CN217380657U

CN217380657U - Impeller machine and aircraft engine

Info

Publication number: CN217380657U
Application number: CN202221215019.4U
Authority: CN
Inventors: 郭晓杰; 吴丽军; 王科
Original assignee: AECC Commercial Aircraft Engine Co Ltd
Current assignee: AECC Commercial Aircraft Engine Co Ltd
Priority date: 2022-05-20
Filing date: 2022-05-20
Publication date: 2022-09-06
Anticipated expiration: 2032-05-20

Abstract

The present disclosure provides a turbomachine and an aircraft engine. The impeller machine comprises at least one blade group, the blade group comprises a plurality of blades distributed along the circumferential direction of the impeller machine, each blade of the blade group comprises a blade body and a flange plate, and a mounting gap extending along the span direction of the blade body in a zigzag mode is formed between the flange plates of two adjacent blades. The aircraft engine comprises the impeller machine. The impeller machinery and the aircraft engine provided by the disclosure can reduce gas or cooling medium leakage.

Description

Impeller machine and aircraft engine

Technical Field

The disclosure relates to the technical field of aircraft engines, in particular to an impeller machine and an aircraft engine comprising the same.

Background

The gas turbine engine consists of a gas compressor, a combustion chamber and a turbine, wherein the gas compressor compresses gas, the combustion chamber provides energy through combustion, and the turbine expands to do work to drive the gas compressor. The turbine has several stages, each stage of turbine blade consists of stator blade fixed onto the turbine casing and rotor blade fixed onto the turbine disc. The blades are positioned in the gas environment of a main runner of an engine, the high temperature resistance is strong, components except the turbine blades, particularly a turbine disc, are limited in high temperature resistance, cooling media are required to be adopted for cooling, and a rim sealing structure is designed at the rim of the root of the turbine blades to prevent gas from leaking into the turbine to cause over-temperature of the components such as the turbine disc and the like or prevent the cooling media from leaking into the main runner.

SUMMERY OF THE UTILITY MODEL

It is an object of the present disclosure to provide a rim seal structure, a turbine and an aircraft engine to reduce gas or cooling medium leakage.

A first aspect of the present disclosure provides an impeller machine, including at least one blade group, where the blade group includes a plurality of blades distributed along a circumferential direction of the impeller machine, and each blade of the blade group includes a blade body and a flange, where a mounting gap extending in a zigzag manner in a span direction of the blade body is provided between the flanges of two adjacent blades.

According to some embodiments of the present disclosure, the flange includes at least one convex portion protruding in an axial direction of the turbomachine, and the mounting gap includes a seal section located between adjacent convex portions of two adjacent blades, the seal section extending zigzag in a span-wise direction of the blade body.

In accordance with some embodiments of the present disclosure,

the at least one blade group comprises a guide blade group, each blade of the guide blade group is a guide blade, the blade body is a guide blade body of the guide blade, the flange plate is a guide blade flange plate of the guide blade, the mounting gap comprises a first mounting gap, and the first mounting gap is formed between two adjacent guide blade flange plates and extends along the span direction of the guide blade body in a zigzag manner; and/or

The at least one blade group comprises a moving blade group, each blade of the moving blade group is a moving blade, the blade body is a moving blade body of the moving blade, the flange plate is a moving blade flange plate of the moving blade, the mounting gap comprises a second mounting gap, and the second mounting gap is formed between two adjacent moving blade flange plates and extends along the span direction of the moving blade body in a zigzag manner.

In accordance with some embodiments of the present disclosure,

the guide vane flange plate and the at least one boss of one of the moving vane flange plates comprise a first boss and a second boss which are arranged at intervals along the radial direction of the impeller machinery, the guide vane flange plate and the at least one boss of the other of the moving vane flange plates comprise a third boss which is arranged between the first boss and the second boss along the radial direction of the impeller machinery, and a sealing cavity is formed between the first boss, the second boss and the third boss.

