Aero-engine intermediate pressure compressor tandem rotor blade assembly structure and method thereof

Abstract

The application belongs to the technical field of assembly design of middle pressure compressor tandem rotor blades in an aero-engine, and particularly relates to an assembly structure and a method of middle pressure compressor tandem rotor blades in an aero-engine, wherein the assembly structure of middle pressure compressor tandem rotor blades in an aero-engine comprises: a rotor disc, an annular clamping groove is formed on the outer edge rear side part; the rear side wall of the annular clamping groove has an opening; front-row rotor blades are formed on the outer edge front side part of the rotor disc in the circumferential direction; rear-row rotor blades have lower edge plates at the roots; each lower edge plate is sequentially clamped into the annular clamping groove through the opening, and the lower edge plates are tightly abutted in the circumferential direction, wherein the lower edge plate opposite to the opening is missing at the opening; and a plug is blocked at the opening.

Description

Technical Field

[0001] This application belongs to the technical field of tandem rotor blade assembly design for aero-engine compressors, specifically relating to an assembly structure and method for tandem rotor blades of aero-engine compressors. Background Technology

[0002] The tandem rotor technology for compressors in aero-engines involves dividing the rotor blades on the rotor disk into two rows with different blade profiles to increase the single-stage pressure ratio of the compressor, reduce the number of stages, and achieve a high thrust-to-weight ratio in aero-engines. Its configuration is as follows: Figure 1 As shown.

[0003] Currently, in aero-engine compressor tandem rotor blades, the front and rear rows of rotor blades are integrally formed on the rotor disk. In order to ensure high performance, the front and rear rows of rotor blades need to be designed with a small gap, which makes it difficult to guide the tool and the tool path is complex, resulting in low yield in machining and manufacturing.

[0004] This application is made in view of the aforementioned technical deficiencies.

[0005] It should be noted that the above background information is only used to assist in understanding the inventive concept and technical solution of this invention, and it does not necessarily belong to the prior art of this application. In the absence of clear evidence that the above information was disclosed on the filing date of this application, the above background information should not be used to evaluate the novelty and inventiveness of this application. Summary of the Invention

[0006] The purpose of this application is to provide an assembly structure and method for tandem rotor blades of a compressor in an aero-engine, so as to overcome or mitigate at least one of the known technical defects.

[0007] The technical solution of this application is:

[0008] One aspect provides an assembly structure for tandem rotor blades of a compressor in an aero-engine, comprising:

[0009] The rotor disk has an annular groove formed on the rear side of its outer edge; the rear side wall of the annular groove has a notch.

[0010] The front row of rotor blades is circumferentially formed on the front side of the outer edge of the rotor disk;

[0011] The rear rotor blades have a lower edge plate at the root; each lower edge plate is sequentially inserted into an annular groove through a notch and abuts against each other in the circumferential direction, wherein the lower edge plate directly opposite the notch is missing at the notch;

[0012] A blockage, used to seal off the opening.

[0013] According to at least one embodiment of this application, in the above-described compressor tandem rotor blade assembly structure of an aero-engine, the rear side of the outer edge of the rotor disk has an inward annular mounting edge.

[0014] The block has an inward connecting edge;

[0015] The inward connecting edge is fastened to the inward annular mounting edge by screws and their locking tabs.

[0016] On the other hand, a method for assembling tandem rotor blades of a compressor in an aero-engine is provided, including:

[0017] The lower edge plates at the root of the rear rotor blades are sequentially inserted through the notch into the annular groove formed on the rear side of the outer edge of the rotor disk, and pressed together in the circumferential direction.

[0018] The block is placed at the opening and abuts against the lower edge plate of the rear rotor blades on both sides to circumferentially stop the rear rotor blades. It also presses against the lower edge plate at the root of the rear rotor blades directly opposite the opening to radially position the rear rotor blades and axially position them.

