CN116085062A - Turbine guide vane in aeroengine - Google Patents

Turbine guide vane in aeroengine Download PDF

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Publication number
CN116085062A
CN116085062A CN202211420411.7A CN202211420411A CN116085062A CN 116085062 A CN116085062 A CN 116085062A CN 202211420411 A CN202211420411 A CN 202211420411A CN 116085062 A CN116085062 A CN 116085062A
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CN
China
Prior art keywords
blade body
cavity
side wall
edge plate
guide pipe
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Pending
Application number
CN202211420411.7A
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Chinese (zh)
Inventor
宋伟
王维岩
吴法勇
丁勇峰
郑占一
王鹏
苏航
吴向宇
吴伟龙
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AECC Shenyang Engine Research Institute
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AECC Shenyang Engine Research Institute
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Filing date
Publication date
Application filed by AECC Shenyang Engine Research Institute filed Critical AECC Shenyang Engine Research Institute
Priority to CN202211420411.7A priority Critical patent/CN116085062A/en
Publication of CN116085062A publication Critical patent/CN116085062A/en
Pending legal-status Critical Current

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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
    • 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
    • F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
    • F01D25/005—Selecting particular materials
    • 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
    • F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
    • F01D25/08—Cooling; Heating; Heat-insulation
    • F01D25/12—Cooling
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2260/00—Function
    • F05D2260/20—Heat transfer, e.g. cooling
    • F05D2260/202—Heat transfer, e.g. cooling by film cooling
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2300/00—Materials; Properties thereof
    • F05D2300/20—Oxide or non-oxide ceramics
    • 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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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Materials Engineering (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)

Abstract

The application belongs to turbine guide vane design technical field in the aeroengine, concretely relates to turbine guide vane in aeroengine includes: the blade body is made of a ceramic matrix composite material and is of a cavity structure, and a plurality of air film holes are formed in the side wall of the blade body, wherein the part of each air film hole located in the pneumatic stagnation point area of the front edge of the blade body is inclined forwards, and the rest part of each air film hole is inclined backwards; the honeycomb duct is made of metal materials, a plurality of impact holes are formed in the side wall of the honeycomb duct, and an impact cavity is formed between the honeycomb duct and the side wall of the blade body.

