BE1030724B1 - TURBOMACHINE ASSEMBLY - Google Patents

Abstract

The invention relates to an assembly (4; 40) for a turbomachine (1) comprising a wall (6) having a guiding surface (8) of an air flow (F1) and a row of stator vanes ( 10; 100) extending from the wall and arranged annularly around an axis (A), said surface being inclined relative to the axis, the assembly being remarkable in that the row of blades comprises first blades (12) extending substantially perpendicular to the axis and second vanes (14) extending substantially perpendicular to the guide surface. The invention also relates to an axial turbomachine comprising a low pressure compressor, a high pressure compressor, and a swan neck vein arranged axially between the low pressure compressor and the high pressure compressor, the turbomachine being remarkable in that the vein contains a set as described above.

Description

Description

TURBOMACHINE ASSEMBLY

Domain

The invention relates to the field of turbomachines and more particularly to the field of axial turbomachine compressors.

Prior art

An axial turbomachine generally comprises two compressors arranged upstream of a combustion chamber, namely a low pressure compressor and a high pressure compressor configured to suck in and compress the air in order to bring it to speeds, pressures and temperatures suitable for good combustion.

In this regard, each compressor comprises a succession of compression stages, each being formed by at least one annular row of rotor blades and at least one annular row of stator blades.

The last row of stator blades (which may be commonly called OGV “Outlet Guide Vane”) is directly followed downstream by structural arms (or “struts”) which cross the air flow path of the axial turbomachine.

The structural arms are generally arranged in a sloping vein (or swan neck between the compressors), i.e. the direction of the air flow circulating in the vein being inclined relative to a longitudinal axis of the axial turbomachine. Document FR 3 027 053 describes an example of such a type of structural arm.

The OGV blades usually have a direction substantially perpendicular to the direction of the flow to limit aerodynamic disturbances which would make the flow unstable to the right of the OGV.

With such an orientation of the OGV blades, the few blades which are circumferentially directly in front of the structural arms have a radially outer part which is close to the structural arms. The proximity between the trailing edge of the OGV and the leading edge of the arms can cause a distortion of the static pressure at the trailing edge of the OGV, which creates a loss of efficiency for the turbomachine.

In order to overcome the problem of static pressure distortion, there is a solution consisting of moving POGV axially away from the structural arms. This solution is not reasonable because it increases the axial length of the turbomachine and therefore its total weight. There is therefore room for improvement to improve the performance of the compressor without affecting its dimensions.

Summary of the invention

Technical problem

The invention aims to resolve the drawbacks of the design of turbomachine compressors of the state of the art. In particular, the invention aims to propose a solution which makes it possible to limit the aerodynamic disturbances at the level of the stator blades upstream of the structural arms while limiting the phenomenon of distortion of the static pressure, in particular in a sloping vein, and this, without complexity and/or addition of additional weight to the turbomachine compressor.

Technical solution

The invention relates to an assembly for a turbomachine comprising a wall having a surface for guiding an air flow and a row of stator vanes extending from the wall and arranged annularly around an axis, said surface being inclined relative to the axis, the assembly being remarkable in that the row of blades comprises first blades extending substantially perpendicular to the axis and second blades extending substantially perpendicular to the guide surface.

Preferably, the inclination of the guide surface relative to the axis results in the air flow comprising the same inclination. In this configuration, the wall of the turbomachine has a curved profile on a longitudinal section and forms an angle of inclination with the longitudinal axis of the turbomachine.

By “row of blades extending annularly”, we mean a single ring of blades arranged circumferentially one after the other. These vanes delimit inter-vane spaces traversed by the air flow.

According to an advantageous embodiment of the invention, the assembly further comprises structural arms axially arranged directly downstream of the row of blades, inter-arm spaces being defined between two circumferentially adjacent structural arms, the first blades circumferentially overlapping the structural arms and the second blades circumferentially overlapping the inter-arm spaces.

According to an advantageous embodiment of the invention, the row of blades is formed by groups of first blades and by groups of second blades, said groups of first and second blades alternating circumferentially.

Advantageously, the circumferential alternation of the groups of first and second blades makes it possible to minimize aerodynamic disturbances by favoring the guidance of the air flow instead of the trailing edge of the row of blades.

