ASSEMBLY FOR A TURBINE ENGINE

Description

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 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 can be commonly called OGV “Outlet Guide Vane”) is directly followed downstream by structural arms (or “struts”) which cross the airflow vein of the axial turbomachine.

The structural arms are generally arranged in a sloping vein (or in a swan neck between the compressors), i.e. the direction of the airflow 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 arms.

OGV blades usually have a direction substantially perpendicular to the flow direction to limit aerodynamic disturbances that would make the flow unstable at the OGV. With such an orientation of the OGV blades, the few blades that are circumferentially directly in front of the structural arms have a radially outer portion that 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 an efficiency loss for the turbomachine.

In order to overcome the problem of static pressure distortion, there is a solution consisting of axially moving the OGV 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 in improving the performance of the compressor without affecting its dimensions.

SUMMARY OF THE INVENTION

Technical Problem

The invention aims to solve the drawbacks of the design of turbomachine compressors of the state of the art. In particular, the invention aims to propose a solution that makes it possible to limit aerodynamic disturbances at 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 complicating and/or adding additional weight to the compressor of the turbomachine.

Technical Solution

The invention relates to an assembly for a turbomachine comprising a wall having a surface for guiding an airflow and a row of stator blades extending from the wall and arranged annularly around an axis, said surface being inclined relative to the axis, wherein the row of blades comprises first blades extending substantially perpendicularly to the axis and second blades extending substantially perpendicularly to the guide surface.

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

By “annularly extending blade row” is meant a single ring of blades arranged circumferentially one after the other. These blades delimit inter-blade spaces traversed by the airflow.

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 promoting the guiding of the airflow instead of the trailing edge of the row of blades.

In addition, the manufacture of such a row is simplified, as is 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, excluding limits, 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 airflow at the trailing edge of the row of blades, thus promoting the regularity and progressiveness of said airflow 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 all the closer to that of the first blades as they are circumferentially close to them, and the third blades having an orientation all the closer to that of the second blades as they are circumferentially close to them.

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

According to an advantageous embodiment of the invention, seen in a radial cross-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 cross-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 along the direction of the airflow with the leading edge of the structural arms, and the trailing edge of the second blades has a second distance along the direction of the airflow 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 makes it possible to promote control of any distortion of the static pressure, particularly by limiting it at the level of the entire span of the leading edges of the structural arms.

According to an advantageous embodiment of the invention, the airflow guide surface forms with the axis the angle of inclination 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, wherein 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 airflow vein, this is notably 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.

Furthermore, the static pressure distortion phenomenon is reduced by the fact that the radial stator blades are moved away from the leading edge of the structural arms, and this, without modifying the axial position of the row of stator blades in the airflow 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, pressure losses are limited and compressor operability is increased, which advantageously improves compressor behavior.

DESCRIPTION OF THE DRAWINGS

FIG. 1 shows a simplified partial view of an axial turbomachine;

FIG. 2 is a schematic illustration of a longitudinal sectional view of an axial turbomachine vein comprising an assembly according to the invention;

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

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

DESCRIPTION OF AN EMBODIMENT

In the following description, the terms “internal”, “inner”, “outer” and “outer” refer to a 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 flow direction of an axial airflow in a main vein of the turbomachine.

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

FIG. 1 shows a simplified partial view of an axial turbomachine 1. In this specific case, it is a dual-flow turbojet, but can also be a turbojet, turbofan, turboprop, turbomotor 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 illustrated).

In operation, the mechanical power of the turbine(s) transmitted via the central shaft to the rotor sets the rotors of the two compressors 3 and 5 in motion. The rotation of the rotor around its longitudinal axis of rotation A thus makes it possible to generate an airflow and to gradually compress the latter up to the inlet of the combustion chamber.

An inlet fan commonly referred to as a fan or blower is coupled to the rotor and generates an incoming airflow F. The blower is arranged upstream of a separation nozzle 11 capable of separating the incoming airflow F into a radially internal airflow, called airflow F1 or primary flow F1, circulating in a primary flow vein 2 and passing through the various aforementioned levels of the turbomachine 1, and a radially external airflow, called secondary flow F2 passing through an annular duct 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 fully detailed in the present description.

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

Referring to FIG. 2, the vein 2 allows the primary flow F1 to be guided between an outer wall 6 having an outer air guide surface 8 and an inner wall 9 having an inner guide surface 7.

The assembly 4, 40 comprises a row of stator blades 10, 100, also called stator grid or OGV, said row of stator blades 10, 100 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 may for example be an internal shroud 9.

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

The wall 6, and in particular the guide surface 8, is inclined relative to the axis A. In this respect, a straight line tangent to the guide surface 8 at the right of the row of stator blades 10, 100 has with the axis A an angle of inclination a of between 5° and 80°, and preferably of between 10° and 60°. The inclination along the angle α results in an inclined guiding of the air along a downward slope, resulting in an overall inclination of the direction of the primary flow F1 relative to the axis A at the right of the assembly 4, 40.

The row of stator blades 10, 100 may 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 marked 12 and subsequently called “first blades”, which extend substantially perpendicular to the 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 14 blades 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 α+90°.

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

Along the direction of the airflow F1, and preferably along a projection onto 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, “distance” is given its usual definition, i.e. 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, at the level of the internal shroud 9, and more precisely at the level of the point P illustrated in FIG. 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 airflow F1.

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

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

Referring to FIG. 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 inter-arm spaces 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 alternating circumferentially (i.e. according to the horizontal orientation of FIG. 3) around the axis of the turbomachine.

Preferably, and in order to maximize stability instead of the blades of the stator blade row 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 10 blades, and even more preferably 5 blades.

In particular, it can be seen in FIG. 3 that the second blades 14 can be axially closer (in a projection on a vertical axis of FIG. 3) to the arms 16 than are the first blades 12.

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

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

Referring to FIG. 4, the row of stator blades 40 comprises, in addition to the first 12 and second 14 blades, a plurality of third blades 30 arranged circumferentially between the first 12 and second 14 blades.

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

The third blades 30 have a respective orientation which is included, excluding limits, 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 progressively 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 blades 100 is mainly due to the progressiveness of the orientation of the third blades 30 between the extreme terminals, i.e. the first 12 and second blades 14.

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

Preferably, the third blades 30 have a respective orientation angle which varies from close to close by at least 0.5° and/or the variation of 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 profile 22 and/or sinusoidal profile 20, such an orientation progression profile makes it possible to further smooth the airflow and to avoid aerodynamic disturbances (for example due to corner effects or sudden variations in the flow guide surfaces).

In summary, the difference between the first and second embodiments 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 embodiments of the invention, and particularly those being circumferentially adjacent to the structural arms, can have a setting or an orientation, also called “tilting”, and which makes it possible to have a different blade profile at 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 spacing more circumferentially 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 may result from the combination of a portion of row 10 of FIG. 3, with a portion of row 100 of FIG. 4.