Description

Turbine guide vane in aeroengine
Technical Field
The application belongs to the technical field of turbine guide vane design in aeroengines, and particularly relates to a turbine guide vane in an aeroengine.
Background
The method is a main way for improving the thrust-weight ratio of the aero-engine turbine by improving the inlet temperature and reducing the structural weight.
The turbine guide vane of the aeroengine is positioned at the inlet part of the turbine and bears higher temperature under the requirement of improving the thrust-weight ratio.
The ceramic matrix composite material has low density and excellent high temperature resistance, and the turbine guide vane manufactured by the ceramic matrix composite material can not only enhance the capability of bearing high temperature, but also reduce the quality, and can effectively improve the thrust-weight ratio of the aeroengine.
Although the turbine guide vane made of the ceramic matrix composite material can bear higher temperature, the turbine guide vane can be well suitable for most working conditions of the aero-engine, under certain extreme working conditions, the born temperature can exceed the bearing capacity of the turbine guide vane, the whole performance of the aero-engine is affected, even dangers occur, and the following two measures are mainly adopted at present:
1) The technical scheme is that the turbine guide vane manufactured by the ceramic matrix composite is designed into a cavity structure, a plurality of air film holes are formed in the side wall of the vane body, cooling air is introduced into the cavity when the aeroengine works, and is discharged through each air film hole, so that the turbine guide vane is cooled;
2) The turbine guide vane manufactured by the ceramic matrix composite is designed into a cavity structure, a plurality of air film holes are formed in the side wall of the vane body, a guide pipe is arranged in the cavity, an impact cavity is formed between the guide pipe and the side wall of the vane body, a plurality of impact holes are formed in the side wall of the guide pipe, a double-layer wall structure is formed, cooling air is introduced into the guide pipe when the aeroengine works, enters the impact cavity through each impact hole, is discharged through each air film hole after impact cooling is performed on the side wall of the vane body, the cooling efficiency is high, the number of the air film holes formed in the side wall of the vane body can be correspondingly reduced, only a small number of air film holes are formed in the front edge and the rear edge of the vane body, the damage to the integrity of fibers in the ceramic matrix composite is reduced, the strength and the toughness of the turbine guide vane are ensured, and the double-layer wall structure of the turbine guide vane manufactured by the ceramic matrix composite is difficult to process and manufacture and has high cost.
The present application has been made in view of the existence of the above-mentioned technical drawbacks.
It should be noted that the above disclosure of the background art is only for aiding in understanding the inventive concept and technical solution of the present invention, which is not necessarily prior art to the present application, and should not be used for evaluating the novelty and the creativity of the present application in the case where no clear evidence indicates that the above content has been disclosed at the filing date of the present application.
Disclosure of Invention
It is an object of the present application to provide a turbine guide vane in an aeroengine that overcomes or alleviates the technical drawbacks of at least one aspect of the known existence.
The technical scheme of the application is as follows:
a turbine guide vane in an aircraft engine comprising:
the blade body is made of a ceramic matrix composite material and is of a cavity structure, and a plurality of air film holes are formed in the side wall of the blade body, wherein the part of each air film hole located in the pneumatic stagnation point area of the front edge of the blade body is inclined forwards, and the rest part of each air film hole is inclined backwards;
the honeycomb duct is made of metal materials, a plurality of impact holes are formed in the side wall of the honeycomb duct, and an impact cavity is formed between the honeycomb duct and the side wall of the blade body.
According to at least one embodiment of the present application, in the turbine guide vane of the aeroengine, the distance between the partial air film holes located in the aerodynamic stagnation area of the front edge of the vane body is smaller.
According to at least one embodiment of the present application, in the turbine guide vane of the aeroengine, a part of the air film hole located at the vane basin side of the vane body front edge is cylindrical, and the distance is 5-6 Dq, wherein Dq is the aperture of the air film hole;
the part of the air film holes positioned on the back side of the blade body front edge blade are transited from the inside to the outside from a round shape to a trapezoid shape, and the interval is 7-10 Dq.
According to at least one embodiment of the present application, in the turbine guide vane in the aeroengine, in the part of the film hole on the back side of the vane body leading edge vane, the outlet divergence angle α=20 ° to 25 °, the outlet minimum width l2=1.5 to 2Dq, and the distance l1=0.6 to 0.8Dq from the outlet boundary to the inlet center.
