CN115898637A - Aeroengine splitter ring and aeroengine - Google Patents

Abstract

The invention relates to an aeroengine splitter ring and an aeroengine. Aeroengine splitter ring includes: the first annular wall comprises a first main body part, a bending part and a first connecting part, wherein the first end of the bending part is connected to the front end of the first main body part, the second end of the bending part extends to the rear end of the first main body part, and the first connecting part is arranged on the inner wall of the first main body part; the second annular wall and the first annular wall form an annular cavity, the second annular wall comprises a second main body part and a second connecting part, the front end of the second main body part is in butt joint with the second end of the bending part, the second connecting part is arranged on the outer wall of the second main body part, and the second connecting part is in butt joint with the first connecting part; the first connecting portion and the second connecting portion divide the annular cavity into a first cavity and a second cavity, a plurality of air holes are arranged between the first connecting portion and the second connecting portion, and the air holes are configured to guide air flow in the second cavity to the first cavity so that the air flow impacts the bending portion. The invention is used for improving the heat exchange coefficient and the heat exchange effect.

Description

Aeroengine splitter ring and aeroengine

Technical Field

The invention relates to the technical field of aerospace, in particular to a flow divider ring of an aero-engine and the aero-engine.

Background

Icing has great influence on flight safety, and particularly for an aircraft engine, once icing occurs, unsmooth air inlet of the engine is caused, and the working efficiency of the engine is reduced; if so, the engine is flamed out and even physically damaged, and serious safety accidents are caused. The physical element behind the fan that separates the inner and outer bypass flows of the aircraft engine splitter ring is typically the component that needs anti-icing.

Disclosure of Invention

Some embodiments of the invention provide an aircraft engine splitter ring and an aircraft engine, which are used for improving the anti-icing effect of the aircraft engine splitter ring.

In one aspect of the present invention, there is provided an aircraft engine splitter ring, comprising:

a first annular wall including a first main body portion, a curved portion and a first connecting portion, wherein a first end of the curved portion is connected to a front end of the first main body portion, a second end of the curved portion extends toward a rear end of the first main body portion, and the first connecting portion is provided on an inner wall of the first main body portion and extends toward the second annular wall; and

the annular cavity is formed between the second annular wall and the first annular wall, the second annular wall comprises a second main body part and a second connecting part, the front end of the second main body part is in butt joint with the second end of the bending part, the second connecting part is arranged on the outer wall of the second main body part and extends towards the first annular wall, and the second connecting part is located on the inner side of the first connecting part and is in butt joint with the first connecting part;

wherein the annular cavity is divided into a first cavity and a second cavity by the first connecting portion and the second connecting portion, a plurality of air holes are arranged between the first connecting portion and the second connecting portion, and the plurality of air holes are configured to guide the airflow in the second cavity to the first cavity so that the airflow impacts the bending portion.

In some embodiments, the front end of the second body portion is provided with a plurality of notches that cooperate with the second end of the curved portion to form a plurality of exhaust ports, each exhaust port of the plurality of exhaust ports being positioned to correspond to a position of a vane of an aircraft engine, the exhaust ports being configured to direct the flow of air within the first cavity toward the vane to which it corresponds.

In some embodiments, the second connecting portion is provided with a plurality of grooves, and the plurality of grooves and the first connecting portion cooperate to form the plurality of air holes.

In some embodiments, the junction of the first connection portion and the first body portion is at the junction of the first end of the curved portion and the front end of the first body portion.

In some embodiments, the connection of the second connecting portion and the second main body portion is located behind the front end of the second main body portion.

In some embodiments, the first annular wall further includes a third connecting portion, the second annular wall includes a fourth connecting portion, the third connecting portion is disposed at the rear end of the first main body portion and extends toward the second annular wall, the fourth connecting portion is disposed at the rear end of the second main body portion and extends toward the first annular wall, the fourth connecting portion is located at the inner side of the third connecting portion and abuts against the third connecting portion, and the fourth connecting portion is provided with an air guiding hole.

In some embodiments, the volume of the second cavity is greater than the volume of the first cavity.

In some embodiments, the thickness of the curved portion is uniform from the first end to the second end thereof, and the thickness of the curved portion is uniform with the thickness of the first body portion.

In one aspect of the invention, an aircraft engine is provided, which comprises the aircraft engine splitter ring.

