CN115163564B - Aeroengine blade tip position machine casket fluting structure - Google Patents

Aeroengine blade tip position machine casket fluting structure Download PDF

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Publication number
CN115163564B
CN115163564B CN202210902671.1A CN202210902671A CN115163564B CN 115163564 B CN115163564 B CN 115163564B CN 202210902671 A CN202210902671 A CN 202210902671A CN 115163564 B CN115163564 B CN 115163564B
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China
Prior art keywords
line
blade tip
casing
rotor blade
highest point
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CN202210902671.1A
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CN115163564A (en
Inventor
沈锡钢
步明繁
冯冬民
翟贤超
王�华
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AECC Shenyang Engine Research Institute
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AECC Shenyang Engine Research Institute
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    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04D—NON-POSITIVE-DISPLACEMENT PUMPS
    • F04D29/00—Details, component parts, or accessories
    • F04D29/40—Casings; Connections of working fluid
    • F04D29/42—Casings; Connections of working fluid for radial or helico-centrifugal pumps
    • F04D29/44—Fluid-guiding means, e.g. diffusers
    • F04D29/441—Fluid-guiding means, e.g. diffusers especially adapted for elastic fluid pumps
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04D—NON-POSITIVE-DISPLACEMENT PUMPS
    • F04D29/00—Details, component parts, or accessories
    • F04D29/66—Combating cavitation, whirls, noise, vibration or the like; Balancing
    • F04D29/661—Combating cavitation, whirls, noise, vibration or the like; Balancing especially adapted for elastic fluid pumps
    • F04D29/666—Combating cavitation, whirls, noise, vibration or the like; Balancing especially adapted for elastic fluid pumps by means of rotor construction or layout, e.g. unequal distribution of blades or vanes
    • 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)
  • Structures Of Non-Positive Displacement Pumps (AREA)

Abstract

The application belongs to the field of design of aero-engines, and particularly relates to a casing slotting structure at an aero-engine blade tip, which comprises a casing, wherein the casing is sleeved on a rotor blade, an annular groove is formed in the position of the rotor blade on the inner wall surface of the casing, and the contour line formed on the axial section of the groove bottom of the annular groove is a groove bottom line a; the contour line of the notch of the annular groove formed on the axial section is a notch line b; an included angle is formed between the groove bottom line a and the groove opening line b, and the opening direction of the included angle is the reverse direction of the flow direction of the inner cartridge receiver of the cartridge receiver; the wear-resistant coating is coated in the annular groove, the thickness of the wear-resistant coating is equal to the minimum distance h between the groove bottom line a and the groove opening line b, airflow generated by the tip of the rotor and moving from the pressure surface side to the suction surface side is restrained, the flow loss generated by the tip of the rotor is reduced, and the efficiency and the surge margin of the gas compressor can be improved.