According to at least one embodiment of the present application, in the turbine guide vane of the aeroengine, the side wall of the guide pipe is provided with a plurality of omega-shaped corrugated convex parts and a plurality of bulges;
the honeycomb duct of the convex part of the omega-shaped corrugation of each part extends axially and is propped against the side wall of the blade body;
each bulge is circumferentially distributed along the guide pipe and is abutted against the side wall of the blade body.
According to at least one embodiment of the present application, in the turbine guide vane of the aeroengine, the cavity is provided with a dividing rib, and the cavity is divided into a leading edge cavity and a trailing edge cavity; wherein, the front edge cavity is close to the front edge of the blade body, and an opening is formed at the root of the blade body; the rear edge cavity is close to the rear edge of the blade body, and an opening is formed at the tip of the blade body;
the honeycomb duct includes:
one end of the front guide pipe extends into the front edge cavity, and the outer wall of the other end of the front guide pipe is provided with a front guide pipe connecting edge;
one end of the rear guide pipe extends into the rear edge cavity, and the outer wall of the other end of the rear guide pipe is provided with a rear guide pipe connecting edge;
the turbine guide vane further includes:
the first bolt fastener is used for connecting the connecting edge of the front guide pipe to the root of the blade body;
and the second bolt fastener is used for connecting the connecting edge of the rear guide pipe with the tip part of the blade body.
According to at least one embodiment of the application, in the turbine guide vane of the aeroengine, the omega-shaped corrugated bulge part on the front guide pipe is one part and is abutted against the boundary between the front edge and the rear edge of the vane body;
the omega-shaped corrugated protruding parts on the rear guide pipe are provided with two parts, wherein one part is propped against the part of the side wall of the blade body close to the back edge blade back, and the other part is propped against the part of the side wall of the blade body close to the back edge blade basin.
According to at least one embodiment of the present application, in the turbine guide vane of the aeroengine, one end of the front guide pipe extending into the front edge cavity is blocked;
one end of the rear flow guide pipe extending into the trailing edge cavity is sealed.
According to at least one embodiment of the present application, in the turbine guide vane of the aeroengine, the method further includes:
the upper edge plate is connected to the tip of the blade body, an upper edge plate cooling cavity is formed in the upper edge plate cooling cavity, an upper edge plate air film hole is formed in the inner side wall of the upper edge plate cooling cavity, and an upper edge plate impact hole is formed in the outer wall of the upper edge plate cooling cavity; the upper edge plate cooling cavity surrounds the tip of the blade body;
the lower edge plate is connected to the root of the blade body, a lower edge plate cooling cavity is formed in the lower edge plate, a lower edge plate air film hole is formed in the inner side wall of the lower edge plate cooling cavity, and a lower edge plate impact hole is formed in the outer wall of the lower edge plate air film hole; the lower edge plate cooling cavity surrounds the root of the blade body.
According to at least one embodiment of the present application, in the turbine guide vane of the aeroengine, the upper edge plate is made of ceramic matrix composite material, the inner side wall of the upper edge plate is formed at the tip of the vane body and is in bonding connection with the outer side wall of the upper edge plate, and an upper edge plate cooling cavity is formed between the upper edge plate and the upper edge plate;
the lower edge plate is made of ceramic matrix composite material, the inner side wall of the lower edge plate is formed at the root of the blade body and is connected with the outer side wall of the blade body in an adhesive mode, and a cooling cavity of the lower edge plate is formed between the lower edge plate and the outer side wall of the blade body.
Drawings
FIG. 1 is a schematic view of a turbine guide vane in an aircraft engine provided in an embodiment of the present application;
FIG. 2 is a schematic view of an assembly of turbine guide vanes in an aircraft engine provided in an embodiment of the present application;
FIG. 3 is a partial cross-sectional view of a turbine guide vane in an aircraft engine provided in an embodiment of the present application;
FIG. 4 is a schematic illustration of a flow conduit provided in an embodiment of the present application;
FIG. 5 is yet another partial cross-sectional view of a turbine guide vane in an aircraft engine provided in an embodiment of the present application;
FIG. 6 is a schematic illustration of pressure distribution throughout a turbine stator blade airfoil sidewall during operation of an aircraft engine provided by an embodiment of the present application;
FIG. 7 is a schematic view of the distribution of heat transfer coefficients throughout the sidewall of a turbine guide vane blade body made of a ceramic matrix composite during operation of an aircraft engine provided by an embodiment of the present application;
FIG. 8 is a schematic view of a portion of a film hole provided in an embodiment of the present application on a back side of a blade leading edge blade;
wherein:
