Sub-component of low-pressure compressor for aircraft turbine engines

Abstract

The present invention relates to a sub-assembly for a low-pressure compressor for an aircraft turbine engine, the sub-assembly comprising a straightener (121) and a rotor hub (6), the straightener being provided with cantilevered blades (7), the rotor hub comprising a cavity (2) covered by an inner shroud (3) opposite to the blades (7), an orifice (5) being formed in the inner shroud (3) to allow airflow to pass from the downstream to the upstream direction of the low-pressure compressor (120).

Description

[0001] Case Analysis This application is a divisional application of Chinese patent application No. 202180045164.4, entitled "Sub-assembly of a low-pressure compressor for an aircraft turbine engine", which entered the Chinese national phase of PCT international patent application PCT / EP2021 / 062437, filed on May 11, 2021. Technical Field

[0002] The present invention relates to a sub-assembly of a low-pressure compressor for an aircraft turbine engine, and such a low-pressure compressor including the sub-assembly. Background Technology

[0003] Typically, aircraft turbine engines are equipped with two compressors (a low-pressure compressor and a high-pressure compressor) to draw in air and compress it to the appropriate speed, pressure, and temperature before it is delivered to the combustion chamber.

[0004] Each of these compressors typically comprises multiple compressor stages aligned along the engine axis and oriented from upstream to downstream. Each stage consists of a movable component (rotor section) extending along the engine axis and a fixed component (stator section) referred to as a "straightener." Thus, the movable and fixed components of a stage alternate along the engine axis. Each component consists of blades, specifically a ring of impeller blades arranged circumferentially around the engine axis. Technical parameters such as the size and geometry of the impeller blades are determined such that the operating conditions of each stage are adapted to the operating conditions of the upstream and / or downstream stages along the engine axis.

[0005] A known low-pressure compressor straightener is equipped with overhanging blades, each of which is attached to the housing via a root and extends approximately radially inward from the root of the blade toward the engine axis. In this way, the head of the overhanging blade, radially opposite to its root, is free and opposite the hub of the rotor, to which the movable components of the stage are connected. This architecture of the straightener obviously requires a gap between the head of each blade (and thus fixed) and the hub of the rotor (and thus capable of self-rotation). When the turbine engine is operating, this gap typically results in the generation and flow of air vortices from downstream to upstream (i.e., in an orientation opposite to the main airflow flowing through the low-pressure compressor), a phenomenon known as “leakage.” These vortices are caused by pressure differences between the inner and outer arcs of the straightener and between the trailing and leading edges of the blades, as the pressure of the main airflow within the compressor increases from upstream to downstream.

[0006] These eddies cause losses in the compressor, thus affecting its efficiency. Therefore, it is preferable to limit compressor losses.

[0007] A known solution to this problem is described in document EP 3 095 963 A1. This solution primarily involves adding an inner shroud to the straightener between the rotor hub and each blade head of the straightener to reconstruct the compressor's internal aerodynamic ductwork at the level of the straightener. In this case, the blade heads are no longer free. The blade heads are attached to the inner shroud at a sealing interface to prevent the formation of leakage eddies. However, since the shroud itself must be positioned between the rotor and stator sections, a clearance between these sections is still required, which in turn leads to air leakage under the inner shroud, resulting in compressor losses. To limit this leakage, as disclosed in EP 3 095 963 A1, the inner shroud is equipped with a wear-resistant coated track, and the rotor hub is equipped with a wiper opposite to this track. However, the integration of these additional components implies an undesirable increase in mass, thereby reducing the efficiency of the low-pressure compressor. In the case of a variable pitch straightener, this increase in mass is further aggravated because each blade of the straightener must also be equipped with a pivot inserted into an inner shroud, which in this case typically consists of two parts that must be joined together. Summary of the Invention

[0008] One object of the present invention is to provide a low-pressure compressor subassembly for an aircraft turbine engine, the low-pressure compressor subassembly including a straightener equipped with overhanging blades, thereby making the low-pressure compressor more efficient.

[0009] To this end, the present invention proposes a sub-assembly for a low-pressure compressor of an aircraft turbine engine, the sub-assembly being oriented along the engine axis to extend from upstream to downstream. Sub-components include: - Rotor, the rotor including a hub extending axially and circumferentially around the engine axis; - Stator, the stator includes a straightener equipped with (suspended) blades that extend generally radially toward the hub; Its features are: - The hub includes a recess oriented toward the engine axis; -The rotor includes an inner protective cover: • The inner cover covers the recessed portion to define a cavity between the hub and the inner cover. Each of the blades includes a free head that at least partially faces the inner shroud and / or cavity; • The inner shroud includes an opening that communicates with the cavity fluid.