In addition, the manufacturing of such a row is simplified, as well as its assembly, in particular with angular sectors of the turbomachine.

According to an advantageous embodiment of the invention, each group of first blades comprises between 3 and 8 blades, preferably 4 blades.

According to an advantageous embodiment of the invention, each group of second blades comprises between 4 and 10 blades, preferably 5 blades.

According to an advantageous embodiment of the invention, the row of blades further comprises a plurality of third blades arranged circumferentially regularly between the first and second blades, said third blades having a respective orientation which is included, limits excluded, between the orientation perpendicular to the axis of the first blades and the orientation perpendicular to the guide surface of the second blades.

Advantageously, the third blades make it possible to further smooth the air flow to the right of the trailing edge of the row of blades, thus making it possible to promote the regularity and progressiveness of said air flow in the vein.

According to an advantageous embodiment of the invention, the respective orientation of the third blades varies progressively in accordance with the circumferential position of the third blades relative to the first and second blades, the third blades having an orientation even closer to that of the first blades. that they are close circumferentially, and the third blades having an orientation all the closer to that of the second blades that they are close circumferentially.

According to an advantageous embodiment of the invention, the third blades have a respective orientation angle which varies step by step by at least 0.5°.

According to an advantageous embodiment of the invention, seen in a radial section, the progression of the orientation of the third blades has a triangular profile.

According to an advantageous embodiment of the invention, seen in a radial section, the progression of the orientation of the third blades has a sinusoidal profile.

According to an advantageous embodiment of the invention, the trailing edge of the first blades has a first distance following the direction of the air flow with the leading edge of the structural arms, and the trailing edge of the second blades has a second distance following the direction of the air flow with the leading edge of the structural arms, said first distance being greater than the second distance.

According to an advantageous embodiment of the invention, all the blades have the same chord length. Preferably, the blades have the same average chord over their entire span. This helps to control any distortion of static pressure, particularly by limiting it across the entire span of the leading edges of the structural arms.

According to an advantageous embodiment of the invention, the air flow guiding surface forms an angle of inclination with the axis of between 10° and 60°.

The invention also relates to an axial turbomachine comprising a low pressure compressor, a high pressure compressor, and a swan neck vein arranged axially between the low pressure compressor and the high pressure compressor, the turbomachine being remarkable in that the vein contains an assembly according to the invention.

Advantages of the invention

The invention is particularly advantageous in that it makes it possible to limit the aerodynamic disturbances of the stator blades in the sloping air flow vein, this is in particular due to the stator blades having either a radial direction or a direction perpendicular to the axis of the turbomachine, thus improving the overall stability of the annular row of stator blades.

In addition, the phenomenon of distortion of the static pressure is reduced by the fact that the radial stator vanes are moved away from the leading edge of the structural arms 5, and this, without modification of the axial position of the row of stator vanes in the air flow vein.

A good compromise is therefore obtained to improve the flow both downstream of the radial blades and downstream of the inclined blades.

In this configuration, the pressure losses are limited and the operability of the compressor is increased, this advantageously makes it possible to improve the behavior of the compressor.

Description of the designs

Figure 1 represents a simplified partial view of an axial turbomachine;

Figure 2 is a schematic illustration of a longitudinal sectional view of an axial turbomachine stream comprising an assembly according to the invention;

Figure 3 schematically represents a view in a radial section of the assembly of Figure 1 and according to a first embodiment of the invention;

Figure 4 schematically represents a view in a radial section of the assembly of Figure 2 and according to a second embodiment of the invention.

Description of an embodiment

In the description which follows, the terms “internal”, “interior”, “external” and “exterior” refer to positioning relative to the longitudinal axis of rotation of a turbomachine. The axial direction corresponds to the direction along the longitudinal axis of rotation of the turbomachine. The radial direction is perpendicular to the longitudinal axis. The annular or circumferential direction is essentially a circular direction around the longitudinal axis. Upstream and downstream refer to the direction of flow of an axial air flow in a main stream of the turbomachine.

