfaces of the turbine blades and the end walls due to the fact that the momentum is too large is reduced. The diffusion efficiency of the cold air is increased, and the momentum reduction helps the cold air to pass through the film cooling hole outlet and cover the turbine blade and the end wall surface with a film more quickly. The small momentum cold air has higher adhesiveness and is easier to adhere to the surfaces of the turbine blades and the end walls, and meanwhile, due to the reduction of the loss of the cold air, the cold air can cover wider surfaces, the average air film coverage rate is improved, the air film cooling efficiency is improved, and the effects of improving the average air film coverage rate by 18 percent and improving the air film cooling efficiency by 20 percent can be expected.
The vortex generator is small in size, the lower wall face is selected as the position of the vortex generator, the structure of the vortex generator can be well protected from being affected by high temperature under the wrapping of the sprayed cold air, specifically, the vortex generator can be well protected from being deformed due to the influence of high temperature when being applied to the inside of the air film cooling hole, and the pneumatic performance change caused by the change of the structure is avoided.
When the air film cooling hole is a common cylindrical air film cooling hole, the position arrangement of the air film cooling hole requires that cold air can fully cover the wall surface to be protected, and meanwhile, the cold air is controlled not to cause the gas flow to be separated from the blades, so that the loss is increased.
When the air film cooling hole is the expanding air film cooling hole, the expanding hole is adopted as the air film cooling hole, the momentum of the sprayed cold air can be reduced, the air film coverage rate is improved, the vortex generator is matched with the expanding hole, and the coverage rate of the air film can be further improved and the air film cooling efficiency is improved through the coupling effect of the vortex generator and the expanding hole. The arrangement of the expanding holes ensures the structural strength and effectively improves the cooling efficiency. FIG. 11 is a comparison graph of the transverse cooling efficiency of a prior art expanding film cooling hole and a film cooling hole for a turbine blade/end wall for the downstream of the expanding hole of the present invention, and the comparison shows that the cooling efficiency near the centerline of the expanding film cooling hole with a vortex generator is significantly improved, which illustrates that the vortex generator enhances the transverse cooling capability, which is also the reason for the improved transverse average cooling efficiency of the expanding film cooling hole with a vortex generator.
The vortex generators are additionally arranged on the inner wall of the air film cooling hole, so that the momentum of a cold air outlet is effectively reduced, the diffusion capacity of cold air is improved, and the coverage rate of a cold air film is improved. The evolution mode is to change the cylindrical hole into an expansion hole, namely, the outlet part of the cylindrical hole is expanded, so that the gas film coverage rate is improved, the sprayed cold air is not easy to separate from an attachment surface, the adhesion of the cold air is strengthened, and the gas film cooling efficiency is further improved.
When the cold air is sprayed from the film cooling hole, the momentum of the sprayed cold air is effectively reduced due to the existence of the vortex generator, and the loss caused by the fact that the cold air cannot be attached to the surfaces of the turbine blade and the end wall due to overlarge momentum is reduced. The diffusion efficiency of the cold air is increased, and the momentum reduction helps the cold air to pass through the film cooling hole outlet and cover the turbine blade and the end wall surface with a film more quickly. The small momentum cold air has higher adhesiveness and is easier to adhere to the surfaces of the turbine blades and the end walls, and meanwhile, due to the reduction of the loss of the cold air, the cold air can cover wider surfaces, the average air film coverage rate is improved, and the air film cooling efficiency is improved, so that the effects of improving the average air film coverage rate by 18 percent and improving the air film cooling efficiency by 20 percent are expected to be achieved.
The vortex generator is small in size, the lower wall face is selected as the position of the vortex generator, the structure of the vortex generator can be well protected from being affected by high temperature under the wrapping of the sprayed cold air, specifically, the vortex generator can be well protected from being deformed due to the influence of high temperature in the application of the interior of the air film cooling hole, and the change of the structure and the change of the pneumatic performance are avoided.
When the air film cooling hole is a cylindrical air film cooling hole, the position arrangement of the air film cooling hole ensures that cold air can fully cover the wall surface to be protected, and meanwhile, the loss of the cold air is controlled to be minimum.
When the air film cooling hole is the expanding air film cooling hole, the expanding hole is adopted as the air film cooling hole, the momentum of the sprayed cold air can be reduced, the air film coverage rate is improved, the vortex generator is matched with the expanding hole, and the coverage rate of the air film can be further improved and the air film cooling efficiency is improved through the coupling effect of the vortex generator and the expanding hole. The arrangement of the expanding holes ensures the structural strength and effectively improves the cooling efficiency.