1-leaf body; 2-a flow guiding pipe; 3-a first bolt fastener; 4-a second bolt fastener; 5-an upper edge plate; 6-lower edge plate.
For the purpose of better illustrating the present embodiments, certain elements of the drawings may be omitted, enlarged or reduced and do not represent the actual product dimensions, and furthermore, the drawings are for illustrative purposes only and are not to be construed as limiting the present patent.
Detailed Description
In order to make the technical solution of the present application and the advantages thereof more apparent, the technical solution of the present application will be more fully described in detail below with reference to the accompanying drawings, it being understood that the specific embodiments described herein are only some of the embodiments of the present application, which are for explanation of the present application, not for limitation of the present application. It should be noted that, for convenience of description, only the portion relevant to the present application is shown in the drawings, and other relevant portions may refer to a general design, and without conflict, the embodiments and technical features in the embodiments may be combined with each other to obtain new embodiments.
Furthermore, unless defined otherwise, technical or scientific terms used in the description of this application should be given the ordinary meaning as understood by one of ordinary skill in the art to which this application belongs. The terms "upper," "lower," "left," "right," "center," "vertical," "horizontal," "inner," "outer," and the like as used in this description are merely used to indicate relative directions or positional relationships, and do not imply that a device or element must have a particular orientation, be configured and operated in a particular orientation, and that the relative positional relationships may be changed when the absolute position of the object being described is changed, and thus should not be construed as limiting the present application. The terms "first," "second," "third," and the like, as used in the description herein, are used for descriptive purposes only and are not to be construed as indicating or implying any particular importance to the various components. The use of the terms "a," "an," or "the" and similar referents in the description of the invention are not to be construed as limited in number to the precise location of at least one. As used in this description, the terms "comprises," "comprising," or the like are intended to cover an element or article that appears before the term and that is listed after the term and its equivalents, without excluding other elements or articles.
Furthermore, unless specifically stated and limited otherwise, the terms "mounted," "connected," and the like in the description herein are to be construed broadly and refer to either a fixed connection, a removable connection, or an integral connection, for example; can be mechanically or electrically connected; can be directly connected or indirectly connected through an intermediate medium, and can also be communicated with the inside of two elements, and the specific meaning of the two elements can be understood by a person skilled in the art according to specific situations.
The present application is described in further detail below in conjunction with fig. 1-7.
A turbine guide vane in an aircraft engine comprising:
the blade body 1 is made of a ceramic matrix composite material and is of a cavity structure, and a plurality of air film holes are formed in the side wall of the blade body, wherein the part of each air film hole located in the pneumatic stagnation point area of the front edge of the blade body is inclined forwards, and the rest part of each air film hole is inclined backwards;
the flow guiding pipe 2 is made of metal materials, a plurality of impact holes are formed in the side wall of the flow guiding pipe, and an impact cavity is formed between the flow guiding pipe and the side wall of the blade body 1.
For the turbine guide vane in the aeroengine disclosed in the above embodiment, it can be understood by those skilled in the art that the guide pipe 2 made of metal is designed to be arranged in the cavity of the vane body 1 made of ceramic matrix composite material, so as to form a double-wall structure, when the aeroengine works, cooling gas can be introduced into the guide pipe 2, and enters the impact cavity through each impact hole, and is discharged through each air film hole after the side wall of the vane body 1 is subjected to impact cooling, so that the cooling efficiency is higher, the number of air film holes formed in the side wall of the vane body 1 can be correspondingly reduced, further, the damage to the integrity of fibers in the ceramic matrix composite material can be greatly reduced, the strength and toughness of the turbine guide vane are ensured, and the guide pipe 2 arranged in the cavity of the vane body 1 made of ceramic matrix composite material is made of metal material, so as to facilitate processing, manufacturing and assembling, and is arranged in the cavity of the vane body 1 made of ceramic matrix composite material, and is not directly contacted with high-temperature gas at the turbine inlet of the aeroengine, so that the temperature is relatively low, and is not easy to be damaged by high-temperature, and the whole aeroengine can be designed to reduce the damage to fiber integrity, and improve the overall wall thickness.