In accordance with some embodiments of the present disclosure,

the at least one boss of one of the bucket platform and the bucket platform includes first and second bosses disposed at intervals in a radial direction of the turbomachine, the at least one boss of the other of the bucket platform and the bucket platform includes a third boss located between the first and second bosses in the radial direction of the turbomachine, wherein,

one of the first mounting gap and the second mounting gap comprises a first sealing section between two adjacent first bosses and a second sealing section between two adjacent second bosses; and/or

The other of the first mounting gap and the second mounting gap includes a third seal section between two adjacent third bosses.

According to some embodiments of the disclosure, a cross-section of at least one of the packing segments is stepped in shape.

According to some embodiments of the disclosure, a cross-section of at least one of the sealing sections is S-shaped.

A second aspect of the present disclosure provides an aircraft engine comprising the impeller machine of the first aspect of the present disclosure.

In the impeller machinery that this disclosure provided, have between the listrium of two adjacent blades along the span of blade body to the installation clearance of zigzag extension, can increase gaseous flow resistance through this installation clearance, consequently do benefit to the leakage that reduces cooling gas and gas to reduce the quantity of cooling gas, reduce aeroengine's the local overtemperature risk of spare part, promote aeroengine's efficiency. The aero-engine provided by the present disclosure has the corresponding advantages of the aforementioned impeller machine.

Other features of the present disclosure and advantages thereof will become apparent from the following detailed description of exemplary embodiments thereof, which proceeds with reference to the accompanying drawings.

Drawings

The accompanying drawings, which are included to provide a further understanding of the disclosure and are incorporated in and constitute a part of this application, illustrate embodiment(s) of the disclosure and together with the description serve to explain the disclosure and not to limit the disclosure. In the drawings:

fig. 1 shows a rim sealing structure of an aircraft engine in the related art.

FIG. 2 illustrates a gas leakage between adjacent vanes of the aircraft engine shown in FIG. 1.

FIG. 3 illustrates a gas leak between adjacent buckets of the aircraft engine shown in FIG. 1.

Fig. 4 shows a mounting gap between adjacent vanes or adjacent buckets of the aircraft engine shown in fig. 1.

Fig. 5 illustrates a first mounting clearance of the turbomachinery of some embodiments of the present disclosure.

Fig. 6 illustrates a second mounting clearance of the turbomachinery of some embodiments of the present disclosure.

FIG. 7 illustrates a cross-section of a first mounting gap or a second mounting gap in some embodiments of the present disclosure.

FIG. 8 illustrates a cross-section of a first mounting gap or a second mounting gap in further embodiments of the present disclosure.

In fig. 1 to 4, each reference numeral represents:

F. a primary air flow; f1, cooling air; f2, fuel gas; FA. An air flow leaking from the first mounting gap; FB. An air flow leaking from the second mounting gap; c', a sealing structure; c1', cooling air supply cavity;

10', a guide vane body; 11', guide vane edge plates; 12', a first sealing sheet;

20', a bucket body; 21', a moving blade flange plate; 22' and a second sealing sheet;

g1', a first mounting gap; g2', a second mounting gap;

in fig. 5 to 8, each reference numeral represents:

10. a guide vane body; 11. a vane flange; 11A, a first boss; 11B, a second boss;

20. a movable blade body; 21. a moving blade flange plate; 21A, a third boss;

g1, first mounting gap; g2, second mounting gap; g11, a first sealing section; g12, a second sealing section; g21, and a third sealing section.

Detailed Description

The technical solutions in the embodiments of the present disclosure will be clearly and completely described below with reference to the drawings in the embodiments of the present disclosure, and it is obvious that the described embodiments are only a part of the embodiments of the present disclosure, and not all of the embodiments. The following description of at least one exemplary embodiment is merely illustrative in nature and is in no way intended to limit the disclosure, its application, or uses. All other embodiments, which can be derived by a person skilled in the art from the embodiments disclosed herein without making any creative effort, shall fall within the protection scope of the present disclosure.