Drawings
FIG. 1 is a three-dimensional view of a vortex generator of the present invention;
FIG. 2 is a top view of a conventional cylindrical film cooling hole;
FIG. 3 is a cross-sectional view of a conventional cylindrical film cooling hole;
FIG. 4 is a top view of the film cooling hole for the turbine endwall proposed by the present invention;
FIG. 5 is a cross-sectional view of FIG. 4;
FIG. 6 is a top view of a film cooling hole for an expanding hole as set forth in the present invention;
FIG. 7a is a cross-sectional view along the hole gas flow direction when the film cooling hole of the present invention is used in an expanding hole;
FIG. 7b is a top view of the film cooling hole of the present invention in the direction of the hole flow when used in an expanding hole;
FIG. 7c is a partial enlarged view of portion A of FIG. 7 a;
FIG. 8 is a schematic view of the arrangement of film cooling holes in the turbine endwall according to the present invention;
FIG. 9 is a schematic illustration of the present invention distributed over a turbine blade;
FIG. 10 is a three-dimensional view of the present invention on a turbine inlet guide vane row;
FIG. 11 is a comparison graph of prior art expanding film cooling holes and film cooling holes for turbine blades/endwalls of the present invention for transverse cooling efficiency downstream of the expanding holes.
In the figure: 1. an inlet guide vane row; 2. a vortex generator; 3. cylindrical gas film cooling holes; 4. an expanding film cooling hole; 5. the expanding film cooling holes of the prior art; 6. an expanding film cooling hole applied to the vortex generator; 7. a leading edge; 8. a trailing edge.
Detailed Description
The embodiment is an air film cooling hole on a guide vane row at the inlet of a turbine; the film cooling holes are divided into blade film cooling holes and end wall film cooling holes according to positions. The blade film cooling holes have the same structural characteristics as the endwall film cooling holes.
The guide vane row comprises a turbine end wall and a plurality of turbine vanes fixed on the outer circumference of the turbine end wall. The blade air film cooling holes are formed in the turbine blades and are distributed along the spanwise direction of the turbine blades. A plurality of groups of film cooling hole groups distributed along the circumferential direction are distributed on the end wall of the turbine, and each film cooling hole group is respectively positioned between adjacent turbine blades; each of the film cooling hole groups includes a plurality of endwall film cooling holes. The guide vane row is in the prior art, and the arrangement mode of the air film cooling holes on the end wall is the same as that on the turbine blade.
When the film cooling hole is an expanding film cooling hole 4, the expanding film cooling hole 4 is composed of an equal-diameter section and an expanding section, and the orifice of the equal-diameter section is an airflow inlet and the orifice of the expanding section is an airflow outlet. The equal-diameter section: the expansion segment is 1: 2. The expanding section is obtained by stretching the expanding point M of the equal diameter section along the chord direction of the turbine blade. The airflow inlet is circular, and the airflow outlet is strip-shaped. The distance between the expansion point M and the outlet of the film cooling hole is 2/3L1；L1The length of the connecting line from the inlet center to the geometric center of the outlet of the expanding hole is in mm. The turbine is arranged along the middle edge of the expansion sectionThe included angle alpha between the chord direction elongated inner surface of the blade and the geometric center line of the expansion section is 10-18 degrees. In the present embodiment, α is 14 °
The cylindrical film cooling hole 3 is a cylindrical film cooling hole 3 in the prior art. The length of the cylindrical gas film cooling hole 3 is L2The aperture is D; l is2: d is more than or equal to 3.0. An included angle beta between the central line of the cylindrical air film cooling hole and the inner surface and the lower surface of the cylindrical air film cooling hole is an injection angle of the cylindrical air film cooling hole, and the injection angle beta is 30-60 degrees in the embodiment, the injection angle beta is 30 degrees
The film cooling hole adopts an expansion type film cooling hole 4 or a cylindrical film cooling hole 3 in the prior art. And a vortex generator 2 is fixed in the air film cooling hole.
The vortex generator 2 is a circular arc-shaped plate. The radius of the arc is 20D-25D, and D is the diameter of the airflow inlet of the film cooling hole. The central angle a between the two ends of the arc length of the vortex generator 2 is 3-5 degrees. The lower end face of the vortex generator 2 is a cambered surface, and the radian of the lower end face is the same as that of the inner surface of the matched air film cooling hole. One side of the vortex generator 2 is the leading edge 7 of the vortex generator 2; the height of the leading edge is 0.3D, and the other side edge of the vortex generator 2 is the tail 8 edge of the vortex generator 2; the height of the trailing edge was 0.5D. The thickness of the vortex generator 2 is 0.5 mm.