[0010] The low-pressure compressor subassembly according to the invention enables the reduction of the negative impacts of leakage eddies that may occur between the impeller head and the rotor without negatively affecting the quality of the low-pressure compressor. In particular, the low-pressure compressor subassembly enables higher efficiency of the low-pressure compressor without the disadvantages of solutions known in the prior art.

[0011] In practice, instead of adding additional elements to the stator to limit the aforementioned leakage vortices, the low-pressure compressor subassembly according to the invention proposes treating the rotor hub by forming a recess defined by an additional inner shroud pierced by orifices, preferably in a non-axisymmetric manner, directly opposite the blades of the straightener, without modifying the stator portion in any way. Specifically, unlike the solution listed in document EP 3 095 963 A1, each blade retains a free head (and thus remains overhanging). Only the rotor portion is modified. The cavity and orifice passively allow airflow to circulate at the level of the straightener, independent of the main airflow in the low-pressure compressor. In this way, any leakage vortices that may form can circulate from downstream to upstream through the cavity via the orifices, as the cavity provides a larger space than the minimum clearance typically left between the blade heads and the hub. Advantageously, the location of the orifices can be selected to bring the airflow of these leakage vortices from downstream to upstream to the desired location at the level of the straightener, or even upstream of the straightener. In this way, the airflow can be influenced to confine leakage eddies to the level of the gap between the stator and rotor, while limiting the negative impact of these leakage eddies on the efficiency of the low-pressure compressor. This advantage is also achieved without altering the blade architecture, which is noteworthy because this architecture offers advantages in terms of mass, aerodynamic efficiency, and design simplicity. Finally, it should be emphasized that the present invention does not reduce the efficiency of the low-pressure compressor by increasing its mass. Therefore, advantageously, the low-pressure compressor sub-assembly (hereinafter referred to as the "sub-assembly") according to the invention results in a higher efficiency of the low-pressure compressor without negatively impacting its mass.

[0012] Since the present invention is based on the fact that the cavity allows for the initial airflow between the rotor and stator, it is highly preferred that the gap between the inner shroud and the blade head be minimized to avoid excessive parallel flow of these leakage vortices between the blade head and the inner shroud. Furthermore, the presence of the cavity makes it preferable to reconstruct the rotor's internal aerodynamic ducts as cleanly and uniformly as possible to avoid aerodynamically affecting the compressor. Therefore, for both reasons, it is preferable that the inner shroud is an extension of the outer surface of the hub, causing the rotor's internal aerodynamic ducts to be (re)formed in a regular manner. Preferably, the initial gap between the hub and the blade head corresponds to the gap between the inner shroud (and / or the hub) and the blade head. This preferred embodiment satisfies the aerodynamic requirements within a low-pressure compressor.

[0013] For completeness, some of the concepts described above, well-known to those skilled in the art, are reviewed herein. A “hub” is the central portion of a conical and / or annular and / or disc-shaped rotating mechanical component. In the case of a low-pressure compressor, the hub of the rotor is a central conical component because the hub of the rotor extends along and around the engine axis. The use of a hub attached to movable blades of the compressor is a typical feature of low-pressure compressors; in high-pressure compressors, a “disc” is generally used instead of a hub. Preferably, the movable blades of the rotor belonging to the same low-pressure compressor stage as the straightener discussed comprise movable blades having roots attached to the hub. The term “overhanging blade” is also well-known to those skilled in the art. The term “overhanging” is defined as remaining above a clearance without direct support from below. In the case of the straightener blades, the straightener blades are radially extending from the outside of the blade towards the inside, wherein the head of each blade is held (“in vacuum”) directly opposite the hub of the rotor (and / or the inner shroud and / or the cavity, in the case of the invention), located horizontally inside the low-pressure compressor. In this case, preferably, the blade root is attached to the outer casing of the low-pressure compressor. Thus, each blade includes an outwardly attached “root,” an inwardly (i.e., unattached) “free head,” an upstream “leading edge,” and a downstream “trailing edge” for engaging the main airflow flowing through the low-pressure compressor. Preferably, the blades in question are circumferentially aligned and have similar profiles. Finally, it should be noted that the reference to “a straightener equipped with blades” does not necessarily limit the blades considered in this statement (and the references below to “each blade in the blades” or “multiple blades”) to all blades of the straightener. Therefore, the selection of straightener blades is within the scope of these statements. However, preferably, all blades of the straightener are considered by the straightener.