In some embodiments, the aircraft engine further comprises a plurality of vanes, the front end of the second body portion is provided with a plurality of notches that cooperate with the second end of the curved portion to form a plurality of exhaust ports, each exhaust port of the plurality of exhaust ports is positioned to correspond to a position of a vane of the plurality of vanes, and the exhaust port is configured to direct the flow of air within the first chamber to the vane to which it corresponds.

Based on the technical scheme, the invention at least has the following beneficial effects:

in some embodiments, the aeroengine flow distribution ring comprises a first annular wall and a second annular wall, an annular cavity is formed between the first annular wall and the second annular wall, the annular cavity is divided into a first cavity and a second cavity by a first connecting part and a second connecting part, the first cavity and the second cavity are communicated through a plurality of air holes, the anti-icing hot air is firstly introduced into the second cavity and then introduced into the first cavity through the plurality of air holes, so that the air flow speed entering the first cavity is improved, the effect of impacting the front edge wall surface of the flow distribution ring is achieved, the heat exchange coefficient and the heat exchange effect are improved, on the premise of ensuring the anti-icing effect, the air entraining amount of the anti-icing hot air is reduced, and the utilization rate of the anti-icing hot air is improved.

Drawings

The accompanying drawings, which are included to provide a further understanding of the invention and are incorporated in and constitute a part of this application, illustrate embodiment(s) of the invention and together with the description serve to explain the invention without limiting the invention. In the drawings:

FIG. 1 is a schematic partial structural view of an aircraft engine splitter ring and vanes provided in accordance with some embodiments of the present invention;

FIG. 2 is a schematic partial structural view of a first annular wall of an aircraft engine diverter ring provided in accordance with some embodiments of the present invention;

FIG. 3 is a schematic partial structural view of a second annular wall of an aircraft engine diverter ring provided in accordance with some embodiments of the present invention;

FIG. 4 is a schematic illustration of the flow direction within an aircraft engine splitter ring provided in accordance with some embodiments of the invention;

FIG. 5 is a schematic illustration of a partial structure of an aircraft engine provided in accordance with some embodiments of the invention.

The reference numbers in the drawings illustrate the following:

1-a first annular wall; 11-a first body portion; 12-a bend; 121-a second end of the bend; 13-a first connection; 14-a third connecting portion;

2-a second annular wall; 21-a second body portion; 211-the front end of the second body portion; 212-a notch; 22-a second connection; 221-grooves; 23-a fourth connection;

3-air holes;

4-an exhaust port;

5-air guiding holes;

61-a first cavity; 62-a second cavity;

100-a shunt ring; 200-guide vanes; 300-a nacelle; 400-support plate; 500-a valve; 600-a gas compressor; 700-a combustion chamber; 800-turbine.

It should be understood that the dimensions of the various parts shown in the figures are not drawn to scale. Further, the same or similar reference numerals denote the same or similar components.

Detailed Description

Various exemplary embodiments of the present invention will now be described in detail with reference to the accompanying drawings. The description of the exemplary embodiments is merely illustrative and is in no way intended to limit the invention, its application, or uses. The present invention may be embodied in many different forms and is not limited to the embodiments described herein. These embodiments are provided so that this disclosure will be thorough and complete, and will fully convey the scope of the invention to those skilled in the art. It should be noted that: the relative arrangement of parts and steps, the composition of materials, numerical expressions and numerical values set forth in these embodiments are to be construed as merely illustrative, and not restrictive, unless specifically stated otherwise.

The use of "first," "second," and similar terms in the present application do not denote any order, quantity, or importance, but rather the terms are used to distinguish one element from another. The word "comprising" or "comprises", and the like, means that the element preceding the word covers the element listed after the word, and does not exclude the possibility that other elements are also covered. "upper", "lower", "left", "right", and the like are used only to indicate relative positional relationships, and when the absolute position of the object being described is changed, the relative positional relationships may also be changed accordingly.

In the present invention, when it is described that a specific device is located between a first device and a second device, there may or may not be an intervening device between the specific device and the first device or the second device. When a particular device is described as being coupled to other devices, that particular device may be directly coupled to the other devices without intervening devices or may be directly coupled to the other devices with intervening devices.