When the vortex generator 2 is arranged in the expanding type film cooling hole 4, the vortex generator 2 is arranged on the inner lower surface of the expanding section, and the distance between the middle point between the front edge and the tail edge of the vortex generator 2 and the geometric center of the outlet of the film cooling hole is 0.25L1。
When installed, the leading edge 7 of the vortex generator 2 is directed towards the airflow inlet of the film-cooling hole and the trailing edge 8 of the vortex generator 2 is directed towards the airflow outlet of the film-cooling hole. And a tangent line of a midpoint between the front edge and the tail edge of the vortex generator 2 and a longitudinal section of the geometric center of the film cooling hole of the expansion section form a line angle theta, wherein the theta is 45-65 degrees. The concave surface of the vortex generator 2 corresponds to the outlet of the film cooling hole. In the present embodiment, θ is 45 °
When the vortex generator 2 is installed in the cylindrical film cooling hole 3, the vortex generator 2 is positioned on the inner and lower surfaces of the cylindrical film cooling hole, and the distance between the midpoint between the front edge and the tail edge of the vortex generator 2 and the center of the outlet of the film cooling hole is 0.25L2. When installed, the leading edge 7 of the vortex generator 2 is directed towards the inlet end of the film-cooling hole and the trailing edge 8 of the vortex generator 2 is directed towards the outlet end of the film-cooling hole. And a line surface angle gamma of 45-65 degrees is formed between the tangent line of the arc midpoint of the vortex generator 2 and the longitudinal section of the expansion section passing through the center line. The concave surface of the vortex generator 2 corresponds to the outlet of the film cooling hole. In the present embodiment, γ is 45 °.
The film cooling holes are respectively arranged on the end wall and the turbine blade.
The film cooling holes distributed on the end wall are positioned between two adjacent turbine blades and are end wall film cooling holes. Two groups of end wall film cooling holes are formed between every two adjacent turbine blades, and the first group of end wall film cooling holes are arranged along the circumferential direction of the end wall; the center-to-center distance between adjacent endwall film cooling holes was 2.3D, and the center of the first endwall film cooling hole in the first group was 1.5D from the root of the suction side of the blade. The center of each endwall film cooling hole of the first group is located at-10% C of the endwallX～15％CXAnd at-10% CX～15％CXAny arrangement of (a); said C isXIs the axial chord length of the turbine blade. In this embodiment, there are 5 endwall film cooling holes in the first set, each endwall film cooling hole centered at-10% C of the endwallX，3％CX，7％CX，11％CX，15％CX。
Second set of endwall film cooling hole distribution locations: when the second set of endwall film cooling holes are arranged, each of the second set of endwall film cooling holes is arranged along the pressure face of the turbine blade. The circle center of the first end wall air film cooling hole in the group is far from the air inlet end of the end wall by the axial chord length 25C of the turbine bladeX% of the other phases of the groupThe center distances of the air film cooling holes of the adjacent end walls are all 10CX% of the total weight of the composition. And the distance from the circle center of each end wall film cooling hole in the second group to the pressure surface is randomly distributed within the range of 1.5-2D.
In this embodiment, the number of the second group of end wall film cooling holes is 5, and the circle center position of each end wall film cooling hole is 25% C of the axial chord length of the turbine blade in sequenceX，35％CX，45％CX，55％CX，65％CX. The distances from the circle center positions of the end wall air film cooling holes to the pressure surface are 1.5D, 1.7D, 1.8D, 1.9D and 2D respectively.
The profile of the turbine blade is known from the prior art. The axial chord length of the turbine blade is CX(ii) a The leading edge point of the turbine blade is 0% CXThe trailing edge point of the turbine blade is 100% CX。
The turbine blade is provided with 15 rows of blade film cooling holes, and the number of the blade film cooling holes in each row is 18-20. The 1 st row to the 6 th row are distributed on a suction surface, the 7 th row to the 15 th row are distributed on a pressure surface, the 1 st row to the 6 th row are distributed on the suction surface, the 7 th row to the 15 th row are distributed on the pressure surface, and the airflow outlets of the blade air film cooling holes are respectively communicated with a cavity between the suction surface and the pressure surface of the turbine blade.
In the blade film cooling holes in each row of the suction surface, the centers of the 1 st row of the blade film cooling holes to the 6 th row of the blade film cooling holes are respectively positioned at 20 percent C of the axial chord length of the turbine bladeX，16％CX，12％CX，4％CX，2％CXAnd 1% of CXTo (3). The centers of the 7 th row blade film cooling holes and the 15 th row blade film cooling holes are respectively positioned at 1% C of the axial chord length of the turbine bladeX，3％CX，6％CX，10％CX，20％CX，30％CX，46％CX，56％CXAnd 75% of CXTo (3).