[0014] For the purposes of this article, it should also be recalled that the "recess" (in this case, in the hub) takes the form of a hollow portion and / or a material removal portion. The recess is preferably formed by processing the hub and preferably extends in both the axial and circumferential directions. The recess is preferably not axisymmetric with respect to the engine axis. A "cavity" is an empty space in a solid (in this case, in the rotor portion under consideration). This cavity is preferably defined by the inner shroud and the hub, or more precisely, by the outer surfaces of the inner shroud and the hub that define the recess. Finally, it should be recalled that the term "orifice" refers to an opening that fluidly communicates the cavity with the outside (in this case, preferably with the outer space separating the impeller head and the inner shroud). In particular, an orifice is a hole that opens on both sides of the inner shroud. These definitions are well known to those skilled in the art and are specified only for completeness. In the case of this invention, each orifice passes through the inner shroud. The fact that the orifice is precisely fluidly connected to the cavity formed by the recess and defined by the inner shroud allows airflow (especially leakage eddies) to pass through the orifice into the cavity, thereby alleviating a significant portion of the airflow in the outer space (corresponding to the gap between the stator and rotor). The term "including orifice" as used herein is understood to mean "including at least two orifices" to enable the capture and injection of this airflow. The invention is by no means limited to the presence of a single cavity and / or a single recess. In particular, the rotor may independently or non-independently include multiple such recesses, cavities, and orifices.

[0015] According to one embodiment of the invention, the recess is formed by at least one deformed portion of the wall of the hub. The wall is preferably formed as a portion of a plate or sheet of metal. The thickness of the wall is preferably substantially constant from upstream to downstream of the recess.

[0016] In one embodiment of the invention, the inner cover is partially formed of an annular plate (or metal sheet) and preferably has a constant thickness.

[0017] According to a preferred embodiment of the invention, a first orifice (at least) extends downstream of the leading edge of one of the blades (opposite to the straightener), and a second orifice (at least) extends upstream of the same leading edge.

[0018] Therefore, the first orifice can preferably be located between the cavity and the head of the blade (at certain moments during the rotation of the hub around the engine axis, the first orifice is opposite the head of the blade). Advantageously, this arrangement of at least two orifices makes it possible to passively control at least a portion of the airflow, which forms a possible leakage vortex by being captured downstream at the level of the first orifice and re-injected upstream through the second orifice, assuming that the two orifices are in fluid communication with the cavity, and that this portion of the airflow flows within the cavity and between the two orifices.

[0019] It should be noted that the terms "downstream" and "upstream" refer to positions along the engine axis as the sole frame of reference. Specifically, mathematically, if the component of a first point in space along the engine axis is less than (and correspondingly greater than) the component of a second point, then the first point in space is upstream (and correspondingly downstream) of the second point in space. Applying the same limitation, if a first point assembly includes at least one point located upstream (and correspondingly downstream) of all points in a second point assembly, then the first point assembly (in this case, abstractly corresponding to an orifice or blade edge) extends "(at least) upstream of the second point assembly" (and correspondingly, "at least downstream of the second point assembly"). This limitation will be distinguished from the case where the first point assembly extends "only (or entirely, completely) upstream of the second point assembly" (and correspondingly, "only downstream of the second point assembly"), which corresponds to the fact that every point in the first point assembly is upstream (and correspondingly downstream) of all points in the second point assembly. Since the terms "upstream" and "downstream" are widely used in this art, these formal expressions will be apparent to those skilled in the art.

[0020] This relationship between the positions of the first and second orifices and the leading edge of the impeller is not limited to the circumferential position of the orifices relative to the leading edge. In particular, when the low-pressure compressor is operating with the hub rotating, this circumferential position may change over time.

[0021] Preferably, according to the foregoing embodiment, the section of the first orifice extends axially between corresponding radial protrusions on the hub and / or inner shroud of the leading and trailing edges of the blade. Specifically, in this case, as the hub rotates about the engine axis, in each rotation of the hub, the head of the blade in question must be opposite the first orifice, and therefore opposite the cavity. Since the clearance between the blade head and the rotor is typically minimal, this allows for more efficient direct capture of leaking vortex airflow into the cavity through the first orifice. In this way, the airflow is controlled and delivered through the cavity to a second orifice upstream of the leading edge, where it is reinjected into the main airflow of the low-pressure compressor. Preferably, the section of the first orifice and the second orifice are at least partially axially aligned, i.e., there exists a straight line, preferably multiple straight lines, parallel to the engine axis intersecting the sections of the first and second orifices.

[0022] The inner shield preferably extends on the outer surface of the hub to (re)form an inner aerodynamic duct. Each of the aforementioned radial protrusions is preferably considered on the outer surface of this duct. In particular, the blade edge is radially projected through an orthogonal projection of each point on this surface, which is obtained by the intersection of the surface with a line perpendicular to the engine axis passing through that point.

[0023] Preferably, in the above-described preferred embodiment, the inner cover includes multiple pairs of first openings and second openings as described above. This description does not exclude the possibility that a pair of second openings is another pair of first openings. Examples of this configuration are given below. Figure 1 As shown. Preferably, the orifices (and / or the first and second orifices) are uniformly distributed on the inner cover.