The figures show the elements schematically and are not represented to scale. Indeed, the figures have been intentionally simplified in order to facilitate understanding and certain dimensions are enlarged to facilitate their reading.

Figure 1 represents a simplified partial view of an axial turbomachine 1. In this specific case it is a double-flow turbojet, but can also be a turbojet, turbofan, turboprop, turboshaft or any other turbomachine.

Those skilled in the art will understand that the invention can also be applied to a turbomachine with a centrifugal compressor.

The turbomachine 1 comprises a first compression level, called low pressure compressor 3, a second compression level, called high pressure compressor 5 (partially shown), as well as a combustion chamber and one or more turbine levels (not shown) .

In operation, the mechanical power of the turbine(s) transmitted via central shaft to the rotor sets in motion the rotors of the two compressors 3 and 5. The rotation of the rotor around its longitudinal axis of rotation À thus makes it possible to generate a flow of air and to gradually compress the latter until it enters the combustion chamber.

An inlet fan commonly referred to as a fan or blower is coupled to the rotor and generates an incoming air flow F. The fan is arranged upstream of a separation nozzle 11 capable of separating the incoming air flow F entering into a radially internal air flow, called air flow F1 or primary flow F1, circulating in a primary flow vein 2 and crossing the different above-mentioned levels of the turbomachine 1, and a radially external air flow, called secondary flow F2 passing through an annular conduit along the machine to then join the primary flow F1 at the turbine outlet. The secondary flow F2 can be accelerated so as to generate a thrust reaction necessary for the flight of an aircraft.

Preferably, the vein 2 has a swan neck shape, commonly called a sloping vein, and preferably, the low pressure compressor 3 comprises at its downstream half an assembly 4, 40 which will be amply detailed in the present description.

Figure 2 is a schematic illustration of a longitudinal sectional view of the vein 2 of an axial turbomachine comprising an assembly 4, 40.

With reference to Figure 2, the vein 2 allows the guidance of the primary flow F1 between an exterior wall 6 having an exterior guide surface 8 of the air and an interior wall 9 having an interior guide surface 7.

The assembly 4, 40 comprises a row of stator vanes 10, 100, also called stator grid or OGV, said row of stator vanes 10,100 extends from the outer wall 6 to the inner wall 9, and extends from the outer wall 6 to the inner wall 9, and extends annularly around the longitudinal axis A of the axial turbomachine. The second guide wall 9 can for example be an internal ferrule 9.

Preferably, the height of the vein 2, i.e. radial distance between the guide surface 8 and the internal ferrule 9, is constant instead of the assembly 4, 40.

The wall 6, and in particular the guide surface 8, is inclined relative to the axis A.

In this regard, a straight line tangent to the guide surface 8 to the right of the row of stator blades 10, 100 presents with the axis At an angle of inclination a between 5° and 80°, and preferably between 10 ° and 60°. The inclination following the angle a results in an inclined guidance of the air following a downward slope, resulting in an overall inclination of the direction of the primary flow F1 relative to the axis A to the right of the assembly 4, 40.

The row of stator blades 10, 100 can belong for example to the low pressure compressor of the axial turbomachine, and said row of stator blades 10, 100 is made up of: - blades denoted 12 and hereinafter called "first blades", which extend substantially perpendicular to axis A; and - blades denoted 14 and hereinafter called "second blades", which extend substantially perpendicular to the guide surface 8.

Preferably, the first 12 and second blades 14 have the same chord length and/or Average profile chord length, particularly in the case where the chord varies over the total extent of each blade.

The second blades 14 have an angle of inclination with the axis A corresponding to the angle a + 90°.

The assembly 4, 40 further comprises structural arms 16 axially arranged directly downstream of the row of stator vanes 10, 100.

Following the direction of the air flow F1, and preferably following a projection on the guide surface 8, the trailing edge 13 of the first blades 12 has a first distance D with the leading edge 17 of the structural arms 16, and the trailing edge 15 of the second blades 14 has a second distance d with said leading edge 17. Here we give “distance” its usual definition, that is to say the shortest distance between the two bodies.

