CN115030778A - Connection structure and connection method of turbine disc and blades of aircraft engine - Google Patents
Connection structure and connection method of turbine disc and blades of aircraft engine Download PDFInfo
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- CN115030778A CN115030778A CN202210754263.6A CN202210754263A CN115030778A CN 115030778 A CN115030778 A CN 115030778A CN 202210754263 A CN202210754263 A CN 202210754263A CN 115030778 A CN115030778 A CN 115030778A
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- 238000000034 method Methods 0.000 title claims abstract description 11
- 238000007789 sealing Methods 0.000 claims description 13
- 230000007704 transition Effects 0.000 description 3
- 230000000694 effects Effects 0.000 description 2
- 238000012986 modification Methods 0.000 description 2
- 230000004048 modification Effects 0.000 description 2
- 238000006467 substitution reaction Methods 0.000 description 2
- 230000003044 adaptive effect Effects 0.000 description 1
- 230000008602 contraction Effects 0.000 description 1
- 230000007547 defect Effects 0.000 description 1
- 230000005489 elastic deformation Effects 0.000 description 1
- 230000007717 exclusion Effects 0.000 description 1
- 238000009434 installation Methods 0.000 description 1
- 230000000750 progressive effect Effects 0.000 description 1
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Classifications
-
- 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
- F01D5/147—Construction, i.e. structural features, e.g. of weight-saving hollow blades
-
- 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
- F01D5/30—Fixing blades to rotors; Blade roots ; Blade spacers
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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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- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Architecture (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
The application belongs to the technical field of connection design of aero-engine turbine disks and blades, and particularly relates to a connection structure of an aero-engine turbine disk and a blade and a connection method thereof, wherein the connection structure comprises: the outer edge of the turbine disc is provided with a plurality of mortises distributed along the circumferential direction, and the side wall of one side of each mortise opening on the turbine disc is provided with an annular mounting groove; the side wall of the inner side of the annular mounting groove is provided with an annular flange; a plurality of turbine blades, the root having a tenon; each tenon is correspondingly clamped into one mortise; the side wall of one side of the baffle ring is provided with an annular mounting boss; the annular mounting boss extends into the annular mounting groove and is provided with an annular folded edge; and the mounting ring is clamped between the annular flange and the annular folded edge, so that each tenon is pressed in the corresponding mortise by the retaining ring.
Description
Technical Field
The application belongs to the technical field of connection design of aero-engine turbine disks and blades, and particularly relates to a connection structure and a connection method of an aero-engine turbine disk and blades.
Background
In the aero-engine, each turbine blade is connected to the outer edge of the turbine disc along the circumferential direction and rotates under the driving of the turbine disc, tenons and grooves are arranged between each turbine blade and the turbine disc in a matched mode, the radial and circumferential positioning is realized, and the axial positioning between each turbine blade and the turbine disc is realized mostly by the baffle ring connected to the turbine disc through bolts.
The present application has been made in view of the above-mentioned technical drawbacks.
It should be noted that the above background disclosure is only for the purpose of assisting understanding of the inventive concept and technical solutions of the present invention, and does not necessarily belong to the prior art of the present patent application, and the above background disclosure should not be used for evaluating the novelty and inventive step of the present application without explicit evidence to suggest that the above content is already disclosed at the filing date of the present application.
Disclosure of Invention
The object of the present application is to provide a connection structure of a turbine disk and a blade of an aircraft engine and a connection method thereof, so as to overcome or alleviate technical defects of at least one aspect of the known existence.
The technical scheme of the application is as follows:
one aspect provides a connection structure of an aircraft engine turbine disk and a blade, comprising:
the outer edge of the turbine disc is provided with a plurality of mortises distributed along the circumferential direction, and the side wall of one side of each mortise opening on the turbine disc is provided with an annular mounting groove; the side wall of the inner side of the annular mounting groove is provided with an annular flange;
a plurality of turbine blades, the root having a tenon; each tenon is correspondingly clamped into one mortise;
the side wall of one side of the baffle ring is provided with an annular mounting boss; the annular mounting boss extends into the annular mounting groove and is provided with an annular folded edge;
and the mounting ring is clamped between the annular flange and the annular folded edge, so that each tenon is pressed in the corresponding mortise by the retaining ring.
According to at least one embodiment of the application, in the connection structure of the aeroengine turbine disc and the blade, the cross section of the annular installation groove is arc-shaped.