[0019] This application has at least the following beneficial technical effects:

[0020] This invention provides an assembly structure and method for tandem rotor blades in an aero-engine compressor. The design features only the front row of rotor blades formed on the rotor disk, while the rear row of rotor blades are assembled by means of notches into annular grooves on the outer edge of the rotor disk. The notches are then sealed with plugs to stop and position the rear row of rotor blades. This design allows for a small gap between the front and rear rows of rotor blades, ensuring high performance and eliminating manufacturing difficulties. Attached Figure Description

[0021] Figure 1 This is a schematic diagram of the compressor tandem rotor blade structure in an aero-engine provided in the embodiments of this application;

[0022] Figure 2 This is a schematic diagram of the compressor tandem rotor blade assembly structure in an aero-engine provided in the embodiments of this application;

[0023] in:

[0024] 1-Rotor disc; 2-Front row rotor blades; 3-Rear row rotor blades; 4-Block.

[0025] To better illustrate this embodiment, some parts in the accompanying drawings may be omitted, enlarged, or reduced, and do not represent the actual product dimensions. Furthermore, the drawings are for illustrative purposes only and should not be construed as limiting this application. Detailed Implementation

[0026] To make the technical solution and advantages of this application clearer, the technical solution of this application will be described in a clearer and more complete manner below with reference to the accompanying drawings. It should be understood that the specific embodiments described herein are only some embodiments of this application, and are only used to explain this application, not to limit this application. It should be noted that, for ease of description, only the parts related to this application are shown in the accompanying drawings. Other related parts can be referred to the general design. In the absence of conflict, the embodiments and technical features in the embodiments of this application can be combined with each other to obtain new embodiments.

[0027] Furthermore, unless otherwise defined, the technical or scientific terms used in this application description shall have the ordinary meaning understood by one of ordinary skill in the art to which this application pertains. The terms "upper," "lower," "left," "right," "center," "vertical," "horizontal," "inner," and "outer," etc., used in this application description to indicate relative direction or positional relationship are used only to indicate relative orientation or positional relationship, and do not imply that the device or component must have a specific orientation, or be constructed and operated in a specific orientation. When the absolute position of the described object changes, its relative positional relationship may also change accordingly, and therefore should not be construed as a limitation on this application. The terms "first," "second," "third," and similar terms used in this application description are used only for descriptive purposes to distinguish different components, and should not be construed as indicating or implying relative importance. The terms "a," "one," or "the," etc., used in this application description should not be construed as an absolute limitation on quantity, but should be construed as indicating the existence of at least one. The terms "including," "comprising," etc., used in this application description mean that the element or object preceding the word covers the element or object listed after the word and its equivalents, without excluding other elements or objects.

[0028] Furthermore, it should be noted that, unless otherwise explicitly specified and limited, terms such as “installation,” “connection,” and “linkage” used in the description of this application should be interpreted broadly. For example, a connection can be a fixed connection, a detachable connection, or an integral connection; it can be a mechanical connection or an electrical connection; it can be a direct connection or an indirect connection through an intermediate medium; or it can be a connection within two components. Those skilled in the art can understand its specific meaning in this application according to the specific circumstances.

[0029] The following is in conjunction with the appendix Figures 1 to 2 This application will be described in further detail.

[0030] One aspect provides an assembly structure for tandem rotor blades of a compressor in an aero-engine, comprising:

[0031] The rotor disk 1 has an annular groove formed on the rear side of its outer edge; the rear side wall of the annular groove has a notch.

[0032] The front row of rotor blades 2 are formed circumferentially on the front side of the outer edge of the rotor disk 1;

[0033] The rear rotor blades 3 have a lower edge plate at the root; each lower edge plate is sequentially inserted into an annular groove through a notch and abuts against each other in the circumferential direction, wherein the lower edge plate directly opposite the notch is missing at the notch;

[0034] Block 4 is used to seal the gap.