The relative arrangement of the components and steps, the numerical expressions, and numerical values set forth in these embodiments do not limit the scope of the present disclosure unless specifically stated otherwise. Meanwhile, it should be understood that the sizes of the respective portions shown in the drawings are not drawn in an actual proportional relationship for the convenience of description. Techniques, methods, and apparatus known to those of ordinary skill in the relevant art may not be discussed in detail, but are intended to be part of the specification where appropriate. In all examples shown and discussed herein, any particular value should be construed as merely illustrative, and not limiting. Thus, other examples of the exemplary embodiments may have different values. It should be noted that: like reference numbers and letters refer to like items in the following figures, and thus, once an item is defined in one figure, further discussion thereof is not required in subsequent figures.

In the description of the present disclosure, it should be understood that the terms "first", "second", etc. are used to define the components, and are used only for convenience of distinguishing the corresponding components, and if not otherwise stated, the terms have no special meaning, and thus, should not be construed as limiting the scope of the present disclosure.

In the description of the present disclosure, it is to be understood that the orientation or positional relationship indicated by the directional terms such as "front, rear, upper, lower, left, right", "lateral, vertical, horizontal" and "top, bottom", etc., are generally based on the orientation or positional relationship shown in the drawings, and are presented only for the convenience of describing and simplifying the disclosure, and in the absence of a contrary indication, these directional terms are not intended to indicate and imply that the device or element being referred to must have a particular orientation or be constructed and operated in a particular orientation, and therefore, should not be taken as limiting the scope of the disclosure; the terms "inner and outer" refer to the inner and outer relative to the profile of the respective component itself.

Fig. 1 shows a rim sealing structure of an aircraft engine in the related art, which is disposed between a guide vane and a movable vane, wherein the guide vane includes a guide vane body 10 ' and a guide vane flange plate 11 ', the movable vane includes a movable vane body 20 ' and a movable vane flange plate 21 ', and a zigzag extending sealing structure C ' is formed between the guide vane flange plate 11 ' and the movable vane flange plate 12 '.

Generally, the cooling air F1 from the cooling air supply chamber C1 'enters the main air flow F through the sealing structure C' as a cooling medium, and part of the fuel gas F2 in the high-pressure region of the main air flow F also enters the sealing structure C ', and the cooling air F1 and the fuel gas F2 are mixed in the sealing structure C' and then discharged into the main air flow F. The above-mentioned rim sealing structure is used for preventing the high-temperature fuel gas F2 in the main air flow F from entering the interior of the aircraft engine.

However, in carrying out the present disclosure, the inventors found that there is a certain mounting gap between adjacent two vanes and between adjacent two buckets. As shown in fig. 2, there is a first mounting gap G1 'between the vane platform plates 11' of two adjacent vanes; as shown in fig. 3, a second mounting gap G2 'exists between the bucket platform plates 12' of two adjacent buckets. At the positions of these mounting gaps, depending on the pressure of the cooling air F1 and the main air flow F, there may be air flows FA leaking from the first mounting gap G1 'through the first mounting gap G1' and air flows FB leaking from the second mounting gap G2 'through the second mounting gap G2'. If the pressure of the main air flow F is higher than the pressure of the cooling air F1, the combustion gases F2 will enter the cooling air supply chamber C1 ' through the first mounting gap G1 ' or the second mounting gap G2 ', causing a local overtemperature of the internal structure of the aircraft engine. If the pressure of the main air flow F is lower than the pressure of the cooling air F1, the cooling air F1 enters the main air flow F from the seal structure C ' through the first mounting gap G1 ' or the second mounting gap G2 ', so that the amount of use of the cooling air F1 increases.