[0024] Preferably, the orifice extends only upstream of the trailing edge of each blade. It is practically advantageous to introduce an orifice that is generally axially opposite to the blade head and upstream of the blade head to capture the airflow and reinject the airflow upstream and / or at the inlet of the straightener into the main airflow of the low-pressure compressor.

[0025] In a preferred embodiment of the invention, the assembly of orifices is generally circumferentially aligned. Preferably, each orifice has a parallelogram profile extending axially at an angle of inclination of up to 60° relative to the engine axis. The advantage of such orifices is that they can be easily pierced one after another (potentially regularly) within the inner casing. As these orifices extend axially, they allow airflow from leaking vortices to be carried upstream across the entire axial range of the orifice, not just at a specific location. The inclination angle further allows the orifices to be more or less circumferentially oriented according to the rotor's rotational motion, preferably such that the blade tip radially faces multiple such orifices communicating with the cavity, thereby enabling better control of the downstream-to-upstream airflow. The shape of this assembly of orifices contributes to improving the overall aerodynamic performance of the aircraft turbine engine and reducing its fuel consumption.

[0026] It should be noted that the fact that all the orifices in the set have parallelogram profiles does not strictly limit the shape of the parallelogram for each orifice. In particular, the set may optionally include orifices that extend further in the axial and / or circumferential directions than other orifices, and / or orifices that have potentially different angles of inclination. However, it is preferred that all orifices in the set have a profile corresponding to a single parallelogram.

[0027] Preferably, each orifice in the assembly extends upstream and downstream of the leading edge of each blade. In this way, each orifice both traps airflow downstream of the leading edge of the blade and reinjects that airflow upstream of those leading edges. More preferably, each orifice in the assembly comprises: - Upstream end, the upstream end is located upstream of the leading edge, at an axial distance from the leading edge, the axial distance being at most 25% of the chord of each blade, preferably between 10% and 25%; - Downstream end, located downstream of the leading edge, at an axial distance from the leading edge, the axial distance being at most 75% of the chord of each blade, preferably between 10% and 75%.

[0028] It should be recalled that the term "axial distance" refers to a distance measured along the engine axis. Specifically, the axial distance between two points in space is the absolute value of the difference between the components of the two points along the engine axis. This (axial) distance between two point assemblies is generally considered to be the minimum (axial) distance between a point in one of these assemblies and a point in another of these assemblies. The "chord" of a blade is a measurement of the extension of that blade along the engine axis, preferably taken at the level of the blade tip. The term "chord of a blade" is well known to those skilled in the art and generally refers to the "axial length" of the blade, which is generally the length measured along the engine axis.

[0029] Values ​​of up to 25% and 75%, along with preferred values ​​associated with those values, indicate that the orifice is capable of extending axially, sufficiently upstream to rigorously reinject airflow upstream of the straightener, and sufficiently downstream (while remaining upstream of the trailing edge of the blade) to effectively capture the airflow along the gap between the blade head and the inner shroud and / or hub.

[0030] In accordance with this spirit, and preferably according to these later embodiments, the axial length of each orifice of the assembly is between 10% and 75% of the chord of each blade.

[0031] According to a preferred embodiment of the invention, in conjunction with the foregoing preferred embodiments, the orifice comprises two circumferentially aligned groups of orifices, the axial distance between these groups of orifices being between 10% and 50% of the chord of each blade. Therefore, airflow can be captured at a selected downstream location of the leading edge of the blade using the orifice of one of these groups of orifices, and the airflow can be reinjected at an upstream location (preferably upstream of these leading edges) through the orifice of the other group of orifices. Thus, airflow is transported between the orifices of these groups of orifices through a cavity without interfering with the airflow at the axial distance separating the two groups of orifices. The points of airflow capture and reinjection can also be selected more precisely, away from the aforementioned axial distance.

[0032] According to a specific embodiment, one orifice group corresponds to the set described above according to the foregoing embodiments, and the orifices of another orifice group in the orifice group each have a different parallelogram profile. Preferably, the orifices of the other orifice group extend axially at another angle of inclination of up to 60° with respect to the engine axis.

[0033] According to a preferred embodiment of the invention, the orifices occupy more than half of the (cylindrical) surface of the inner shroud, which extends axially between the entire upstream and downstream ends of all the orifices. In other words, between the aforementioned entire ends, the inner shroud comprises an orifice surface area larger than the material itself.

[0034] According to a specific embodiment of the invention, the cavity is divided into a plurality of units fixed relative to the impeller and defined by edges, each unit extending in both the axial and circumferential directions along an axially inclined direction, such that each unit comprises: -Downstream section, the downstream section radially faces the space separating the first and second blades in the circumferentially aligned blades; -Upstream section, the upstream section radially faces another space that separates the third and fourth blades in the circumferentially aligned blades, at least one of the third and fourth blades being different from the first and second blades.