In the operation process of the aeroengine, high-temperature gas can reside in the pneumatic stagnation point area of the front edge of the blade body of the turbine guide blade, and is positioned at the position, close to the front edge, of the She Shenshe basin to form a local high-temperature and high-pressure area.
In some alternative embodiments, in the turbine guide vane of the aeroengine, the distance between partial air film holes in the aerodynamic stagnation area of the front edge of the vane body is smaller, that is, the arrangement density is larger, so that a larger amount of cooling gas can be discharged in the area, and the cooling effect of the air film holes is ensured.
In some optional embodiments, in the turbine guide vane of the aeroengine, a part of the air film hole on the vane basin side of the vane body front edge is cylindrical, and the interval is 5-6 Dq, wherein Dq is the aperture of the air film hole;
the part of the air film holes positioned on the back side of the blade body front edge blade are transited from the inside to the outside from a round shape to a trapezoid shape, and the interval is 7-10 Dq.
In the operation process of the aeroengine, the temperature and pressure distribution of the side wall of the blade body of the turbine guide blade are uneven, as shown in fig. 6, the temperature distribution is approximately equal to the pressure distribution, and the heat exchange coefficient distribution of the side wall of the blade body of the turbine guide blade made of ceramic matrix composite material is uneven, as shown in fig. 7, in the turbine guide blade of the aeroengine disclosed by the embodiment, the spacing between the partial air film holes Q12/Q13/K1-K6/Q1 on the blade basin side of the front edge of the blade body is designed to be 5-6 Dq, the spacing between the partial air film holes Q2-Q6 on the back side of the blade body front edge of the blade body is designed to be 7-10 Dq, namely, the air film holes in the high-temperature, high-pressure and low-heat exchange coefficient areas are designed to have relatively large density, the air film holes in the low-temperature and low-pressure and high-heat exchange areas are designed to have relatively small density, the method can balance the discharge amount of cooling gas at each position, balance the cooling temperature of each position of the blade body, and cool each position of the blade body uniformly, in addition, the partial gas film holes Q12/Q13/K1-K6/Q1 positioned at the front edge blade basin side of the blade body are designed to be cylindrical, so that the damage to the ceramic matrix composite material in the partial region can be reduced, the partial gas film holes Q2-Q6 positioned at the back side of the front edge blade side of the blade body are designed to be transited from the inside to the outside to be trapezoid, namely, the effective flow area of the gas film holes is gradually enlarged, and the outlet part can cover a larger area, so that the cooling effect on the part of the blade body is improved, the density of the gas film holes can be reduced to a larger extent, and the strength of the blade body can be ensured as a whole.
In some alternative embodiments, in the turbine guide vane in the aeroengine, in the part of the film hole on the back side of the vane body leading edge vane, the outlet divergence angle α=20 ° to 25 °, specifically may be 22 °, the minimum outlet width l2=1.5 to 2Dq, and the distance l1=0.6 to 0.8Dq from the outlet boundary to the inlet center, so as to ensure the coverage characteristic of the part of the film hole for discharging the cooling gas on the outer wall surface of the vane body.
In some alternative embodiments, in the turbine guide vane of the aeroengine, the side wall of the guide pipe 2 is provided with a plurality of omega-shaped corrugated convex parts and a plurality of bulges;
the honeycomb duct 2 of the omega-shaped corrugated bulge part of each part extends axially and is propped against the side wall of the blade body 1;
each bulge is circumferentially distributed along the guide pipe 2 and is abutted against the side wall of the blade body 1.
The ceramic matrix composite material and the metal material have different thermal expansion coefficients, the thermal expansion coefficients are larger, under the high temperature condition, serious uncoordinated deformation is easy to occur between the blade body 1 made of the ceramic matrix composite material and the guide pipe 2 made of the metal material, and larger local stress is generated, so that damage occurs, in the turbine guide vane of the aeroengine disclosed by the embodiment, the omega-shaped corrugated protruding part which extends along the axial direction on the side wall of the guide pipe 2 is designed to abut against the side wall of the blade body 1, on one hand, the positioning of the guide pipe 2 in the cavity of the blade body 1 can be realized, the occurrence of flutter and damage are prevented, on the other hand, the omega-shaped corrugated protruding part has larger deformation capability, and the blade body 1 and the guide pipe 2 can be protected from being damaged by deformation and absorption of larger stress generated by uncoordinated deformation of the blade body 1 and the guide pipe 2 under the high temperature condition.