Description

Aeroengine blade tip department cartridge receiver fluting structure
Technical Field
The application belongs to the field of aero-engine design, in particular to a slotting structure of a casing at an aero-engine blade tip.
Background
For a high-pressure compressor of an aircraft engine, rotor and stator clearance control and rotor and stator collision and abrasion are always the key points of design. Improper clearance control between the rotor blade and the casing can cause collision and abrasion, and if the rotor blade directly collides with the casing, the blade can be damaged by curling, corner falling and the like. Depending on the material chosen, a titanium fire may also be initiated. And too large clearance reservation can cause aerodynamic loss at the blade tip and influence performance. Most of the flow paths of the casing of the compressor adopt a straight design, and the corresponding blade tip parts are coated with abradable coatings, so that a larger blade tip clearance is adopted to avoid the collision and abrasion of a rotor and a stator, and the performance loss is caused, as shown in fig. 4.
1. The clearance between the blades and the casing coating causes the rotor blade tips to generate airflow moving from the pressure surface side to the suction surface side, which causes flow loss at the blade tips, thereby reducing the efficiency and surge margin of the compressor.
2. Because rotor blade tip portion warp and the quick-witted casket warp discordance, reasonable apex clearance design degree of difficulty is big, easily leads to the engine during operation apex and quick-witted casket to bump and grind the damage.
Disclosure of Invention
In order to solve the problems, the application provides a slotting structure of a casing at the blade tip of an aeroengine; the method comprises the following steps:
the rotor blade is sleeved with the casing, an annular groove is formed in the position of the rotor blade on the inner wall surface of the casing, and the contour line formed on the axial section of the groove bottom of the annular groove is a groove bottom line a; the contour line of the notch of the annular groove formed on the axial section is a notch line b; an included angle is formed between the groove bottom line a and the groove opening line b, and the opening direction of the included angle is the reverse direction of the flow direction of the cartridge receiver in the cartridge receiver; the inside of the annular groove is coated with a wear-resistant coating, and the thickness of the wear-resistant coating is equal to the minimum distance h between the groove bottom line a and the groove opening line b.
Preferably, the size of the included angle is θ, and θ satisfies the following formula:
θ＝α-β；
the first straight line is formed by connecting the highest point of the front edge with the highest point of the rear edge when the rotor blade tip is in a static state, and the second straight line is formed by connecting the highest point of the front edge with the highest point of the rear edge when the rotor blade tip is in a working state;
beta is an included angle between a third straight line and a fourth straight line, the third straight line is formed by the highest point when the highest point of the rotor blade tip front edge corresponds to the highest point of the static state of the casing part corresponding to the rotor blade tip rear edge, and the fourth straight line is formed by the highest point of the rotor blade tip rear edge corresponding to the highest point of the working state of the casing part corresponding to the rotor blade tip rear edge.
Preferably, both ends of the annular groove have rounded corners.
Preferably, the radius R of the fillet satisfies: r =1+abs (Sin θ).
The application has the advantages that:
1. the annular groove with the inclined groove bottom inhibits the airflow generated by the rotor blade tip and moving from the pressure surface side to the suction surface side, reduces the flow loss generated by the blade tip, and can improve the efficiency and surge margin of the gas compressor.
2. The annular groove with the inclined groove bottom enables the coating to be worn more uniformly, and plays a role in protecting the apex angle weak area of the front edge of the blade tip.
Drawings
FIG. 1 is a schematic axial cross-sectional view of a slotted structure of a casing at a blade tip of an aircraft engine according to a preferred embodiment of the present application;
FIG. 2 is a schematic contour line of a rotor blade tip at rest and a rotor blade tip at work;
FIG. 3 is a schematic contour line diagram of a static state of a casing part corresponding to a rotor blade tip and a working state of the casing part corresponding to the rotor blade tip;
FIG. 4 is a schematic axial sectional view of a casing slotting structure at a blade tip of a conventional aircraft engine.
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 accompanying 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 some, but not all embodiments of the present application. The embodiments described below with reference to the accompanying 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 obtained by a person of ordinary skill in the art without any inventive work based on the embodiments in the present application are within the scope of protection of the present application. Embodiments of the present application will be described in detail below with reference to the drawings.
As shown in fig. 1, the present application provides a casing slotting structure at the blade tip of an aircraft engine; the method comprises the following steps:
the rotor blade is sleeved with the casing, an annular groove is formed in the position of the rotor blade on the inner wall surface of the casing, and the contour line formed on the axial section of the groove bottom of the annular groove is a groove bottom line a; the contour line of the notch of the annular groove formed on the axial section is a notch line b; an included angle is formed between the groove bottom line a and the groove opening line b, and the opening direction of the included angle is the reverse direction of the flow direction of the inner cartridge receiver of the cartridge receiver; a wear-resistant coating is coated in the annular groove, and the thickness of the wear-resistant coating is equal to the minimum distance h between the groove bottom line a and the groove opening line b; said another way is that: the slot line a is not parallel to the slot line b, the tail end of the slot line b is taken as an original point O, the rotation angle of the slot line a towards the outer side direction of the casing is also the included angle between the slot line a and the slot line b, and the minimum distance h between the slot line a and the slot line b is also the offset distance of the slot line a; compared with the traditional slotting structure of the casing at the blade tip of the aero-engine shown in fig. 4, the slotting structure of the traditional casing is used for accommodating the wear-resistant coating, and the scheme is not used for milling a slot along the casing, but is used for designing a blade tip chute.
Furthermore, the included angle is theta, the size of theta is determined by the strength calculation deformation value of the rotor blade tip and the casing in the working state, and theta satisfies the following formula:
θ＝α-β；
wherein alpha is an included angle between a first straight line and a second straight line, the first straight line is formed by connecting the highest point of the front edge with the highest point of the rear edge when the rotor blade tip is in a static state, and the second straight line is formed by connecting the highest point of the front edge with the highest point of the rear edge when the rotor blade tip is in a working state;
beta is the included angle between the third straight line and the fourth straight line, the third straight line is formed by the highest point when the highest point is connected with the highest point when the rotor blade tip front edge corresponds to the static state of the casing part, the fourth straight line is formed by the highest point when the highest point is connected with the highest point when the rotor blade tip rear edge corresponds to the working state of the casing part, and referring to fig. 2, wherein alpha is the contour line x of the rotor blade tip static state 1 x 2 Contour line x of working state of rotor blade tip 1 ’x 2 ' Angle of x 1 Is the highest point of the leading edge, x, when the rotor blade tip is in a static state 2 Is the highest point of the trailing edge when the rotor blade tip is in a static state, and the contour line x 1 x 2 Is a point x 1 And point x 2 The connecting line of (2); x is a radical of a fluorine atom 1 ' is the leading edge of the rotor blade tip in working stateHighest point, x 2 ' is the highest point of the trailing edge when the rotor blade tip is in working state, and the contour line x 1 ’x 2 ' is a point x 1 ' AND Point x 2 ' is connected; referring to FIG. 3, β is the contour line w of the rotor blade tip corresponding to the stationary position of the casing 1 w 2 Contour line w of working state of casing part corresponding to rotor blade tip 1 ’w 2 Angle of' w 1 Is the highest point, w, of the rotor blade tip leading edge in the static state corresponding to the casing part 2 Is the highest point of the rotor blade tip trailing edge corresponding to the casing part in a static state, and the contour line w 1 w 2 Is point w 1 And point w 2 The connecting line of (1); w is a 1 The highest point, w, of the rotor blade tip trailing edge in the operating state of the casing 2 The highest point of the rotor blade tip trailing edge corresponding to the casing part in the working state, the contour line w 1 ’w 2 ' is point w 1 ' AND point w 2 ' is connected;
the strength calculation and the previous trial run experience prove that the blade tip of the blade is not uniformly deformed, and the theta angle is properly compensated, so that the abrasion of the tip part of the front edge of the blade tip is reduced as much as possible by the coating during working.
Further, both ends of the annular groove have rounded corners.
Further, the radius R of the fillet satisfies: r =1+ abs (Sin theta), a better pneumatic profile can be formed, the air flow generated by the rotor blade tip and moving from the pressure surface side to the suction surface side is reduced, the flow loss generated by the blade tip is reduced, and the efficiency and the surge margin of the compressor can be improved.
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 (4)