The centers of the blade film cooling holes in the rows from 1 to 4 and the centers of the blade film cooling holes in the rows from 10 to 15 are the same in the spanwise position of the turbine blade, wherein the center of the blade film cooling hole in the row from 1 is located at 4% of the spanwise position of the turbine blade, and the center distance of the rest adjacent blade film cooling holes is 4.7% of the spanwise position of the turbine blade.
The blade film cooling holes of the rows 5 to 9 are staggered, and the staggered distance is 2.4% of the spanwise direction of the turbine blade. The distances from the circle centers of the film cooling holes of the 1 st row of the blades in the 5 th row to the circle centers of the film cooling holes of the 1 st row of the 9 th row to the end wall are 6.4%, 4%, 6.4%, 4% and 6.4% of the spanwise direction of the turbine blade in sequence.
The injection angles beta of the air film cooling holes in the 1 st row to the 15 th row on the turbine blade are 50.0 degrees, 43.5 degrees, 41.0 degrees, 90.0 degrees, 34.0 degrees, 30.0 degrees, 40.0 degrees, 37.0 degrees, 30.0 degrees and 34.0 degrees in sequence.
Claims (10)
1. The aeroengine turbine film cooling hole with the vortex generator is characterized in that a plurality of blade film cooling holes are formed in each turbine blade and are respectively distributed on the end wall of each turbine blade on the turbine end wall; the blade film cooling holes on the turbine blades are arranged along the spanwise direction of the turbine blades, and the blade film cooling holes on the turbine end wall are arranged along the circumference of the turbine end wall; the air film cooling holes are an expansion air film cooling hole and a cylindrical air film cooling hole; the blade film cooling holes on the turbine end wall are divided into a plurality of film cooling hole groups, and each film cooling hole group is respectively positioned between adjacent turbine blades; vortex generators are fixed in the air film cooling holes; the vortex generator is a circular arc-shaped plate; the radius of the arc is 20D-25D, and D is the diameter of the airflow inlet of the film cooling hole; the central angle a between the two ends of the arc length of the vortex generator is 3-5 degrees; the lower end face of the vortex generator is a cambered surface, and the radian of the lower end face is the same as that of the inner surface of the matched air film cooling hole.
2. The aircraft engine turbine film cooling hole having a vortex generator as claimed in claim 1, wherein one side of said vortex generator is a leading edge of said vortex generator; the height of the front edge is 0.3D, and the other side edge of the vortex generator is the tail edge of the vortex generator; the height of the trailing edge is 0.5D; the thickness of the vortex generator is 0.5 mm.
3. The aircraft engine turbine film cooling hole having a vortex generator as claimed in claim 1, wherein when said vortex generator is installed in an extended film cooling hole, the vortex generator is installed on the inner lower surface of the extended section such that the distance between the midpoint between the leading edge and the trailing edge of the vortex generator and the geometric center of the exit of the film cooling hole is 0.25L1(ii) a The front edge of the vortex generator faces to the airflow inlet of the film cooling hole, and the tail edge of the vortex generator faces to the airflow outlet of the film cooling hole; a tangent line of a midpoint between the front edge and the tail edge of the vortex generator and a longitudinal section of the geometric center of the film cooling hole of the expansion section form a line angle theta, wherein the theta is 45-65 degrees; the concave surface of the vortex generator corresponds to the outlet of the air film cooling hole.
4. The aircraft engine turbine film cooling hole having vortex generators of claim 1, wherein said vortex generators are located on the lower surface of the cylindrical film cooling hole when said vortex generators are installed in the cylindrical film cooling hole, and the distance between the midpoint between the leading edge and the trailing edge of the vortex generators and the center of the exit of the film cooling hole is 0.25L2And the leading edge of the vortex generator faces the inlet end of the film cooling hole, and the trailing edge of the vortex generator faces the outlet end of the film cooling hole; a linear surface angle gamma of 45-65 degrees is formed between the tangent line of the arc midpoint of the vortex generator and the longitudinal section of the expansion section passing through the center line; the concave surface of the vortex generator corresponds to the outlet of the air film cooling hole.