For the turbine guide vane in the aeroengine disclosed in the above embodiment, those skilled in the art can understand that the bulge distributed along the circumferential direction on the guide pipe 2 is designed to lean against the side wall of the blade body 1 and cooperate with the omega-shaped corrugated bulge at each place, so that the reliability of positioning the guide pipe 2 in the cavity of the blade body 1 can be enhanced, and the point between the guide pipe and the side wall of the blade body 1 is in local limited point contact, so that the turbulence of the cooling gas flowing in the impact cavity can be increased, the cooling effect can be enhanced, and the larger pressure loss caused to the cooling gas flowing can be avoided.
In some alternative embodiments, in the turbine guide vane of the aeroengine, the cavity is provided with a dividing rib, and the cavity is divided into a leading edge cavity and a trailing edge cavity; wherein, the front edge cavity is close to the front edge of the blade body 1, and an opening is formed at the root of the blade body 1; the rear edge cavity is close to the rear edge of the blade body 1, and an opening is formed at the tip of the blade body 1;
the draft tube 2 includes:
one end of the front guide pipe extends into the front edge cavity, and the outer wall of the other end of the front guide pipe is provided with a front guide pipe connecting edge;
one end of the rear guide pipe extends into the rear edge cavity, and the outer wall of the other end of the rear guide pipe is provided with a rear guide pipe connecting edge;
the turbine guide vane further includes:
the first bolt fastener 3 connects the connecting edge of the front guide pipe to the root of the blade body 1 and is far away from the high-temperature area of the inlet core of the turbine of the aeroengine, so that the high-temperature damage is avoided, and the high-temperature damage is avoided;
and a second bolt fastener 4 is used for connecting the connecting edge of the rear guide pipe to the tip part of the blade body 1.
In the turbine guide vane in the aeroengine disclosed in the above embodiment, during specific application, the cooling gas from the inner ring of the combustion chamber of the aeroengine may be introduced into the front guide pipe, the cooling gas from the inner ring of the combustion chamber of the aeroengine may enter into the front edge cavity through the impact hole on the side wall of the front guide pipe, impact cool the side wall of the front edge cavity, then flow out from the air film hole on the side wall of the front edge of the blade body 1, and the cooling gas from the outer ring of the combustion chamber of the aeroengine may be introduced into the rear guide pipe, the cooling gas from the outer ring of the combustion chamber of the aeroengine may enter into the rear edge cavity through the impact hole on the side wall of the rear guide pipe, impact cool the side wall of the rear edge cavity, and then flow out from the air film hole on the side wall of the rear edge of the blade body 1.
For the turbine guide vane in the aeroengine disclosed in the above embodiment, it may be further understood by those skilled in the art that the side wall of the vane body 1 in the turbine guide vane of the aeroengine has a larger pressure gradient along the chord direction, the pressure at the front edge of the vane body is far greater than the pressure at the rear edge of the vane body, the front edge cavity near the front edge of the vane body 1 and the front guide pipe thereof are designed to introduce the cooling gas in the inner ring of the combustion chamber, the higher pressure is provided, the cooling gas flowing out of the air film hole near the front edge of the vane body can be ensured to have a sufficient pressure margin, the cooling gas introduced into the outer ring of the combustion chamber by the rear edge cavity near the rear edge of the vane body 1 and the rear guide pipe thereof are designed to have a lower pressure, the cooling gas flowing out of the air film hole near the rear edge of the vane body can be ensured to have a relatively lower pressure, the pressure adapted to the rear edge of the vane body can be convenient for controlling the cooling gas flowing out of the air film hole in each chord direction region of the vane body 1, the temperature gradient along the chord direction can be further reduced, the vane body side wall is prevented from being caused to generate a larger non-uniform deformation, the vane body can be effectively utilized, and the local tip part of the vane body 1 can be prevented from weakening the local tip part of the vane body 1.
For the turbine guide vane in the aeroengine disclosed in the above embodiment, those skilled in the art can also understand that, in the design of the turbine guide vane, the front guide pipe is connected to the root of the vane body 1 through the front guide pipe connecting edge by the first bolt fastener 3, the rear guide pipe is connected to the tip of the vane body 1 through the rear guide pipe connecting edge by the second bolt fastener 4, and the first bolt fastener 3 and the second bolt fastener 4 can release the stress generated by the uncooled deformation of the vane body 1 made of the ceramic matrix composite material and the guide pipe 2 made of metal due to the high temperature by using the gaps between the first bolt fastener 3 and the corresponding bolt holes, so as to avoid the high temperature damage.
In some optional embodiments, in the turbine guide vane of the aeroengine, the omega-shaped corrugated bulge on the front guide pipe is one part, and the part abutting against the boundary between the front edge and the rear edge of the vane body is a part with larger temperature gradient;