As can be seen in fig. 4, either the first mounting gap G1 'or the second mounting gap G2' is straight-through. In such a straight-through installation gap, the gas flow resistance is low, which is not favorable for preventing the gas F2 from entering the sealing structure C' or reducing the usage amount of the cooling gas F1.

As shown in FIG. 1, for locations intermediate the vane and bucket platform 11 ', 12', gas leakage between the structural gaps may be reduced by installing first and second seal flaps 12 ', 22'. However, for the positions of the two ends of the guide vane flange plate 11 'and the movable vane flange plate 12', namely the positions for forming the rim sealing structure, the installation gap is still in a straight-through type, and the gas leakage is difficult to reduce by the mode of installing the sealing sheets.

Considering that the pressure of the main air flow F is not uniformly distributed along the circumferential direction at the outlet position of the rim sealing structure, the high-pressure area has periodic distribution, the cooling air F1 and the fuel gas F2 are still easy to leak through the straight-through installation gap of the rim sealing structure, and the leakage causes the increase of the cooling air F1, increases the local over-temperature risk of the parts of the aircraft engine, and has adverse effect on the efficiency of the aircraft engine.

To ameliorate the above problem, embodiments of the present disclosure provide a turbomachine and an aircraft engine.

As shown in fig. 5 to 8, an embodiment of the present disclosure provides an impeller machine including at least one blade group, where the blade group includes a plurality of blades distributed along a circumferential direction of the impeller machine, each blade of the blade group includes a blade body and a rim plate, and a mounting gap extending zigzag along a span direction of the blade body is provided between the rim plates of two adjacent blades.

The blades of the blade group may be a guide blade group composed of a plurality of guide blades, or a movable blade group composed of a plurality of movable blades. The shape and location of the mounting gap that meanders along the span of the vane body may be determined according to the requirement of the turbomachinery to reduce gas leakage. The mounting gap extending in a zigzag manner along the span direction of the blade body may extend in a zigzag manner or in a curved manner. The gap formed by the flanges, which may be two adjacent blades, may be partially the above-mentioned zigzag-extending mounting gap, for example, the zigzag-extending mounting gap may be located at a position where the first sealing sheet 12 'is not provided at both ends of the guide flange 11' in fig. 1 and a position where the second sealing sheet 22 'is not provided at both ends of the movable flange 12' in fig. 1; the gaps formed by the flanges of two adjacent blades may all be the mounting gaps extending in a zigzag manner.

Among the impeller machinery that this disclosed embodiment provided, have between the listrium of two adjacent blades along the span of blade body to the installation clearance of tortuous extension, can increase gaseous flow resistance through this installation clearance, consequently do benefit to and reduce the leakage of gas and gas for the cooling to reduce the quantity of gas for the cooling, reduce aeroengine's the local super temperature risk of spare part, promote aeroengine's efficiency.

In some embodiments, the platform includes at least one boss projecting in an axial direction of the turbomachine, and the mounting gap includes a seal section between adjacent bosses of two adjacent blades, the seal section extending zigzag in a span-wise direction of the blade body.

In this embodiment, the section of obturating extends and is located between the adjacent bellying of two adjacent blades to the zigzag along the exhibition of blade body, can form the structure of obturating in such position that is difficult to install the piece of obturating between the bellying, does benefit to and increases gaseous flow resistance in the position that is difficult to install the piece of obturating, reduces the leakage of gas and gas for the cooling, promotes the effect of obturating.

In some embodiments, as shown in fig. 5, at least one of the blade groups includes a guide blade group, each blade of the guide blade group is a guide blade, the blade body is a guide blade body 10 of the guide blade, the edge plate is a guide blade edge plate 11 of the guide blade, the mounting gap includes a first mounting gap G1, and the first mounting gap G1 is formed between two adjacent guide blade edge plates 11 and extends zigzag along the span direction of the guide blade body 10.