5. The aircraft engine turbine film cooling hole with vortex generators as claimed in claim 1, wherein the expanding film cooling hole is composed of a constant diameter section and an expanding section, and the orifice of the constant diameter section is an air flow inlet and the orifice of the expanding section is an air flow inletAn airflow outlet; the equal-diameter section: 1:2 of expansion segment; the expansion section is obtained by extending the expansion point M of the equal-diameter section along the chord direction of the turbine blade; the airflow inlet is circular, and the airflow outlet is strip-shaped; the distance between the expansion point M and the outlet of the film cooling hole is 2/3L1；L1The length of a connecting line between the inlet center and the outlet geometric center of the expanding hole is in mm; the included angle alpha between the inner surface of the expansion section elongated along the chord direction of the turbine blade and the geometric center line of the expansion section is 10-18 degrees.
6. The aircraft engine turbine film cooling hole with swirl generators of claim 1, wherein the cylindrical film cooling hole has a length L2The aperture is D; l is2: d is more than or equal to 3.0; and an included angle beta between the central line of the cylindrical gas film cooling hole and the inner surface and the lower surface of the cylindrical gas film cooling hole is an injection angle of the cylindrical gas film cooling hole, and the injection angle beta is 30-60 degrees.
7. The aircraft engine turbine film cooling hole with a vortex generator as claimed in claim 1, wherein the film cooling holes of the first end wall of the two groups of film cooling holes between the two turbine blades are arranged along the circumferential direction of the end wall; the center-to-center distance between each adjacent endwall film cooling hole is 2.3D, and the center of the first endwall film cooling hole in the first group is 1.5D away from the root of the suction surface of the blade; the center of each endwall film cooling hole of the first group is located at-10% C of the endwallX～15％CXAnd at-10% CX～15％CXAny arrangement of (a); said C isXIs the axial chord length of the turbine blade; second set of endwall film cooling hole distribution locations: when the second set of endwall film cooling holes are arranged, arranging each of the second set of endwall film cooling holes along the pressure face of the turbine blade; the circle center of the first end wall air film cooling hole in the group is far from the air inlet end of the end wall by the axial chord length 25C of the turbine bladeX% and the center distance of each other adjacent end wall film cooling hole in the group is 10CXPercent; and the distance from the circle center of each end wall film cooling hole in the second group to the pressure surface is randomly distributed within the range of 1.5-2D.
8. The aero engine turbine film cooling hole with vortex generators of claim 1 wherein the number of the blade film cooling holes on the turbine blade is 15, and the number of the blade film cooling holes in each row is the same, and is 18 to 20; the 1 st row to the 6 th row are distributed on a suction surface, the 7 th row to the 15 th row are distributed on a pressure surface, the 1 st row to the 6 th row are distributed on the suction surface, the 7 th row to the 15 th row are distributed on the pressure surface, and the airflow outlets of the blade air film cooling holes are respectively communicated with a cavity between the suction surface and the pressure surface of the turbine blade.
9. The aircraft engine turbine film cooling hole with vortex generators of claim 8, wherein the centers of the 1 st row blade film cooling holes to the 6 th row blade film cooling holes in each row of blade film cooling holes on the suction surface are respectively located at 20% C of the axial chord length of the turbine bladeX，16％CX，12％CX，4％CX，2％CXAnd 1% of CXAt least one of (1) and (b); the centers of the 7 th row blade film cooling holes and the 15 th row blade film cooling holes are respectively positioned at 1% C of the axial chord length of the turbine bladeX，3％CX，6％CX，10％CX，20％CX，30％CX，46％CX，56％CXAnd 75% of CXAt least one of (1) and (b); the injection angles beta of the air film cooling holes in the 1 st row to the 15 th row on the turbine blade are 50.0 degrees, 43.5 degrees, 41.0 degrees, 90.0 degrees, 34.0 degrees, 30.0 degrees, 40.0 degrees, 37.0 degrees, 30.0 degrees and 34.0 degrees in sequence.
10. The aircraft engine turbine film cooling hole with a vortex generator as claimed in claim 8, wherein the centers of the rows of blade film cooling holes in the rows 1 to 4 and the centers of the rows of blade film cooling holes in the rows 10 to 15 are the same in the spanwise position of the turbine blade, wherein the center of the row 1 blade film cooling hole is located at 4% of the spanwise direction of the turbine blade, and the center distance of the other adjacent blade film cooling holes is 4.7% of the spanwise direction of the turbine blade;
the blade air film cooling holes in the rows from 5 th row to 9 th row are staggered, so that the staggered distance is 2.4% of the spanwise direction of the turbine blade; the distances from the circle centers of the film cooling holes of the 1 st row of the blades in the 5 th row to the circle centers of the film cooling holes of the 1 st row of the 9 th row to the end wall are 6.4%, 4%, 6.4%, 4% and 6.4% of the spanwise direction of the turbine blade in sequence.
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| CN113006879B (en) | 2023-06-23 |
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