two omega-shaped corrugated protruding parts on the rear guide pipe are arranged, wherein one part is abutted against the side wall of the blade body 1 close to the back edge blade, and the other part is abutted against the side wall of the blade body 1 close to the back edge blade basin, and is a position with larger temperature gradient.
For the turbine guide vane in the aeroengine disclosed in the above embodiment, those skilled in the art can understand that the vane body 1 and the guide pipe 2 are designed to contact with omega-shaped corrugated convex parts at a larger temperature gradient, that is, contact with the most serious part of uncoordinated deformation through the omega-shaped corrugated convex part line, so that the vane body 1 and the guide pipe 2 can be efficiently absorbed through deformation to generate larger stress due to uncoordinated deformation, direct contact of corresponding parts is avoided under a high temperature condition, and the vane body 1 and the guide pipe 2 are protected from damage.
For the turbine guide vane in the aeroengine disclosed by the embodiment, it can be further understood by those skilled in the art that one omega-shaped corrugated bulge on the front guide pipe is designed to be abutted against the front edge and the rear edge of the vane body, the front edge cavity in the vane body is divided at the part, the partition control of the cooling air discharge amount of the air film holes on two sides of the vane body can be facilitated, the vane body is uniformly cooled, the uneven deformation of the vane body is avoided, and when the aeroengine is applied specifically, two omega-shaped corrugated bulge on the front guide pipe can be designed, one part is abutted against the front edge and the rear edge of the vane body, and the other part is abutted against the part of the vane back close to the front edge.
In some alternative embodiments, in the turbine guide vane of the aeroengine, one end of the front guide pipe extending into the front edge cavity is blocked;
one end of the rear flow guide pipe extending into the trailing edge cavity is sealed.
In some alternative embodiments, the turbine guide vane in the aeroengine further comprises:
the upper edge plate 5 is connected to the tip of the blade body 1, an upper edge plate cooling cavity is formed in the upper edge plate, an upper edge plate air film hole is formed in the inner side wall of the upper edge plate cooling cavity, and an upper edge plate impact hole is formed in the outer wall of the upper edge plate air film hole; the upper edge plate cooling cavity surrounds the tip of the blade body 1;
the lower edge plate 6 is connected to the root of the blade body 1, a lower edge plate cooling cavity is formed in the lower edge plate, a lower edge plate air film hole is formed in the inner side wall of the lower edge plate cooling cavity, and a lower edge plate impact hole is formed in the outer wall of the lower edge plate air film hole; the lower edge plate cooling cavity surrounds the root of the blade body 1.
According to the turbine guide vane in the aeroengine disclosed by the embodiment, when the aeroengine works, cooling gas can be introduced into the upper edge plate cooling cavity through the upper edge plate impact holes, the cooling gas entering the upper edge plate cooling cavity can cool the inner side wall of the upper edge plate 5, and then flows out through the upper edge plate air film holes, so that the upper edge plate 5 can be prevented from being damaged by high temperature, the upper edge plate 5 has higher cooling efficiency, and the high-temperature effect of high-temperature gas at the turbine inlet of the aeroengine on the second bolt fastening piece 4 can be blocked.
For the design of the lower edge plate 6 in the turbine guide vane of the aeroengine disclosed in the above embodiment, reference is made to the above explanation of the design of the upper edge plate 5, and no further explanation is given here.
In some alternative embodiments, in the turbine guide vane of the aeroengine, the upper edge plate 5 is made of ceramic matrix composite material, the inner side wall of the upper edge plate is formed at the tip of the vane body 1 and is in bonding connection with the outer side wall of the vane body, and an upper edge plate cooling cavity is formed between the upper edge plate and the upper edge plate cooling cavity, so that the upper edge plate cooling cavity can be processed conveniently;
the lower edge plate 6 is made of ceramic matrix composite material, the inner side wall of the lower edge plate is formed at the root of the blade body 1 and is connected with the outer side wall of the blade body in an adhesive mode, and a lower edge plate cooling cavity is formed between the lower edge plate and the inner side wall, so that the lower edge plate cooling cavity can be conveniently processed.
In the description, each embodiment is described in a progressive manner, and each embodiment is mainly described by the differences from other embodiments, so that the same similar parts among the embodiments are mutually referred.
Having thus described the technical aspects of the present application with reference to the preferred embodiments illustrated in the accompanying drawings, it should be understood by those skilled in the art that the scope of the present application is not limited to the specific embodiments, and those skilled in the art may make equivalent changes or substitutions to the relevant technical features without departing from the principles of the present application, and those changes or substitutions will now fall within the scope of the present application.