In some embodiments, as shown in fig. 6, at least one blade group includes a blade group, each blade of the blade group is a blade, the blade body is a blade body 20 of the blade, the platform is a blade platform 21 of the blade, the mounting gap includes a second mounting gap G2, and the second mounting gap G2 is formed between two adjacent blade platform 21 and meanders along a span direction of the blade body 20.

In some embodiments, the at least one blade group may include a guide blade group and a moving blade group arranged side by side in an axial direction of the impeller machine. For example, the guide blade group may be disposed on the upstream side in the airflow direction of the turbo machine, and the moving blade group may be disposed on the downstream side in the airflow direction of the turbo machine; the rotor blade group may be provided on the upstream side in the airflow direction of the turbo machine, and the guide blade group may be provided on the downstream side in the airflow direction of the turbo machine.

In some embodiments, the at least one boss of one of the vane and

bucket platform

11, 21 comprises first and

second bosses

11A, 11B spaced radially of the turbomachine, the at least one boss of the other of the vane and

bucket platform

11, 21 comprises a third boss 21A radially of the turbomachine between the first and

second bosses

11A, 11B, 21A forming a seal cavity therebetween.

For example, in the embodiment shown in fig. 5 and 6, the at least one boss of the vane platform 11 includes a first boss 11A and a second boss 11B, the at least one boss of the bucket platform 21 includes a third boss 21A, and a seal cavity similar to the seal structure C' in fig. 1 may be formed among the first boss 11A, the second boss 11B, and the third boss 21A.

For another example, in some embodiments not shown, the at least one boss of the bucket platform 21 includes a first boss and a second boss, the at least one boss of the bucket platform 11 includes a third boss, and a seal cavity similar to the seal structure C' of fig. 1 may be formed between the first boss, the second boss, and the third boss.

In this embodiment, not only form between the adjacent stator vane flange plate 11 along the span of stator vane body 10 to first installation gap G1 of zigzag extension, form between the adjacent movable vane flange plate 21 along the span of movable vane body 20 to second installation gap G2 of zigzag extension, stator vane flange plate 11 and movable vane flange plate 21 form the seal chamber of zigzag extension moreover, turbomachinery radially and axially all can form the structure of obturating, do benefit to the flow resistance of increase gas on different directions, reduce the leakage of cooling gas and gas, promote the effect of obturating.

In some embodiments, as shown in FIGS. 5 and 6, the at least one boss of one of the vane and

bucket platform

11, 21 comprises first and

second bosses

bucket platform

second bosses

11A, 11B, wherein one of the first mounting gap G1 and the second mounting gap G2 includes a first sealing section G11 between adjacent first bosses 11A and a second sealing section G12 between adjacent second bosses 11B, and/or the other of the first mounting gap G1 and the second mounting gap G2 includes a third obturating section G21 located between two adjacent third bosses 21A.

For example, in the embodiment shown in fig. 5 and 6, the at least one boss of the vane platform 11 includes a first boss 11A and a second boss 11B, the at least one boss of the bucket platform 21 includes a third boss 21A, the first mounting gap G1 includes a first seal segment G11 between two adjacent first bosses 11A and a second seal segment G12 between two adjacent second bosses 11B, and the second mounting gap G2 includes a third seal segment G21 between two adjacent third bosses 21A.

For another example, in some embodiments not shown, the at least one boss of the bucket platform 21 includes a first boss and a second boss, the at least one boss of the bucket platform 11 includes a third boss, the second mounting gap G2 includes a first seal segment between two adjacent first bosses and a second seal segment between two adjacent second bosses, and the first mounting gap G1 includes a third seal segment between two adjacent third bosses.

In this embodiment, between two adjacent first bellying 11A, between two adjacent second bellying 11B and between two adjacent third bellying 21A all form the section of obturating of zigzag extension, do benefit to the flow resistance of increase gas in a plurality of positions, promote to reduce the leakage of gas and gas for the cooling, promote the effect of obturating.