Preferably, the trailing edge 13 of the first blades 12 is axially at the same level as the trailing edge 15 of the second blades 14, to the right of the internal shroud 9, and more precisely to the right of the point P illustrated in Figure 2. For this purpose, the trailing edge 15 of the second blades 14 does not axially exceed the trailing edge 13 of the first blades 12 in the direction of circulation of the air flow F1.

The first distance D is greater than the second distance d.

Advantageously, this makes it possible to avoid the axial distance of the entire row of blades 10, 100 relative to the leading edge 17 of the structural arms 16 while preventing the radially external part of the first blades from being too close. arms. The distortion of the flow can be limited by avoiding the phenomenon of loss of static pressure at the level of the structural arms 16.

Figure 3 schematically represents a view in a radial section of assembly 4 of Figure 2 and according to a first embodiment of the invention.

With reference to Figure 3, the structural arms 16 delimit inter-arm spaces 18 between each two circumferentially adjacent structural arms 16.

In this configuration, a group 12' formed by a plurality of first blades 12, called first group 12', circumferentially overlaps the structural arms 16, while a second group 14' formed by second blades 14, circumferentially overlaps the interspaces. -arm 18, the latter being partly covered by the second blades 14.

In this regard, the row of stator blades 10 is formed by the first group 12' and second group 14', the latter alternate circumferentially (that is to say according to the horizontal orientation of Figure 3) around the axis of the turbomachine.

Preferably, and in order to maximize the stability instead of the blades of the row of stator blades 10, the first group 12' comprises between 2 and 10 blades, and preferably between 3 and 8 blades, and even more preferably 4 blades, the second group 14' comprises between 2 and 12 blades, and preferably between 4 and blades, and even more preferably 5 blades.

We see in particular in Figure 3 that the second blades 14 can be closer axially (in a projection on a vertical axis of Figure 3) to the arms 16 than are the first blades 12.

Figure 4 schematically represents a view in a radial section of the assembly 40 of Figure 2 and according to a second embodiment of the invention.

Indeed, Figure 4 particularly represents the schematization of a progression of orientation of the stator blades 100 of Figure 2.

With reference to Figure 4, the row of stator vanes 40 comprises, in addition to the first 12 and second vanes 14, a plurality of third vanes 30 arranged circumferentially between the first 12 and second vanes 14.

It should be noted that the third blades 30 are not visible and therefore are not illustrated in Figure 2.

The third blades 30 have a respective orientation which is included, limits excluded, between the orientation of the first blades 12 and the orientation of the second blades 14.

The respective orientation of the third blades 30 varies gradually and regularly in accordance with the circumferential position of the third blades 30 relative to the first 12 and second blades 14,

In this configuration, the progressive orientation of the stator vanes 100 is mainly due to the progressive orientation of the third vanes 30 between the extreme terminals, i.e. the first 12 and second vanes 14.

In this regard, the third blades 30 have an orientation that is all the closer to that of the first blades 12 as they are circumferentially close to it. Indeed, the closer the third blade 30 is annularly to the first blade 12,

and the closer the orientation of the third blade 30 will be to the inclination of the first blade 12.

Similarly, the third blades 30 have an orientation that is all the closer to that of the second blades 14 as they are circumferentially close to it.

Preferably, the third blades 30 have a respective orientation angle which varies step by step by at least 0.5° and/or the variation in the angle of the third blades 30 is preferably less than 20°.

The variation in orientation of the third blades is regular and is therefore linked to the number of third blades.

The progression of the orientation of the third blades 30 has a triangular 22 and/or sinusoidal 20 profile, such an orientation progression profile makes it possible to further smooth the air flow and avoid aerodynamic disturbances (for example due to corner effects or sudden variations in the flow guiding surfaces).

In summary, the difference between the first and second embodiment of the invention results in the circumferential evolution of the inclination of the blades forming the stator grid.

All the blades, according to the first and second embodiment of the invention, and particularly those being circumferentially adjacent to the structural arms, can have a wedging or an orientation, also called "tilting", and which makes it possible to have a profile of different blades at the level of the stator grid in order to homogenize as much as possible the pressure value at the leading edge of the blades forming said stator grid. The tilting can be adjusted independently for each blade.