Claims (10)

1. A turbine guide vane in an aircraft engine, comprising:
the blade body (1) is made of a ceramic matrix composite material and is of a cavity structure, and a plurality of air film holes are formed in the side wall of the blade body, wherein the part of each air film hole located in the pneumatic stagnation point area of the front edge of the blade body is inclined forwards, and the rest part of each air film hole is inclined backwards;
the honeycomb duct (2) is made of metal materials, a plurality of impact holes are formed in the side wall of the honeycomb duct, and an impact cavity is formed between the honeycomb duct and the side wall of the blade body (1) and is arranged in the cavity.
2. The aircraft engine turbine guide vane of claim 1,
the distance between partial air film holes in the pneumatic standing point area of the front edge of the blade body is smaller.
3. The aircraft engine turbine guide vane of claim 1,
the part of the air film holes positioned at the front edge of the blade body and at the side of the blade basin are cylindrical, the distance is 5-6 Dq, wherein Dq is the aperture of the air film holes;
the part of the air film holes positioned on the back side of the blade body front edge blade are transited from the inside to the outside from a round shape to a trapezoid shape, and the interval is 7-10 Dq.
4. A turbine guide vane in an aircraft engine according to claim 3,
in the part of the air film hole on the back side of the blade body front edge blade, the outlet expansion angle alpha=20-25 degrees, the outlet minimum width L2=1.5-2 Dq, and the distance L1=0.6-0.8 Dq from the outlet boundary to the inlet center.
5. A turbine guide vane in an aircraft engine according to claim 1, wherein,
the side wall of the flow guide pipe (2) is provided with a plurality of omega-shaped corrugated convex parts and a plurality of bulges;
the honeycomb duct (2) of the omega-shaped corrugated bulge part of each part extends axially and is propped against the side wall of the blade body (1);
each bulge is circumferentially distributed along the guide pipe (2) and is abutted against the side wall of the blade body (1).
6. The aircraft engine turbine guide vane of claim 1,
the cavity is provided with a dividing rib which divides the cavity into a front edge cavity and a rear edge cavity; wherein, the front edge cavity is close to the front edge of the blade body (1), and an opening is formed at the root of the blade body (1); the rear edge cavity is close to the rear edge of the blade body (1), and an opening is formed at the tip of the blade body (1);
the draft tube (2) comprises:
one end of the front guide pipe extends into the front edge cavity, and the outer wall of the other end of the front guide pipe is provided with a front guide pipe connecting edge;
one end of the rear guide pipe extends into the rear edge cavity, and the outer wall of the other end of the rear guide pipe is provided with a rear guide pipe connecting edge;
the turbine guide vane further includes:
a first bolt fastener (3) connects the connecting edge of the front guide pipe to the root of the blade body (1);
and the second bolt fastener (4) is used for connecting the connecting edge of the rear guide pipe with the tip part of the blade body (1).
7. A turbine guide vane in an aircraft engine according to claim 6, characterized in that,
the omega-shaped corrugated bulge part on the front flow guide pipe is provided with a part which is propped against the boundary of the front edge and the rear edge of the blade body;
two omega-shaped corrugated protruding parts are arranged on the rear guide pipe, wherein one part is propped against the side wall of the blade body (1) close to the back of the rear edge blade, and the other part is propped against the side wall of the blade body (1) close to the rear edge blade basin.
8. A turbine guide vane in an aircraft engine according to claim 6, characterized in that,
one end of the front flow guide pipe extending into the front edge cavity is blocked;
one end of the rear flow guide pipe extending into the trailing edge cavity is sealed.
9. A turbine guide vane in an aircraft engine according to claim 1, characterized in that,
further comprises:
the upper edge plate (5) is connected to the tip of the blade body (1), an upper edge plate cooling cavity is formed in the upper edge plate cooling cavity, an upper edge plate air film hole is formed in the inner side wall of the upper edge plate cooling cavity, and an upper edge plate impact hole is formed in the outer wall of the upper edge plate air film hole; the upper edge plate cooling cavity surrounds the tip of the blade body (1);
the lower edge plate (6) is connected to the root of the blade body (1), a lower edge plate cooling cavity is formed in the lower edge plate cooling cavity, lower edge plate air film holes are formed in the inner side wall of the lower edge plate cooling cavity, and lower edge plate impact holes are formed in the outer wall of the lower edge plate cooling cavity; the lower edge plate cooling cavity surrounds the root of the blade body (1).
10. The aircraft engine turbine guide vane of claim 9,
the upper edge plate (5) is made of ceramic matrix composite material, the inner side wall of the upper edge plate is formed at the tip part of the blade body (1) and is connected with the outer side wall of the blade body in an adhesive manner, and an upper edge plate cooling cavity is formed between the upper edge plate and the outer side wall;
the lower edge plate (6) is made of ceramic matrix composite material, the inner side wall of the lower edge plate is formed at the root of the blade body (1) and is connected with the outer side wall of the blade body in an adhesive mode, and a lower edge plate cooling cavity is formed between the lower edge plate and the outer side wall.
CN202211420411.7A 2022-11-15 2022-11-15 Turbine guide vane in aeroengine Pending CN116085062A (en)

Priority Applications (1)

Application Number Priority Date Filing Date Title
CN202211420411.7A CN116085062A (en) 2022-11-15 2022-11-15 Turbine guide vane in aeroengine

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
CN202211420411.7A CN116085062A (en) 2022-11-15 2022-11-15 Turbine guide vane in aeroengine

Publications (1)

Publication Number Publication Date
CN116085062A true CN116085062A (en) 2023-05-09

Family

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Family Applications (1)

Application Number Title Priority Date Filing Date
CN202211420411.7A Pending CN116085062A (en) 2022-11-15 2022-11-15 Turbine guide vane in aeroengine

Country Status (1)

Country Link
CN (1) CN116085062A (en)

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