All terms (including technical and scientific terms) used herein have the same meaning as commonly understood by one of ordinary skill in the art to which this invention belongs unless specifically defined otherwise. It will be further understood that terms, such as those defined in commonly used dictionaries, should be interpreted as having a meaning that is consistent with their meaning in the context of the relevant art and will not be interpreted in an idealized or overly formal sense unless expressly so defined herein.

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.

The physical element behind the fan that separates the air flow in the interior and the exterior ducts of an aircraft engine splitter ring is typically a component that requires anti-icing. The anti-icing bleed air of the splitter ring of the aircraft engine plays an important role in preventing the splitter ring from being frozen. If the splitter ring freezes, the safety and the performance of the engine are greatly influenced, and the working stability of the engine is influenced. Excessive use of hot gas from the high-pressure compressor for anti-icing also affects engine performance.

Based on this, some embodiments of the present disclosure provide an aeroengine splitter ring for improving the anti-icing effect of the aeroengine splitter ring.

In the embodiments of the present disclosure, "front" and "rear" are defined in terms of the flow direction of the airflow outside the splitter ring, with "front" being upstream of the airflow outside the splitter ring and "rear" being downstream of the airflow outside the splitter ring. The central axis of the aircraft engine is defined as an inner side and an outer side, one side close to the central axis of the aircraft engine is the inner side, and one side far away from the central axis of the aircraft engine is the outer side.

The air flow outside the splitter ring contains water drops, ice crystals and the like, the temperature is low, the splitter ring is easy to freeze, and the anti-icing hot gas exchanges heat with the splitter ring by introducing the anti-icing hot gas into the splitter ring, so that the splitter ring is prevented from freezing.

As shown in fig. 1, in some embodiments, an aircraft engine diverter ring 100 includes a first annular wall 1 and a second annular wall 2.

As shown in fig. 2, the first annular wall 1 includes a first main body portion 11, a bent portion 12, and a first connecting portion 13.

A first end of the bending portion 12 is connected to a front end of the first body portion 11, and optionally, the first end of the bending portion 12 is integrally formed with the front end of the first body portion 11. The second end 121 of the curved portion 12 extends toward the rear end of the first body portion 11. The curved portion 12 is the leading edge component of the diverter ring 100.

Alternatively, the bending portion 12 is U-shaped, the opening of the bending portion 12 faces the rear end of the first body portion 11, and the first end and the second end of the bending portion 12 are both ends of the U-shaped opening.

The first connecting portion 13 is provided on the inner wall of the first body portion 11 and extends toward the second annular wall 2.

The second annular wall 2 is connected to the first annular wall 1, and an annular cavity is formed between the second annular wall 2 and the first annular wall 1. The second annular wall 2 is close to the median axis of the aircraft engine with respect to the first annular wall 1. That is, the first annular wall 1 is located inside the second annular wall 2.

As shown in fig. 3, the second annular wall 2 includes a second main body portion 21 and a second connecting portion 22.

As shown in fig. 1 and 3, the front end 211 of the second body portion 21 abuts against the second end 121 of the bending portion 12.

The second connection portion 22 is provided on the outer wall of the second body portion 21 and extends toward the first annular wall 1, and the second connection portion 22 is located inside the first connection portion 13 and abuts against the first connection portion 13.

As shown in fig. 1, in which the first and second connecting portions 13 and 22 divide the annular chamber into a first chamber 61 and a second chamber 62, the bent portion 12 is located in the first chamber 61. A plurality of air vents 3 are provided between the first connection portion 13 and the second connection portion 22, and the plurality of air vents 3 are configured to guide the air flow in the second chamber 62 toward the first chamber 61 so that the air flow impacts the bent portion 12.

Alternatively, the plurality of air holes 3 are provided in a circle between the first connection portion 13 and the second connection portion 22. The plurality of air holes 3 may be uniformly distributed or periodically distributed.

In the above embodiment, the first connecting portion 13 and the second connecting portion 22 separate the ring cavity into the first cavity 61 and the second cavity 62, the bending portion 12 is located in the first cavity 61, that is to say, the first cavity 61 is close to the front end of the flow distribution ring 100, the first cavity 61 and the second cavity 62 are communicated through the plurality of air holes 3, the anti-icing hot gas is firstly introduced into the second cavity 62, and then is introduced into the first cavity 61 through the plurality of air holes 3, so that the air velocity entering the first cavity 61 is increased, the impact bending portion 12 is played, that is, the effect of impacting the front edge wall surface of the flow distribution ring 100 is achieved, the heat exchange coefficient and the heat exchange effect are improved, on the premise that the anti-icing effect is ensured, the air entrainment of the anti-icing hot gas is reduced, and the utilization rate of the anti-icing hot gas is improved.