According to at least one embodiment of the application, in the connection structure of the aeroengine turbine disc and the blade, the annular mounting boss abuts against the side wall outside the annular mounting groove.
According to at least one embodiment of the application, in the connection structure of the aeroengine turbine disk and the blade, the mounting ring is provided with a notch.
According to at least one embodiment of the present application, in the above-described connection structure of an aircraft engine turbine disk and a blade, 360 ° x < α <360 ° x +1 °;
wherein,
r2 is the radius of the inner side wall of the annular mounting groove;
r1 is the inner diameter of the mounting ring;
alpha is the angle of the notch on the mounting ring.
According to at least one embodiment of the application, in the connection structure of the aircraft engine turbine disc and the blade, the mounting ring is provided with an annular stop edge, and the annular stop edge abuts against the inner side of the annular flange.
According to at least one embodiment of the application, in the connection structure of the aeroengine turbine disc and the blades, a plurality of air introducing holes are formed in the side wall of the turbine disc, which faces away from the baffle ring;
an air guide channel is arranged between each mortise and the corresponding tenon; each air guide channel is correspondingly communicated with one air guide hole;
the side wall of the baffle ring is provided with an inner annular boss and an outer annular boss;
the inner annular bosses are tightly pressed on the tenons and the side wall of the turbine disc, and are communicated with the air guide channels in the space between the inner annular bosses and the annular mounting bosses;
the outer annular boss is pressed on the roots of the turbine blades and communicated with the space between the inner annular boss and the gaps between the roots of the turbine blades.
According to at least one embodiment of the application, in the connection structure of the aero-engine turbine disc and the blades, an inner annular seal groove is formed between the inner annular boss and the side wall of the turbine disc;
an outer annular sealing groove is formed between the outer annular boss and the side wall of the turbine disc;
the connection structure of aeroengine turbine dish and blade still includes:
the inner annular sealing ring is arranged in the inner annular sealing groove;
and the outer annular sealing ring is arranged in the outer annular sealing groove.
According to at least one embodiment of the application, in the connection structure of the aeroengine turbine disc and the blades, L4< L1;
L2+L5>L3；
wherein,
l1 is the height difference of the root and the tenon of each turbine blade in the radial direction thereof;
l2 is the height difference of the side wall of the turbine disk and the inner side of the annular flange in the axial direction;
l3 is the height difference between the side wall surface of the annular folding edge facing the baffle ring and the inner annular boss when the baffle ring is in a natural state;
l4 is the height difference between the inner annular boss and the outer annular boss under the natural state of the baffle ring;
l5 is the thickness of the mounting ring.
In another aspect, a method for connecting a turbine disk and a blade of an aircraft engine is provided, which includes:
mounting the tenon at the root of each turbine blade into a corresponding tenon groove at the outer edge of the turbine disc;
expanding the mounting ring to enable the inner diameter of the mounting ring to be larger than the outer diameter of the annular flange, and then plugging the mounting ring into the annular mounting groove in the side wall surface of the turbine disc;
contracting the mounting ring to enable the outer diameter of the mounting ring to be smaller than the inner diameter of the annular flanging, further extending the annular mounting boss on the side wall of the baffle ring into the annular mounting groove, and compressing to enable the annular flanging on the annular mounting boss to be located in the annular mounting groove relative to the mounting ring;
and loosening the mounting ring to enable the annular stop edge on the mounting ring to abut against the inner side of the annular folding edge, and loosening the annular mounting boss to enable the annular folding edge and the annular flange to clamp the main mounting ring.
Drawings
FIG. 1 is a schematic view of a connection structure of a turbine disk and a blade of an aircraft engine provided by an embodiment of the application;
FIG. 2 is a schematic illustration of a turbine disk provided by an embodiment of the present application;
FIG. 3 is a schematic view of a turbine blade provided by an embodiment of the present application;
FIG. 4 is a schematic view of a retainer ring provided in accordance with an embodiment of the present application;
FIG. 5 is a schematic cross-sectional view of a mounting ring provided by an embodiment of the present application;
wherein:
1-a turbine disk; 2-a turbine blade; 3-a baffle ring; 4, mounting a ring; 5-inner annular sealing ring; 6-outer annular sealing ring;
l1 is the height difference of the root and the tenon of each turbine blade in the radial direction thereof;
l2 is the height difference of the side wall of the turbine disk and the inner side of the annular rib in the axial direction;
l3 is the height difference between the side wall surface of the annular folding edge facing the baffle ring and the inner annular boss when the baffle ring is in a natural state;
l4 is the height difference between the inner annular boss and the outer annular boss under the natural state of the baffle ring;
l5 is the thickness of the mounting ring.