1. The utility model provides an aeroengine apex department receiver fluting structure which characterized in that includes:
the rotor blade is sleeved with the casing, an annular groove is formed in the position of the rotor blade on the inner wall surface of the casing, and the contour line formed on the axial section of the groove bottom of the annular groove is a groove bottom line a; the contour line of the notch of the annular groove formed on the axial section is a notch line b; an included angle is formed between the groove bottom line a and the groove opening line b, and the opening direction of the included angle is the reverse direction of the flow direction of the inner cartridge receiver of the cartridge receiver; the inside of the annular groove is coated with a wear-resistant coating, and the thickness of the wear-resistant coating is equal to the minimum distance h between the groove bottom line a and the groove opening line b.
2. The aero-engine blade tip casing grooving structure of claim 1 wherein the included angle is θ, and θ satisfies the following equation:
θ＝α-β；
wherein alpha is an included angle between a first straight line and a second straight line, the first straight line is formed by connecting the highest point of the front edge with the highest point of the rear edge when the rotor blade tip is in a static state, and the second straight line is formed by connecting the highest point of the front edge with the highest point of the rear edge when the rotor blade tip is in a working state;
beta is an included angle between a third straight line and a fourth straight line, the third straight line is formed by the highest point when the highest point of the front edge of the rotor blade tip corresponds to the highest point of the static state of the casing part corresponding to the rear edge of the rotor blade tip, and the fourth straight line is formed by the highest point of the rear edge of the rotor blade tip corresponding to the highest point of the working state of the casing part corresponding to the rear edge of the rotor blade tip.
3. The slot structure of the casing at the blade tip of the aircraft engine as claimed in claim 2, wherein both ends of the annular groove have rounded corners.
4. The aircraft engine blade tip casing grooving structure according to claim 3, wherein the fillet radius R satisfies: r =1+abs (Sin θ).
CN202210902671.1A 2022-07-29 2022-07-29 Aeroengine blade tip position machine casket fluting structure Active CN115163564B (en)

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CN115163564B true CN115163564B (en) 2023-03-24

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Citations (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
RU32532U1 (en) * 2003-03-26 2003-09-20 Федеральное государственное унитарное предприятие "Центральный научно-исследовательский автомобильный и автомоторный институт" Turbocharger
CN101046162A (en) * 2006-03-30 2007-10-03 斯奈克玛 Device for fixing ring sectors on the casing of a jet engine
CN215860973U (en) * 2021-01-18 2022-02-18 中国航发商用航空发动机有限责任公司 Compressor rotor blade

Patent Citations (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
RU32532U1 (en) * 2003-03-26 2003-09-20 Федеральное государственное унитарное предприятие "Центральный научно-исследовательский автомобильный и автомоторный институт" Turbocharger
CN101046162A (en) * 2006-03-30 2007-10-03 斯奈克玛 Device for fixing ring sectors on the casing of a jet engine
CN215860973U (en) * 2021-01-18 2022-02-18 中国航发商用航空发动机有限责任公司 Compressor rotor blade

Non-Patent Citations (1)

* Cited by examiner, † Cited by third party
Title
大涵道比涡扇发动机风扇/增压级试验件结构设计及验证;张岩等;《航空发动机》;20200215;全文 *

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