In some embodiments, the volume of the second cavity 62 is greater than the volume of the first cavity 61. The air current that second chamber 62 introduced leads to first chamber 61 through a plurality of gas pockets 3, the gas volume syneresis, the air speed improves, because the air speed is very fast, therefore, when the air current jets into first chamber 61, flexion 12 can be strikeed, form the structure of an impact heat transfer, this kind of impact structure can make the speed of air current improve bigger, the air disturbance in first chamber 61 has been strengthened, convection current heat transfer ability has been improved, the reinforcing heat transfer effect, under the prerequisite of guaranteeing anti-icing effect, reduce the bleed air volume of anti-icing steam, the utilization ratio of anti-icing steam is improved.

Since the guide vanes 200 inside the splitter ring 100 are also prone to icing, and the icing of the guide vanes 200 causes a large blockage area, the guide vanes 200 need to be anti-iced.

In some embodiments, the front end 211 of the second body portion 21 is provided with a plurality of notches 212, the plurality of notches 212 cooperates with the second end 121 of the curved portion 12 to form a plurality of exhaust ports 4, the position of each exhaust port 4 of the plurality of exhaust ports 4 is configured to correspond to the position of a vane 200 of the aircraft engine, and the exhaust ports 4 are configured to direct the airflow within the first cavity 61 towards its corresponding vane 200.

Because gas vent 4 guides the air current in first chamber 61 to its stator 200 that corresponds, every gas vent 4 corresponds a stator 200, and 4 exhaust gas of gas vent carry out steam gas mould to stator 200 and cover the processing, play the purpose of reutilization anti-icing steam, under the prerequisite of guaranteeing anti-icing effect, reduce the bleed volume of anti-icing steam, improve the utilization ratio of anti-icing steam.

In some embodiments, as shown in fig. 4, the airflow in the second cavity 62 is accelerated through the plurality of air holes 3 and then flows into the first cavity 61, and impacts the bending portion 12 to heat the front edge of the splitter ring 100, which is a first utilization of the anti-icing hot gas, and then the airflow in the first cavity 61 is discharged through the plurality of air outlets 4, each air outlet 4 corresponds to one guide vane 200, and the airflow in the first cavity 61 can be intensively impacted to each guide vane 200 through each air outlet 4, so that the airflow flow is high, the flow speed is high, the hot gas-mold covering treatment can be performed on the guide vane 200 to prevent the guide vane 200 from icing, which is a second utilization of the anti-icing hot gas, and on the premise of ensuring the anti-icing effect, the air entrainment amount of the anti-icing hot gas is reduced, and the utilization rate of the anti-icing hot gas is improved.

As shown in fig. 3, in some embodiments, the second connecting portion 22 is provided with a plurality of grooves 221, and the plurality of grooves 221 cooperate with the first connecting portion 13 to form a plurality of air holes 3, as shown in fig. 1. The groove 221 is formed in the second connecting portion 22, and the machining process is simple and easy to achieve.

Optionally, the groove comprises a semi-circular groove.

As shown in fig. 1 and 2, in some embodiments, the connection of the first connection portion 13 to the first main body portion 11 is located at the connection of the first end of the bent portion 12 and the front end of the first main body portion 11.

As shown in fig. 1 and 3, in some embodiments, the connection point of the second connecting portion 22 and the second main body portion 21 is located behind the front end 211 of the second main body portion 21.

The curved portion 12, the first connection portion 13, the second connection portion 22, and the front end 211 of the second body portion 21 form a first cavity 61.

In some embodiments, the first annular wall 1 further comprises a third connection 14. The third connecting portion 14 is provided at the rear end of the first body portion 11 and extends toward the second annular wall 2.

The second annular wall 2 comprises a fourth connection 23. The fourth connecting portion 23 is provided at the rear end of the second body portion 21 and extends toward the first annular wall 1.