For a better explanation of the present embodiment, some parts of the drawings may be omitted, enlarged or reduced, and do not represent the size of an actual product, and furthermore, the drawings are for illustrative purposes only and should not be construed as limiting the present patent.
Detailed Description
In order to make the technical solutions and advantages of the present application clearer, the technical solutions of the present application will be further clearly and completely described in the following detailed description with reference to the accompanying drawings, and it should be understood that the specific embodiments described herein are only some of the embodiments of the present application, and are only used for explaining the present application, but not limiting the present application. It should be noted that, for convenience of description, only the parts related to the present application are shown in the drawings, other related parts may refer to general designs, and the embodiments and technical features in the embodiments in the present application may be combined with each other to obtain a new embodiment without conflict.
In addition, unless otherwise defined, technical or scientific terms used in the description of the present application shall have the ordinary meaning as understood by one of ordinary skill in the art to which the present application belongs. The terms "upper", "lower", "left", "right", "center", "vertical", "horizontal", "inner", "outer", and the like used in the description of the present application, which indicate orientations, are used only to indicate relative directions or positional relationships, and do not imply that the devices or elements must have a specific orientation, be constructed and operated in a specific orientation, and when the absolute position of the object to be described is changed, the relative positional relationships may be changed accordingly, and thus, should not be construed as limiting the present application. The use of "first," "second," "third," and the like in the description of the present application is for descriptive purposes only to distinguish between different components and is not to be construed as indicating or implying relative importance. The use of the terms "a," "an," or "the" and similar referents in the context of describing the application is not to be construed as an absolute limitation on the number, but rather as the presence of at least one. The word "comprising" or "comprises", and the like, when used in this description, is intended to specify the presence of stated elements or items, but not the exclusion of any other elements or items.
Further, it should be noted that, unless otherwise explicitly stated or limited, the terms "mounted," "connected," and the like as used in the description of the present application are to be construed broadly, e.g., the connection may be a fixed connection, a detachable connection, or an integral connection; can be mechanically or electrically connected; they may be directly connected or indirectly connected through an intermediate medium, or they may be connected through the inside of two elements, and those skilled in the art can understand their specific meaning in this application according to the specific situation.
The present application is described in further detail below with reference to fig. 1 to 5.
One aspect provides a connection structure of an aircraft engine turbine disk and a blade, comprising:
the outer edge of the turbine disc 1 is provided with a plurality of mortises distributed along the circumferential direction, and the side wall of one side of each mortise opening on the turbine disc is provided with an annular mounting groove; the side wall of the inner side of the annular mounting groove is provided with an annular flange;
a plurality of turbine blades 2, the roots of which have tenons; each tenon is correspondingly clamped into one mortise;
a baffle ring 3, one side wall of which is provided with an annular mounting boss; the annular mounting boss extends into the annular mounting groove and is provided with an annular folded edge;
and the mounting ring 4 is clamped between the annular flange and the annular folded edge, so that the tenon is pressed in the corresponding mortise by the retaining ring 3.
To the connection structure of the aircraft engine turbine disc and the blades disclosed in the above embodiment, a skilled person in the art can understand that the side wall of the inside of the annular mounting groove on the side wall of each mortise opening on the turbine disc 1 is provided with an annular flange, the annular mounting boss on the flange ring 3 extends into the annular mounting groove, the mounting ring 4 is clamped between the annular flange and the annular flange of the annular mounting boss, so as to fix the flange ring 3, and the tenon at the root of each turbine blade 2 is pressed in the corresponding mortise, thereby realizing circumferential positioning between each turbine blade 2 and the turbine disc 1, without arranging a plurality of bolt holes on the turbine disc 1, avoiding intermittent damage to the overall structure of the turbine disc, ensuring the overall rigidity of the turbine disc 1, ensuring the service life of the turbine disc 1, and facilitating disassembly and assembly.
In some optional embodiments, in the above-mentioned connection structure between the aircraft engine turbine disc and the blade, the cross section of the annular mounting groove is arc-shaped, so as to avoid generating a large stress concentration during the operation of the aircraft engine, and ensure the service life of the turbine disc 1, and correspondingly, the corresponding transition portion between the annular mounting boss and the annular flange may also be designed to be an adaptive arc shape.