[0035] This statement applies at any time, even during hub rotation, so at each such time, there are always downstream and upstream sections opposite different spaces between two successively aligned blades. In this way, the airflow captured downstream at the axial level of the blades and / or the space between the two blades is reinjected upstream at another blade and / or another space between the two blades. This allows for improved compressor performance by limiting disturbances that might be caused by excessively large cavities. In a particular embodiment of the unit, the fourth blade corresponds to the first blade, such that the third, first, and second blades are aligned circumferentially in this order. This particular embodiment of the unit advantageously allows the airflow of leaking vortices to be captured at the level of the inner arc of the blade (in this case, the first blade), and this airflow is extracted from the outer arc of the same blade. This facilitates the reinjection of this airflow into the main airflow. The edges of the unit may be beveled to optimize airflow within each unit. Preferably, each orifice is in fluid communication with a single unit.

[0036] In a general embodiment of the invention, the edge of the orifice is beveled to have an open profile. This profile preferably opens outward downstream of the leading edge of the impeller to facilitate airflow capture and inward upstream of the leading edge of the impeller to facilitate airflow reinjection upstream of the straightener.

[0037] Typically and preferably, the sub-assemblies according to the invention include an outer (and fixed) housing. Each blade of the straightener includes a root attached to the housing and extends generally radially inward from the root of the blade toward the free head of the blade.

[0038] The present invention also proposes a low-pressure compressor stage comprising a sub-assembly according to the invention. In this case, the rotor is preferably equipped with movable blades, each of which includes a root attached to the hub and extends generally radially outward from the root. The present invention also proposes a low-pressure compressor for an aircraft turbine engine, the low-pressure compressor comprising a stage and / or sub-assembly according to the invention. Preferred embodiments and advantages of the sub-assemblies according to the invention, with necessary modifications, are applicable to the stage and the low-pressure compressor of this low-pressure compressor.

[0039] Finally, the present invention provides an aircraft turbine engine equipped with a low-pressure compressor according to the invention. The preferred embodiments and advantages of the low-pressure compressor according to the invention, with necessary modifications, are applicable to this aircraft turbine engine.

[0040] The use of the verb "including" and its variants, as well as its inflections, in this document does not preclude the existence of other elements besides those mentioned above. The use of the indefinite article "a" or the definite article "the" to introduce an element does not preclude the existence of multiple such elements. The terms "first," "second," "third," etc., used within the scope of this document are merely for distinguishing similar elements and do not imply any order among these elements.

[0041] It should be recalled that the present invention relates to the technical field of compressors (and particularly, low-pressure compressors) for aircraft turbine engines. This is a very specific technical field with particular technical limitations on compressors. In particular, this technical field should not be confused with and / or mixed with the separate field of turbines for aircraft turbine engines. In particular, it should be recalled that the purpose of a compressor is to compress the air entering the aircraft turbine engine at the inlet, while the purpose of a turbine is to expand the gas at the outlet of the combustion chamber of the aircraft turbine engine. In particular, the functions, locations, and technical limitations (e.g., speed, temperature, exposure to external debris, etc.) associated with the operation of compressors and turbines for aircraft turbine engines are particularly distinct. Those skilled in the art of compressors for aircraft turbine engines (not to mention the very specific technical background of the invention described in the prior art) will not consult and will not be inspired by the prior art related to turbines for aircraft turbine engines to develop the present invention related to compressors, unless it becomes clear from that prior art how the many technical differences between these technical fields can be considered. Attached Figure Description

[0042] Other features and advantages of the invention will become apparent from the following detailed description, and with reference to the accompanying drawings for understanding the description, in which: - Figure 1A partial schematic diagram of a cross-section of a low-pressure compressor subassembly according to an embodiment of the present invention is shown; - Figure 2 and Figure 3 A schematic diagram of a partial orifice and a straightener blade arrangement according to an embodiment of the present invention is shown, the straightener blade arrangement having a profile projected onto the internal aerodynamic duct of the low-pressure compressor; - Figure 4A and Figure 4B Schematic diagrams and partial three-dimensional and two-dimensional projection views of cavities divided into units in the rotor hub according to embodiments of the present invention are shown respectively. - Figure 5 A simplified schematic cross-section of an aircraft turbine engine according to an embodiment of the present invention is shown.

[0043] The figures in the accompanying drawings are not drawn to scale. Generally, similar reference numerals are used in the figures to refer to similar elements. Within the scope of this document, identical or similar elements may have the same reference numerals. Furthermore, the presence of reference numerals or letters in the figures should not be considered limiting, even if such numbers or letters are indicated in the claims. Detailed Implementation

[0044] This section provides a detailed description of preferred embodiments of the invention. The invention is described with reference to specific embodiments and the accompanying drawings, but is not limited to these embodiments and drawings. The figures and / or drawings described below are illustrative only and not restrictive.