The fourth connecting portion 23 is located inside the third connecting portion 14 and abuts against the third connecting portion 14, and the fourth connecting portion 23 is provided with the air vent 5.

The introduced anti-icing hot gas enters the second cavity 61 from the air guide holes 5, then enters the first cavity 61 through the plurality of air holes 3, impacts the bending part 12, performs anti-icing on the splitter ring 100, and finally flows to the guide vane 200 through the plurality of air exhaust holes 4 to perform anti-icing on the guide vane 200.

In some embodiments, the thickness of the bending portion 12 from the first end to the second end thereof is uniform, and the thickness of the bending portion 12 is uniform to the thickness of the first body portion 11.

The bending portion 12 is a front edge of the shunt ring 100, and a thickness of a bending point of the bending portion 12 is generally thicker than thicknesses of a first end and a second end of the bending portion 12, so as to increase a strength of the front edge of the shunt ring 100, in the embodiment of the present disclosure, the first cavity 61 is disposed at the bending portion 12, a thickness of the bending portion 12 from the first end to the second end thereof is uniform, the thickness of the bending portion 12 is consistent with that of the first main body portion 11, the thickness of the bending portion 12 is not increased, a space can be increased for the first cavity 61, the thickness of the front edge of the shunt ring 100 is reduced, an air flow heat exchange coefficient can be increased, ice prevention of ice-proof hot air on the front edge of the shunt ring 100 is facilitated, and the connection between the first connection portion 13 and the second connection portion 22 can play a supporting role, so as to enhance the strength of the front edge of the shunt ring 100, and the strength of the shunt ring 100 cannot be affected by the thinness of the bending portion 12.

Some embodiments also provide an aircraft engine that includes the aircraft engine diverter ring 100 described above.

In some embodiments, the aircraft engine further includes a plurality of vanes 200, the front end of the second body portion 21 is provided with a plurality of notches 212, the plurality of notches 212 and the second end 121 of the curved portion 12 cooperate to form a plurality of exhaust ports 4, a position of each exhaust port 4 of the plurality of exhaust ports 4 is configured to correspond to a position of a vane 200 of the plurality of vanes 200, and the exhaust port 4 is configured to direct the airflow within the first cavity 61 toward its corresponding vane 200.

Alternatively, one exhaust port 4 corresponds to one guide vane 200, or a plurality of exhaust ports 4 correspond to one guide vane 200.

In some embodiments, the volume of the second cavity 62 is greater than the volume of the first cavity 61. Set up an anti-icing loculus at splitter ring 100's front end, can make the airstream volume reduce, increase the air current velocity of flow, introduce the loculus through a plurality of gas pockets 3 with the air current in loculus simultaneously, can further improve the air current velocity of flow, make the air current strike flexion 12 after getting into the loculus, improve the heat transfer effect, and then improve splitter ring 100's anti-icing effect, the air current in the loculus passes through gas vent 4 and leads to stator 200, prevent ice to stator 200, the utilization ratio of steam has been improved, it is too big to have alleviated splitter ring anti-icing bleed demand under certain operating mode, and stator icing scheduling problem.

As shown in fig. 5, in some embodiments, the aircraft engine further includes, in addition to the splitter ring 100 and the guide vane 200, a nacelle 300, a strut 400, a bleed air pipe, a valve 500, a compressor 600, a combustion chamber 700, and a turbine 800, as shown in fig. 1 and 4, the bleed air pipe leads anti-icing hot gas from the compressor 500, the valve 500 is disposed in the bleed air pipe, the bleed air pipe is connected to the bleed air hole 5, the valve 500 opens the valve 500, the anti-icing hot gas enters the bleed air pipe from the compressor 500, enters the second cavity 62 of the splitter ring 100 through the bleed air hole 5, then enters the first cavity 61 through the plurality of air holes 3, and finally flows to the guide vane 200 through the exhaust port 4, so as to enhance heat exchange, thereby achieving hot gas anti-icing of the splitter ring and the guide vane.

The structure of the splitter ring 100 provided by the embodiment of the present disclosure improves the heat convection capability and uniformity of the inner wall surface of the splitter ring 100, reduces the amount of demand of the anti-icing hot gas, and simultaneously performs anti-icing on the guide vane 200, thereby improving the utilization rate of the anti-icing hot gas.