In some alternative embodiments, in the above-mentioned connection structure of the aircraft engine turbine disk and the blade, the annular mounting boss abuts against the sidewall wall outside the annular mounting groove, so as to realize the centering and positioning of the baffle ring 3.
In some optional embodiments, in the above connection structure of a turbine disk and a blade of an aero-engine, the mounting ring 4 has a notch, so that the expansion and contraction of the mounting ring 4 facilitate the disassembly and assembly of the connection structure of the turbine disk and the blade of the aero-engine, which may be specifically referred to the connection method of the turbine disk and the blade of the aero-engine disclosed in the embodiments of the present application.
In some alternative embodiments, in the above connection structure of the aeroengine turbine disk and the blade, the angle is 360 ° × 1-R2/R1< α <360 ° × 1-R2/R1+1 °, wherein R2 is the radius of the inner side wall of the annular mounting groove; r1 is the inner diameter of the mounting ring 4; alpha is the angle of the notch on the mounting ring 4, and the notch on the mounting ring 4 can be ensured to have a smaller angle on the premise of meeting the use requirement.
In some optional embodiments, in the above connection structure of the aircraft engine turbine disc and the blade, the mounting ring 4 has an annular stop edge, and the annular stop edge abuts against the inner side of the annular flange, so as to realize centering and positioning of the mounting ring 4 and prevent the mounting ring 4 from falling between the annular flange and the annular flange.
In some alternative embodiments, in the above-mentioned connection structure of the aircraft engine turbine disk and the blade, the side wall of the turbine disk 1 facing away from the baffle ring 3 is provided with a plurality of air introducing holes;
an air guide channel is arranged between each mortise and the corresponding tenon; each air guide channel is correspondingly communicated with one air guide hole;
the side wall of the baffle ring 3 is provided with an inner annular boss and an outer annular boss;
the inner annular bosses are tightly pressed on the tenons and the side wall of the turbine disc 1 and are communicated with the air guide channels in the space between the inner annular bosses and the annular mounting bosses;
the outer annular boss is pressed against the roots of the turbine blades 2, and the space between the outer annular boss and the inner annular boss is communicated with the gap between the roots of the turbine blades 2.
For the connection structure of the aero-engine turbine disc and the blade disclosed in the above embodiment, it can be understood by those skilled in the art that when the size of the baffle ring 3 is too large, the strength is weakened, it is difficult to ensure reliable pressing of the tenon at the root of each turbine blade 2 in the corresponding mortise, the inner annular boss on the sidewall of the baffle ring 3 is designed to press on each tenon and the sidewall of the turbine disc 1, and the outer annular boss on the root of each turbine blade 2, the elastic deformation force of the baffle ring 3 can be utilized to ensure reliable pressing of the tenon at the root of each turbine blade 2 in the corresponding mortise through the corresponding size design, and the arc transition of the inner annular boss and the annular mounting boss and the arc transition of the inner annular boss and the outer annular boss can be designed to reduce the local stress of the baffle ring 3, the service life of the baffle ring 3 is ensured.
For the connection structure of the aircraft engine turbine disc and the blades disclosed in the above embodiment, those skilled in the art can also understand that gas can flow into the space between the annular boss on the inner side of the baffle ring 3 and the annular mounting boss through the bleed holes on the side wall of the turbine disc 1, and that gas flow can flow the gap between the roots of the turbine blades 2 into the space between the annular boss on the outer side of the baffle ring 3 and the annular boss on the inner side of the baffle ring 3, so as to cool the turbine disc 1 and the baffle ring 3 thereof, and prolong the service life of the turbine disc 1 and the baffle ring 3 in a high-temperature environment.
In some optional embodiments, in the above connection structure between the aircraft engine turbine disc and the blade, an inner annular sealing groove is formed between the inner annular boss and the sidewall of the turbine disc 1;
an outer annular sealing groove is formed between the outer annular boss and the side wall of the turbine disc 1;
the connection structure of aeroengine turbine dish and blade still includes:
an inner annular seal ring 5 arranged in the inner annular seal groove;
and an outer annular seal ring 6 arranged in the outer annular seal groove.