[0045] In some of these figures, references are shown as abstract geometric reference frames, primarily for quantifying and / or visualizing the characteristics of embodiments of the invention. In this context, the terms "axial," "circumferential," and "radial" directions are used, corresponding respectively to directions parallel to the engine axis, generally circular around the engine axis, and perpendicular to the engine axis. The reference frames in the figures show these directions (with orientations) labeled X, R, and Y, respectively. By misuse of the notation for similar elements, the engine axis will also be referred to as X, since the corresponding vector X has the same orientation and direction as the engine axis. The terms "axially," "radially," and "circumferentially" are derived from the terms "axial," "radial," and "circumferential," respectively, and have similar preferred meanings. Furthermore, the terms "circumferential" and "radial" preferably refer to polar coordinate systems known to those skilled in the art in each plane perpendicular to the engine axis. The terms "inward" and "inner" naturally correspond to orientations toward the engine axis X in the radial direction, and the terms "outward" and "outer" correspond to opposite orientations in that direction.

[0046] Figure 5 An axially dual-flow aircraft turbine engine 100 is shown, which continuously includes a fan 110, a low-pressure compressor 120, a high-pressure compressor 130, a combustion chamber 160, a high-pressure turbine 140, and a low-pressure turbine 150 along the engine axis X. These components are known to those skilled in the art. In operation, the mechanical power of the low-pressure turbine 150 and the high-pressure turbine 140 is transmitted to the low-pressure compressor 120 and the high-pressure compressor 130 via a low-pressure shaft 101 and a high-pressure shaft 102, respectively, and to the fan 110 via the low-pressure shaft 101. As is known, the fan 110 enables the generation of a primary airflow 106 passing through the aircraft turbine engine 100 in a primary airflow duct and a secondary airflow 107 flowing outward around the compressors 120, 130 and the turbines 140, 150.

[0047] The low-pressure compressor 120 includes movable blades 122 and a straightener 121 consisting of fixed blades, the movable blades and the straightener alternating around the engine axis X along the engine axis X. Figure 1 A cross-section of a subassembly 1 of a low-pressure compressor 120 according to a preferred embodiment of the invention is shown. The subassembly 1 includes one of the straighteners 121 described above. The straightener includes a plurality of circumferentially aligned fixed overhang blades 7 having similar profiles. Each blade 7 includes a root 72 attached to the housing 9 of the low-pressure compressor 120, and a free head 71 (i.e., not attached to any other element) opposite the root 72 along the radial extension direction of the blade 7. Specifically, the blades 7 extend generally radially inward. The blades 7 also include a leading edge 7A primarily oriented upstream and a trailing edge 7B downstream. The straightener 121 causes the flow of the main airflow 106 to return parallel to the engine axis X, while increasing the pressure of the main airflow and reducing its absolute velocity.

[0048] Subassembly 1 also includes a rotor portion (or rotor) comprising a hub 6 that extends circumferentially about and rotates about the engine axis X. The rotor portion typically includes movable blades extending generally radially outward from a root attached to the hub 6 toward the housing 9. The fixed blades forming the straightener 121 and the movable blades of the rotor are then assembled adjacent to each other along the engine axis X.

[0049] Typically, the free head 71 of the suspended impeller 7 faces (or similarly faces) the hub 6 radially, but since the impeller 7 is fixed and the hub 6 is rotating, they do not contact each other. Therefore, only a small gap (or space) 10 exists between the free head 71 of the impeller 7 and the hub 6. This gap leads to the generation and flow of air from the leakage vortex from downstream to upstream, resulting in a loss of efficiency in the low-pressure compressor 120. To mitigate this effect, the present invention proposes to process the hub 6 to form at least one recess (i.e., an inwardly recessed portion) 8, which is at least radially opposite the passage position of the free head 71 of the impeller 7.

[0050] It is also proposed to add an inner shroud 3 to the rotor portion to cover the recess 8, thereby extending the outer surface of the hub 6 without the recess and restoring the inner aerodynamic duct 4 that was originally present before the hub 6 was treated. In this way, the gap 10 between the inner aerodynamic duct 4 and the free head 71 of the blade 7 is substantially maintained. The inner shroud 3 is preferably attached to the hub 6 by welding. This operation also makes it possible to define a cavity 2 defined axially and radially inward by the recess 8 of the hub 6 and radially outward by the inner shroud 3. The cavity 2 preferably has a radially truncated depth in the range of 5% to 20% of the radial length of the blade 7. The cavity 2 preferably has an axially truncated length in the range of 50% to 150% of the axial width of the blade 7.