Based on the embodiments of the invention described above, the technical features of one of the embodiments can be advantageously combined with one or more other embodiments without explicit negatives.

Although some specific embodiments of the present invention have been described in detail by way of illustration, it should be understood by those skilled in the art that the above illustration is only for the purpose of illustration and is not intended to limit the scope of the invention. It will be understood by those skilled in the art that various changes may be made in the above embodiments or equivalents may be substituted for elements thereof without departing from the scope and spirit of the invention. The scope of the invention is defined by the appended claims.

Claims (10)

1. An aircraft engine diverter ring, comprising:

the first annular wall (1) comprises a first main body part (11), a bending part (12) and a first connecting part (13), wherein the first end of the bending part (12) is connected to the front end of the first main body part (11), the second end (121) of the bending part (12) extends towards the rear end of the first main body part (11), and the first connecting part (13) is arranged on the inner wall of the first main body part (11) and extends towards the second annular wall (2); and

a second annular wall (2) forming an annular cavity with the first annular wall (1), wherein the second annular wall (2) comprises a second main body part (21) and a second connecting part (22), the front end (211) of the second main body part (21) is butted with the second end (121) of the bending part (12), the second connecting part (22) is arranged on the outer wall of the second main body part (21) and extends towards the first annular wall (1), and the second connecting part (22) is positioned on the inner side of the first connecting part (13) and is butted with the first connecting part (13);

wherein the first and second connections (13, 22) divide the annular chamber into a first chamber (61) and a second chamber (62), a plurality of air holes (3) being provided between the first and second connections (13, 22), the plurality of air holes (3) being configured to direct an air flow within the second chamber (62) towards the first chamber (61) such that the air flow impinges on the curved portion (12).

2. The aircraft engine diverter ring according to claim 1, wherein the front end (211) of the second body portion (21) is provided with a plurality of notches (212), the plurality of notches (212) cooperating with the second end (121) of the bend (12) to form a plurality of exhaust ports (4), the position of each exhaust port (4) of the plurality of exhaust ports (4) being configured to correspond to the position of a vane (200) of an aircraft engine, the exhaust port (4) being configured to direct the flow of air within the first cavity (61) towards the vane (200) to which it corresponds.

3. The aircraft engine diverter ring according to claim 1, wherein the second connection portion (22) is provided with a plurality of grooves (221), the plurality of grooves (221) cooperating with the first connection portion (13) to form the plurality of air holes (3).

4. The aircraft engine diverter ring according to claim 1, wherein the connection of the first connection portion (13) to the first main body portion (11) is at the connection of the first end of the curved portion (12) to the front end of the first main body portion (11).

5. The aircraft engine diverter ring according to claim 1, wherein the junction of the second connecting portion (22) and the second body portion (21) is located aft of a forward end (211) of the second body portion (21).

6. The aircraft engine diverter ring according to claim 1, wherein the first annular wall (1) further comprises a third connecting portion (14), the second annular wall (2) comprises a fourth connecting portion (23), the third connecting portion (14) is provided at a rear end of the first main body portion (11) and extends toward the second annular wall (2), the fourth connecting portion (23) is provided at a rear end of the second main body portion (21) and extends toward the first annular wall (1), the fourth connecting portion (23) is located inside the third connecting portion (14) and abuts against the third connecting portion (14), and the fourth connecting portion (23) is provided with an air bleed hole (5).

7. The aircraft engine diverter ring according to claim 1, wherein the volume of the second chamber (62) is greater than the volume of the first chamber (61).

8. The aircraft engine diverter ring according to claim 1, wherein the thickness of the curved portion (12) is uniform from the first end to the second end thereof, the thickness of the curved portion (12) corresponding to the thickness of the first body portion (11).

9. An aircraft engine including an aircraft engine diverter ring according to any one of claims 1 to 8.

10. The aircraft engine according to claim 9, comprising a plurality of vanes (200), wherein the front end of the second body portion (21) is provided with a plurality of notches (212), wherein the plurality of notches (212) cooperate with the second end (121) of the curved portion (12) to form a plurality of exhaust ports (4), wherein the position of each exhaust port (4) of the plurality of exhaust ports (4) is configured to correspond to the position of a vane (200) of the plurality of vanes (200), and wherein the exhaust port (4) is configured to direct the air flow within the first cavity (61) towards the vane (200) to which it corresponds.

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