In some alternative embodiments, in the above-described connection structure of the aero-engine turbine disk and the blade, L4< L1; l2+ L5> L3; wherein L1 is a height difference between the root and the tenon of each turbine blade 2 in the radial direction thereof; l2 is the height difference of the side wall of the turbine disc 1 and the inner side of the annular flange in the axial direction; l3 is the height difference between the annular flange facing the side wall surface of the baffle ring 3 and the inner annular boss when the baffle ring 3 is in a natural state; l4 is the height difference between the inner annular boss and the outer annular boss of the retainer ring 3 in the natural state; l5 is the thickness of collar 4 to guarantee that the assembly is accomplished the back, can reliably compress tightly the tongue and groove that the outer fringe of turbine disc 1 corresponds with the tenon of each turbine blade 2 root, specific numerical value can be confirmed by relevant technical staff when using by finite element analysis, guarantees that when aeroengine during operation, baffle 3 still can be with the reliable tongue and groove that compresses tightly in turbine disc 1 outer fringe corresponds of tenon of each turbine blade 2 root.
In another aspect, a method for connecting a turbine disk and a blade of an aircraft engine is provided, which includes:
installing tenons at the roots of the turbine blades 2 into mortises corresponding to the outer edge of the turbine disc 1;
expanding the mounting ring 4 to enable the inner diameter of the mounting ring 4 to be larger than the outer diameter of the annular flange, and further plugging the mounting ring 4 into the annular mounting groove on the side wall surface of the turbine disc 1;
contracting the mounting ring 4 to enable the outer diameter of the mounting ring 4 to be smaller than the inner diameter of the annular flanging, further extending the annular mounting boss on the side wall of the baffle ring 3 into the annular mounting groove, and compressing to enable the annular flanging on the annular mounting boss to be located in the annular mounting groove relative to the mounting ring 4;
and loosening the mounting ring 4 to enable the annular stop edge on the mounting ring 4 to abut against the inner side of the annular flanging, and loosening the annular mounting boss to enable the annular flanging and the annular flange to clamp the main mounting ring 4.
In order to facilitate the connection of the connection structure of the aero-engine turbine disc and the blade disclosed in the above embodiment, a person skilled in the art may design or select a corresponding connection tool with reference to the above connection method of the aero-engine turbine disc and the blade disclosed in the above embodiment.
For the connection method of the aero-engine turbine disk and the blade disclosed in the above embodiment, the connection structure of the aero-engine turbine disk and the blade is implemented, the description is simple, specific relevant parts can be referred to the relevant description of the connection structure part of the aero-engine turbine disk and the blade, and the technical effects can also be referred to the technical effects of the relevant part of the connection structure of the aero-engine turbine disk and the blade, which are not described again here.
The embodiments are described in a progressive mode in the specification, the emphasis of each embodiment is on the difference from the other embodiments, and the same and similar parts among the embodiments can be referred to each other.
Having thus described the present application in connection with the preferred embodiments illustrated in the accompanying drawings, it will be understood by those skilled in the art that the scope of the present application is not limited to those specific embodiments, and that equivalent modifications or substitutions of related technical features may be made by those skilled in the art without departing from the principle of the present application, and those modifications or substitutions will fall within the scope of the present application.
Claims (10)
1. A connection structure of an aeroengine turbine disk and blades, characterized by comprising:
the outer edge of the turbine disc (1) is provided with a plurality of mortises distributed along the circumferential direction, and the side wall of one side of each mortise opening is provided with an annular mounting groove; the side wall of the inner side of the annular mounting groove is provided with an annular flange;
a plurality of turbine blades (2), the root having a tenon; each tenon is correspondingly clamped into one mortise;
the side wall of one side of the baffle ring (3) is provided with an annular mounting boss; the annular mounting boss extends into the annular mounting groove and is provided with an annular folded edge;
and the mounting ring (4) is clamped between the annular flange and the annular folded edge, so that the tenon is pressed in the corresponding mortise by the retaining ring (3).
2. The aircraft engine turbine disk to blade connection structure according to claim 1,
the cross section of the annular mounting groove is arc-shaped.
3. The aircraft engine turbine disk to blade connection structure according to claim 1,
the annular mounting boss is abutted against the side wall on the outer side of the annular mounting groove.
4. The aircraft engine turbine disk to blade connection structure according to claim 1,
the mounting ring (4) is provided with a notch.
5. The aircraft engine turbine disk to blade connection structure according to claim 4,
360°×(1-R2/R1)<α<360°×(1-R2/R1)+1°；
wherein,
r2 is the radius of the inner side wall of the annular mounting groove;
r1 is the inner diameter of the mounting ring (4);
alpha is the angle of the notch on the mounting ring (4).