[0051] The inner shroud 3 includes an orifice 5 communicating with the cavity 2. At least one orifice communicates with the gap 10. Each free head 71 faces the inner aerodynamic duct 4, more precisely, at least partially facing the inner shroud 3 and / or the cavity 2, and optionally partially facing the hub 6. The orifices 5 are arranged such that when the hub 6 rotates in a regular and / or periodic manner, the free head 71 of the impeller 7 faces downstream of the first orifice 51 (or orifice segment) and the second orifice 52 (or other orifice segment). Therefore, the airflow from the leakage vortex (in Figure 1 (Indicated by arrows) can be captured through these first orifices 51, transported from downstream to upstream within the cavity 2, and re-injected upstream through these second orifices 52.

[0052] Figure 2 and Figure 3 Each is shown with an internal aerodynamic duct 4 (shown as unfolded circumferentially and expanded in a plane), the profile of the blade 7 projected radially onto the internal aerodynamic duct and including an orifice 5. A reference portion introduced for the blade is applied in a similar manner. The leading edge 7A and trailing edge 7B of the blade 7 are located on lines 70A and 70B, respectively.

[0053] The orifice 5 comprises a set of orifices 5 (or a first set of orifices 5A), which extend axially and circumferentially along the axial length C and all have a parallelogram profile. The preferred inclination angle α of the orifice relative to the engine axis X is between 0° and 60°, preferably between 30° and 45°. Along the circumference of the inner aerodynamic duct 4, the orifices have a constant circumferential width F and are separated by a circumferential space E, preferably smaller than the circumferential width F. Each of these orifices 5 includes an upstream end 11 at an axial distance A from line 70A and a downstream end 12 at an axial distance A' = CA from line 70A, where the axial distance A is between 5% and 15% of the chord B of the blade 7, and the axial distance A' is preferably between 10% and 75% of the chord B of the blade 7. Specifically, all the orifices 5 are located only upstream of the trailing edge 7B of the blade 7, because it is precisely at the axial level of the free head 71 of the blade that the airflow from the leakage vortex must be captured and carried upstream of the leading edge 7A of the blade 7. These parallelogram-shaped orifices 5 have the advantage of being very easy to design and to effectively achieve the technical effects desired by the present invention.

[0054] exist Figure 2 In this case, the orifice 5 extends axially over a large axial length C, preferably between 60% and 80% of the chord B of the blade 7. Thus, these orifices 5 extend continuously from the downstream end 12 located between the leading edge 7A and the trailing edge 7B of the blade 7 to the upstream end 11 located upstream of the leading edge 7A of the blade 7, thereby capturing airflow over the entire length of the orifice to carry the airflow from downstream to upstream through the cavity 2 and reinjecting the airflow into the main airflow upstream of the straightener 121. Simultaneously, the tilt angle α means that the downstream section of the first orifice 51 extends relative to the blade between the leading edge 7A and the trailing edge 7B of the blade 7, while the upstream section of the second orifice 52 extends upstream (and downstream) of the leading edge 7A of the same blade 7. These two sections are at least partially axially aligned, i.e., there exists a straight line, preferably multiple straight lines, parallel to the engine axis X intersecting the straight line.

[0055] exist Figure 3 In this case, the orifice 5 extends axially over a shorter axial length C, preferably between 10% and 40% of the chord B of the impeller 7. The orifice is connected via cavity 2 to a second set of orifices 5B, which is located downstream of the first set of orifices 5A at an axial distance D, preferably between 10% and 50% of the chord B of the impeller 7. The orifices 5 of the second set of orifices 5B have a geometric profile independent of the orifices of the first set of orifices 5A (in... Figure 3In the case of a parallelogram with a zero inclination angle (i.e., a rectangle), the second set of orifices extends axially along an axial length C', preferably between 10% and 40% of the chord B of the blade 7. The second set of orifices has a circumferential width F', which is preferably greater than the circumferential space E' that separates the orifices in pairs. The orifices 5 of the second set of orifices 5B are primarily dedicated to capturing the airflow of the leaking vortex between the leading edge 7A and the trailing edge 7B of the blade 7, while the orifices 5 of the first set of orifices 5A are more dedicated to re-injecting this airflow upstream of the straightener 121 in a controlled manner. In particular, preferably, at least one section of the first orifice 51 of the second set of orifices 5B extends opposite to the blade between the leading edge 7A and the trailing edge 7B of the blade 7, while at least one section of the second orifice 52 of the first set of orifices 5A extends upstream (and downstream) of the leading edge 7A of the same blade 7, and these two sections are at least partially axially aligned in the aforementioned sense.