[0035] Regarding the compressor tandem rotor blade assembly structure of the aero-engine disclosed in the above embodiments, those skilled in the art will understand that its design only has the front row of rotor blades 2 formed on the rotor disk 1, while the rear row of rotor blades 3 is assembled by means of notches and assembled into the annular grooves on the outer edge of the rotor disk 1. The notches are sealed with plugs 4 to stop and position the rear row of rotor blades 2. The front and rear rows of rotor blades can be designed to have a small gap, ensuring high performance indicators, and there are no problems with processing and manufacturing difficulties.

[0036] In some optional embodiments, in the above-described compressor tandem rotor blade assembly structure of the aero-engine, the rear part of the outer edge of the rotor disk 1 has an inward annular mounting edge.

[0037] Block 4 has an inward connecting edge;

[0038] The inward connecting edge is fastened to the inward annular mounting edge by screws and their locking tabs.

[0039] On the other hand, a method for assembling tandem rotor blades of a compressor in an aero-engine is provided, including:

[0040] The lower edge plates of the three roots of the rear rotor blades are sequentially inserted through the notch into the annular groove formed on the rear side of the outer edge of the rotor disk 1, and they are pressed together in the circumferential direction.

[0041] The block 4 is placed at the opening and abuts against the lower edge plate of the rear rotor blades 3 on both sides to stop the rear rotor blades 3 circumferentially, and is pressed against the lower edge plate of the root of the rear rotor blades 3 directly opposite the opening to radially position the rear rotor blades 3 and axially position the rear rotor blades 3.

[0042] The above-described method for assembling tandem rotor blades of aero-engine compressors is used to assemble the tandem rotor blade assembly structure of aero-engine compressors disclosed in the above-described embodiments. The description is relatively simple. For specific details, please refer to the relevant description of the tandem rotor blade assembly structure of aero-engine compressors. The technical effects can also be referred to the technical effects of the relevant parts of the tandem rotor blade assembly structure of aero-engine compressors, which will not be repeated here.

[0043] The various embodiments in the specification are described in a progressive manner, with each embodiment focusing on the differences from other embodiments. The same or similar parts between the various embodiments can be referred to each other.

[0044] The technical solution of this application has been described in conjunction with the preferred embodiments shown in the accompanying drawings. Those skilled in the art should understand that the protection scope of this application is obviously not limited to these specific embodiments. Without departing from the principles of this application, those skilled in the art can make equivalent changes or substitutions to the relevant technical features, and the technical solutions after these changes or substitutions will all fall within the protection scope of this application.

Claims

1. An assembly method of a high-pressure compressor tandem-rotor blade of an aircraft engine for assembling an assembly structure of a high-pressure compressor tandem-rotor blade of an aircraft engine, characterized by, The serial rotor blade assembly structure of the intermediate pressure compressor of the aero-engine comprises: a rotor disc (1) having an annular clamping groove formed at a rear side of an outer edge of the rotor disc (1), and a notch formed at a rear side wall of the annular clamping groove; a front row of rotor blades (2) formed at a front side of the outer edge of the rotor disc (1) in a circumferential direction; a rear row of rotor blades (3) each having a lower edge plate, and each of the lower edge plates being sequentially clamped into the annular clamping groove through the notch and abutting against each other in the circumferential direction, wherein the lower edge plate opposite to the notch is absent at the notch; a plug (4) sealing the notch; the rear side of the outer edge of the rotor disc (1) having an inward annular mounting edge; the plug (4) having an inward connecting edge; the inward connecting edge being fastened to the inward annular mounting edge by a screw and a locking plate of the screw; the serial rotor blade assembly method of the intermediate pressure compressor of the aero-engine comprising: clamping the lower edge plates of the root of the rear row of rotor blades (3) into the annular clamping groove formed at the rear side of the outer edge of the rotor disc (1) through the notch in sequence, and abutting against each other in the circumferential direction; sealing the plug (4) at the notch, abutting against the lower edge plates of the rear row of rotor blades (3) on both sides, circumferentially stopping the rear row of rotor blades (3), and pressing the lower edge plate of the root of the rear row of rotor blades (3) opposite to the notch, radially positioning the rear row of rotor blades (3), and axially positioning the rear row of rotor blades (3).

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