6. The aircraft engine turbine disk to blade connection structure according to claim 1,
the mounting ring (4) is provided with an annular stop edge which abuts against the inner side of the annular folded edge.
7. The aircraft engine turbine disk to blade connection structure according to claim 1,
the side wall of the turbine disc (1) back to one side of the baffle ring (3) is provided with a plurality of air guide holes;
an air guide channel is arranged between each mortise and the corresponding tenon; each air-entraining channel is correspondingly communicated with one air-entraining hole;
the side wall of the baffle ring (3) is provided with an inner annular boss and an outer annular boss;
the inner annular bosses are tightly pressed on the tenons and the side wall of the turbine disc (1), and are communicated with the air guide channels in the space between the inner annular bosses and the annular mounting bosses;
the outer annular boss is tightly pressed on the roots of the turbine blades (2), and the outer annular boss is communicated with the space between the inner annular bosses to form a gap between the roots of the turbine blades (2).
8. The aircraft engine turbine disk to blade connection structure according to claim 7,
an inner annular sealing groove is formed between the inner annular boss and the side wall of the turbine disc (1);
an outer annular sealing groove is formed between the outer annular boss and the side wall of the turbine disc (1);
the connection structure of aeroengine turbine dish and blade still includes:
an inner annular seal ring (5) disposed in the inner annular seal groove;
and the outer annular sealing ring (6) is arranged in the outer annular sealing groove.
9. The aircraft engine turbine disk to blade connection structure according to claim 7,
L4<L1；
L2+L5>L3；
wherein,
l1 is the height difference of the root and the tenon of each turbine blade (2) in the radial direction thereof;
l2 is the height difference between the side wall of the turbine disc (1) and the inner side of the annular rib in the axial direction;
l3 is the height difference between the side wall surface of the annular folding edge facing the baffle ring (3) and the inner annular boss when the baffle ring (3) is in a natural state;
l4 is the height difference between the inner annular boss and the outer annular boss when the baffle ring (3) is in a natural state;
l5 is the thickness of the mounting ring (4).
10. A method for connecting a turbine disk and a blade of an aircraft engine, comprising:
installing tenons at the roots of the turbine blades (2) into mortises corresponding to the outer edge of the turbine disc (1);
the mounting ring (4) is expanded, so that the inner diameter of the mounting ring (4) is larger than the outer diameter of the annular flange, and the mounting ring (4) is plugged into the annular mounting groove on the side wall surface of the turbine disc (1);
contracting the mounting ring (4) to enable the outer diameter of the mounting ring (4) to be smaller than the inner diameter of the annular flanging, further extending the annular mounting boss on the side wall of the baffle ring (3) into the annular mounting groove, and pressing to enable the annular flanging on the annular mounting boss to be located in the annular mounting groove relative to the mounting ring (4);
and loosening the mounting ring (4) to enable the annular stop edge on the mounting ring (4) to abut against the inner side of the annular folded edge, and loosening the annular mounting boss to enable the annular folded edge and the annular flange to clamp the main mounting ring (4).
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202210754263.6A CN115030778A (en) | 2022-06-28 | 2022-06-28 | Connection structure and connection method of turbine disc and blades of aircraft engine |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202210754263.6A CN115030778A (en) | 2022-06-28 | 2022-06-28 | Connection structure and connection method of turbine disc and blades of aircraft engine |
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| Publication Number | Publication Date |
|---|---|
| CN115030778A true CN115030778A (en) | 2022-09-09 |
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| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN202210754263.6A Pending CN115030778A (en) | 2022-06-28 | 2022-06-28 | Connection structure and connection method of turbine disc and blades of aircraft engine |
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Cited By (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN115791142A (en) * | 2023-02-09 | 2023-03-14 | 中国航发四川燃气涡轮研究院 | Axial limiting blade structure and configuration method |
| CN116104586A (en) * | 2023-04-11 | 2023-05-12 | 中国航发沈阳发动机研究所 | Locking and fixing structure of turbine rotor blade and turbine disk |
-
2022
- 2022-06-28 CN CN202210754263.6A patent/CN115030778A/en active Pending
Cited By (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN115791142A (en) * | 2023-02-09 | 2023-03-14 | 中国航发四川燃气涡轮研究院 | Axial limiting blade structure and configuration method |
| CN116104586A (en) * | 2023-04-11 | 2023-05-12 | 中国航发沈阳发动机研究所 | Locking and fixing structure of turbine rotor blade and turbine disk |
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