[0056] according to Figure 4A and Figure 4B In the preferred embodiment of the invention shown, the cavity 2 is divided into units 2A, 2B, and 2C, which are defined by edges 23 typically formed in the hub 6. Each such unit 2A, 2B, and 2C extends at least partially radially relative to the blades 7 and circumferentially at a principal angle β preferably between 10° and 60°, such that the upstream segment 22 of each such unit 2A is radially opposite to the space between the third blade (37) and the first blade (17) of the blade 7, while the downstream segment 21 of the unit 2A is radially opposite to the space between the first blade (17) and the second blade (27), which are different from the aforementioned blades. In this way, airflow from the leakage vortex is carried from downstream to upstream in unit 2A, from the space between two blades 7 of a pair of consecutive blades 7 to another space between two blades 7 of another pair of consecutive blades 7. For example, units 2A, 2B, and 2C can be implemented by embossing the walls of the hub 8.

[0057] In summary, the present invention relates to a sub-assembly 1 of a low-pressure compressor 120 for an aircraft turbine engine 100, the sub-assembly including a straightener 121 and a rotor hub 6, the straightener being equipped with suspended blades 7, the rotor hub including a cavity 2 covered by an inner shroud 3 facing the blades 7, an orifice 5 being fitted in the inner shroud 3 to allow airflow to flow from downstream to upstream of the low-pressure compressor 120.

[0058] The present invention has been described above with reference to specific embodiments, which are merely illustrative and should not be considered limiting. Generally, it will be apparent to those skilled in the art that the present invention is not limited to the examples described and / or illustrated above.

Claims

1. A sub-assembly for a low-pressure compressor of an aircraft turbine engine, the sub-assembly being oriented along an engine axis to extend from upstream to downstream, the sub-assembly comprising: - A rotor, the rotor including a hub extending axially and circumferentially around the engine axis; - Stator, the stator including a straightener equipped with blades that extend generally radially toward the hub; in: - The hub includes a recess oriented toward the engine axis; -The rotor includes an inner protective cover: The inner cover covers the recess to define a cavity between the hub and the inner cover. Each of the blades includes a free head that at least partially faces the inner shroud and / or the cavity; ○ The inner protective cover includes an opening that is in fluid communication with the cavity; Furthermore, the inner protective cover is formed by an annular plate portion.

2. The sub-component according to claim 1, characterized in that, The first orifice extends downstream of the leading edge of at least one of the blades, and the second orifice extends upstream of the leading edge.

3. The sub-component according to claim 2, characterized in that, At least one portion of the first orifice extends axially between the hub portion of the leading and trailing edges of the blade and / or the corresponding radial protrusions on the inner shroud.

4. The sub-component according to claim 2 or 3, characterized in that, The inner protective cover includes multiple pairs of first openings and second openings.

5. The sub-component according to any one of claims 1 to 3, characterized in that, The assembly of orifices is generally aligned circumferentially, and each orifice has a parallelogram profile extending axially at an angle (α) of up to 60° relative to the engine axis.

6. The sub-component according to claim 5, characterized in that, Each of the orifices in the set extends upstream and downstream of the leading edge of each blade.

7. The sub-component according to claim 6, characterized in that, Each of the orifices in the set includes: - The upstream end, located upstream of the leading edge, at an axial distance (A) from the leading edge, the axial distance being at most 25% of the chord of each blade; - Downstream end, the downstream end being located downstream of the leading edge at an axial distance (A') from the leading edge, the axial distance being at most 75% of the chord of each blade.

8. The sub-component according to claim 5, characterized in that, The axial length (C) of each orifice of the assembly is between 10% and 75% of the chord of each blade.

9. The sub-component according to claim 8, characterized in that, The orifice comprises two orifice groups aligned circumferentially, the axial distance (D) between the orifice groups and each other is between 10% and 50% of the chord of each blade, and one orifice group corresponds to the set, and the orifices of the other orifice group in the orifice group each have a different parallelogram profile.

10. The sub-component according to any one of claims 1 to 3, characterized in that, The orifice comprises two orifice groups aligned in the circumferential direction, the axial distance (D) between the orifice groups and each other being between 10% and 50% of the chord of each blade.

11. The sub-component according to any one of claims 1 to 3, characterized in that, The cavity is divided into units defined by edges, each unit extending in both the axial and circumferential directions along an axially inclined direction, such that each unit comprises: - Downstream section, the downstream section radially facing the space between the first and second blades of the blades that are successively aligned in the circumferential direction; - An upstream section, which radially faces another space between a third and a fourth blade in a circumferentially aligned impeller, at least one of the third and fourth blades being different from the first and second blades.

12. The sub-component according to any one of claims 1 to 3, characterized in that, The edges of the opening are beveled to create an open profile.

13. The sub-component according to any one of claims 1 to 3, characterized in that: - The recess is formed by at least one deformed portion of the wall of the hub.

14. A low-pressure compressor for an aircraft turbine engine, characterized in that, The low-pressure compressor includes a sub-component according to any one of claims 1 to 3.

15. An aircraft turbine engine, characterized in that, The aircraft turbine engine is equipped with a low-pressure compressor as described in claim 14.

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