In some embodiments, the cross-section of at least one of the packing sections is stepped in shape. For example, as shown in fig. 7, the cross-sections of the first seal segment G11, the second seal segment G12, and the third seal segment G21 may each have a stepped shape.

In some embodiments, the cross-section of at least one of the sealing sections is S-shaped. For example, as shown in fig. 8, the cross-sections of the first sealing section G11, the second sealing section G12, and the third sealing section G21 may each be S-shaped.

Some embodiments of the present disclosure also provide an aircraft engine comprising the aforementioned impeller machine. The aircraft engine has the advantages of the impeller machine.

Finally, it should be noted that: the above examples are intended only to illustrate the technical solutions of the present disclosure and not to limit them; although the present disclosure has been described in detail with reference to preferred embodiments, those of ordinary skill in the art will understand that: modifications to the embodiments of the disclosure or equivalent replacements of parts of the technical features may be made, which are all covered by the technical solution claimed by the disclosure.

Claims

1. An impeller machine, characterized by comprising at least one blade group, wherein the blade group comprises a plurality of blades distributed along the circumferential direction of the impeller machine, each blade of the blade group comprises a blade body and a flange plate, and a mounting gap which extends along the span direction of the blade body in a zigzag mode is arranged between the flange plates of two adjacent blades.

2. The turbomachine of claim 1, wherein the rim plate includes at least one boss protruding in an axial direction of the turbomachine, and the mounting gap includes a seal segment between adjacent bosses of two adjacent blades, the seal segment extending zigzag in a span-wise direction of the blade body.

3. The turbomachinery of claim 2,

the at least one blade group comprises a guide blade group, each blade of the guide blade group is a guide blade, the blade body is a guide blade body (10) of the guide blade, the edge plate is a guide blade edge plate (11) of the guide blade, the mounting gap comprises a first mounting gap (G1), and the first mounting gap (G1) is formed between two adjacent guide blade edge plates (11) and extends along the span-wise zigzag of the guide blade body (10); and/or

The at least one blade group comprises a blade group, each blade of the blade group is a movable blade, the blade body is a movable blade body (20) of the movable blade, the flange plate is a movable blade flange plate (21) of the movable blade, the mounting gap comprises a second mounting gap (G2), and the second mounting gap (G2) is formed between two adjacent movable blade flange plates (21) and extends along the span-wise zigzag of the movable blade body (20).

4. The turbomachinery of claim 3,

the at least one boss of one of the stator vane platform (11) and the moving blade platform (21) comprises a first boss (11A) and a second boss (11B) arranged at intervals in the radial direction of the impeller machine, the at least one boss of the other of the stator vane platform (11) and the moving blade platform (21) comprises a third boss (21A) located between the first boss (11A) and the second boss (11B) in the radial direction of the impeller machine, and a sealing cavity is formed between the first boss (11A), the second boss (11B) and the third boss (21A).

5. The turbomachinery of claim 3,

the at least one protrusion of one of the stator vane platform (11) and the bucket platform (21) comprises a first protrusion (11A) and a second protrusion (11B) arranged at a radial distance of the turbomachine, the at least one protrusion of the other of the stator vane platform (11) and the bucket platform (21) comprises a third protrusion (21A) located radially of the turbomachine between the first protrusion (11A) and the second protrusion (11B), wherein,

one of the first mounting gap (G1) and the second mounting gap (G2) includes a first sealing section (G11) between two adjacent first bosses (11A) and a second sealing section (G12) between two adjacent second bosses (11B); and/or

The other of the first mounting gap (G1) and the second mounting gap (G2) includes a third obturating section (G21) between two adjacent third bosses (21A).

6. The turbomachinery of any one of claims 2 to 5, wherein the cross-section of at least one of the seal segments is stepped in shape.

7. The turbomachinery of any one of claims 2 to 5, wherein at least one of the seal segments has an S-shaped cross-section.

8. An aircraft engine, characterized in that it comprises a turbomachine according to any one of claims 1 to 7.