Furthermore, the elimination of the phenomenon of distortion of the static pressure at the leading edge of the blades of the stator grid can also be obtained by further circumferentially spacing the blades located annularly adjacent to the structural arms, said spacing concerns all the blades of the first or second embodiment.

It should be noted that the invention is not limited to the examples described in the figures. The teachings of the present invention may in particular be applicable to another type of turbomachine, and each technical characteristic of each illustrated example is applicable to the other examples.

For example, a row of stator blades can result from the combination of a portion of a row in Figure 3, with a portion of a row 100 in Figure 4.

Claims (14)

Claims 1. Assembly (4; 40) for a turbomachine (1) comprising a wall (6) having a guiding surface (8) of an air flow (F1) and a row of stator vanes (10; 100) s extending from the wall (6) and arranged annularly around an axis (A), said surface (8) being inclined relative to the axis (A), the assembly (4; 40) being characterized in that the row of blades (10; 100) comprises first blades (12) extending substantially perpendicular to axis (A) and second blades (14) extending substantially perpendicular to the guide surface (8). 2. Assembly (4; 40) according to claim 1, characterized in that it further comprises structural arms (16) axially arranged directly downstream of the row of blades (10; 100), inter-arm spaces (18) being defined between two circumferentially adjacent structural arms (16), the first vanes (12) circumferentially overlapping the structural arms (16) and the second vanes (14) circumferentially overlapping the inter-arm spaces (18). 3. Assembly (4) according to one of claims 1 or 2, characterized in that the row of blades (10) is formed by groups of first blades (12') and by groups of second blades (14) , said groups of first (12) and second blades (14') alternating circumferentially. 4. Assembly (4) according to claim 3, characterized in that each group of first blades (12') comprises between 3 and 8 blades (12), preferably 4 blades (12). 5. Assembly (4) according to one of claims 3 and 4, characterized in that each group of second blades (14') comprises between 4 and 10 blades (14), preferably 5 blades (14). 6. Assembly (40) according to one of claims 1 or 2, characterized in that the row of blades (100) further comprises a plurality of third blades (30) arranged circumferentially regularly between the first (12) and second blades (14), said third blades (30) having a respective orientation which is included, limits excluded, between the orientation perpendicular to the axis (A) of the first blades (12) and the orientation perpendicular to the guide surface (8) second blades (14). 7. Assembly (40) according to claim 6, characterized in that the respective orientation of the third vanes (30) varies progressively in accordance with the circumferential position of the third vanes (30) relative to the first (12) and second vanes (14). ), the third blades (30) having an orientation all the closer to that of the first blades (12) as they are circumferentially close to it, and the third blades (30) having an orientation all the closer to that of the second blades (14) that they are close circumferentially. 8. Assembly (40) according to claim 7, characterized in that the third blades (30) have a respective orientation angle which varies step by step by at least 0.5°. 9. Assembly (40) according to claim 8, characterized in that, seen in a radial section, the progression of the orientation of the third blades (30) presents a triangular profile (22). 10. Assembly (40) according to claim 7, characterized in that, seen in a radial section, the progression of the orientation of the third blades (30) presents a sinusoidal profile (20). 11. Assembly (4; 40) according to one of claims 2 to 10, characterized in that the trailing edge (13) of the first blades (12) has a first distance (D) following the direction of the air flow (F1) with the leading edge (17) of the structural arms (16), and the trailing edge (15) of the second blades (14) has a second distance (d) following the direction of the air flow (F1) ) with the leading edge (17) of the structural arms (16), said first distance (D) being greater than the second distance (d). 12. Assembly (4; 40) according to one of claims 1 to 11, characterized in that all the blades (12, 14, 30) have the same chord length. 13. Assembly (4; 40) according to one of claims 1 to 12, characterized in that the guide surface (8) of the air flow (F1) forms an angle of inclination with the axis (A). (a) between 10° and 60°. 14. Axial turbomachine (1) comprising a low pressure compressor (3), a high pressure compressor (5), and a swan neck vein (2) arranged axially between the low pressure compressor (3) and the high pressure compressor ( 5), the turbomachine being characterized in that the vein (2) contains an assembly (4; 40) according to one of claims 1 to 13.