Aircraft turbomachine compressor rectifier including leakage flow guidance grooves, and aircraft turbomachine including such a rectifier

Abstract

Aircraft turbomachine compressor stator (10) having an axis (X) and comprising a plurality of blades (12), a radially internal platform (14) having an upstream edge (14C) disposed, relative to the airfoil (12A) of each blade (12), on the leading edge side (12A1) of the blades (12), the upstream edge (14C) having as many grooves (18) as there are blades (12), the grooves (18) being regularly distributed around the axis (X), a single groove (18) being disposed between two adjacent blades (12). Figure for the abbreviation: Fig. 4

Description

Title of the invention: Aircraft turbomachine compressor rectifier comprising leakage flow orientation grooves, and aircraft turbomachine comprising such a rectifier. Technical field

[0001] The present disclosure relates to an aircraft turbomachine compressor rectifier, in particular a high-pressure compressor rectifier, and an aircraft turbomachine equipped with such a rectifier.

[0002] The term "aircraft turbomachine" refers to all gas turbine devices that produce motive power, among which a distinction is made in particular between turbojets, which provide the thrust necessary for propulsion by reaction to the high-speed ejection of hot gases, and turboshaft engines, in which motive power is supplied by the rotation of a drive shaft. For example, turboshaft engines are used as engines for helicopters. Turboprop engines (turboshafts driving a propeller) are also turboshaft engines used as aircraft engines. Prior art

[0003] An aircraft turbomachine compressor typically comprises a series of stages, each including a runner and a rectifier. The rectifier is located, in the direction of gas flow, immediately downstream of the runner, with the gases flowing from upstream to downstream. The main function of the compressor runner is to increase the pressure and temperature of the gases as it rotates. The rectifier's function is to straighten the gas flow along the axial direction of the aircraft turbomachine, thereby converting the kinetic energy of the gases acquired as they pass through the runner into static pressure (which reduces the gas flow velocity). In other words, in the direction of gas flow, the gas pressure downstream of a rectifier is greater than the gas pressure upstream of the rectifier.Such a pressure differential can generate an undesirable gas leak from downstream to upstream of the rectifier, particularly through the gaps around the radially internal platform of the rectifier. This phenomenon is illustrated in [Fig. 1] showing a compressor 105: a leakage flow EF interferes with the main gas flow F, this leakage flow EF bypassing the radially internal platform 106A of the rectifier 106, which is positioned between two rotating wheels 107A. This leakage flow EF can, in particular, disturb the main flow at the root of the rectifier blade, disrupt the boundary layers, disrupt the local flow directions of the main flow, and impose unwanted local stresses on the rotating wheel and rectifier blades.

[0004] A first solution to limit this leakage flow EF is to provide a sealing system 108 between the shaft 107B of the rotor 107, which includes the rotating wheels 107A, and the platform 106A. However, this system 108 is generally insufficient. A complementary solution is known from FR3084395. However, this latter solution is complex and expensive.

[0005] There is therefore a need to prevent disturbances to the main flow F caused by the leakage flow EF in an efficient and inexpensive manner. Description of the invention

[0006] One embodiment relates to an aircraft turbomachine compressor stator, the stator having an axis and comprising a plurality of blades regularly distributed circumferentially around the axis, each blade having an aerodynamic profile having a leading edge and a trailing edge, a radially internal platform having a radially internal face, a radially external face and an upstream edge disposed, relative to the aerodynamic profile of each blade, on the side of the leading edge of the blades, the upstream edge extending radially between the radially internal face and the radially external face, the upstream edge having as many grooves as blades, the grooves being regularly distributed around the axis, each groove having an inlet formed in the radially internal face and an outlet formed in the radially external face, a single outlet of a single groove being disposed between two adjacent blades,each groove extending radially between the radially internal face and the radially external face and having a non-zero axial depth over the entire radial extent of the upstream edge between the radially internal face and the radially external face.

[0007] Generally, and unless otherwise specified, in this description, upstream and downstream are defined with respect to the normal flow direction of the fluid (from upstream to downstream) through the rectifier, and more generally, the aircraft turbomachine. Furthermore, the axial direction corresponds to the direction of the rectifier's axis, which, once mounted on the aircraft turbomachine, corresponds to the axis of rotation of the aircraft turbomachine's rotating elements. A radial direction is a direction perpendicular to the axis. The circumferential or azimuthal direction corresponds to the direction describing a ring around the axis. The three directions axial, radial, and azimuthal correspond respectively to the directions defined by the ridge, radius, and angle in a cylindrical coordinate system. Finally, unless otherwise specified, the adjectives inside and outside are used with reference to a radial direction, such that the inside (i.e.,The radially inner part of an element is closer to the axis than the outer (i.e. radially outer) part of the same element.

[0008] Hereafter and unless otherwise indicated, "rectifier" means "aircraft turbomachine compressor rectifier".

[0009] The radially internal platform is disposed on the radially internal side with respect to the airfoil. For example, the straightener may comprise a radially internal platform and a radially external platform. The optional radially external platform is disposed on the radially external side with respect to the airfoil.

[0010] Depending on the circumferential direction, there is an alternation between the presence of a groove on the upstream edge and the presence of a blade. The grooves can form channels for channeling the leakage flow along the upstream edge, from the radially inner face to the radially outer face of the platform. The inlet of each groove is configured to receive the leakage flow from the radially inner face, while the outlet is configured to discharge and guide the leakage flow to the radially outer face of the inner platform, into the main gas stream, within the main flow. The grooves can be relatively simple to form and allow the leakage flow to be guided between the blades of the straightener (i.e.By diverting it away from the stator blades (and in particular from the leading edge of the stator blades), a certain degree of control over this leakage flow is achieved, significantly reducing its impact on the main flow. Furthermore, such grooves on the upstream edge can decrease the flexural stiffness of the radially internal platform, which can help dissipate energy and reduce potential local stresses.

[0011] For example, the grooves can be configured to guide a leakage flow parallel to the blade chords, the chords being considered at the level of the internal radial platform, the chord of a blade being a geometric segment connecting the leading edge to the trailing edge. Such a configuration can be relatively simple to implement and help guide the leakage flow between the stator blades, allowing some control of the leakage flow and significantly reducing the impact of the leakage flow on the main flow.

[0012] In certain embodiments, the connections between the different faces of the groove, at the bottom of the groove or with the rest of the platform, may include rounded edges or fillets. This helps to avoid potential local stress concentrations.

[0013] In certain embodiments, an inter-blade zone can be defined as a circumferential portion of the upstream edge delimited by the tangents to a skeleton line at the leading edge of two adjacent blades, the outlet of each groove being arranged circumferentially, in whole or in part, within a central portion of an inter-blade zone, the central portion extending over approximately 50%, for example, over approximately 25% of the total extent along the circumferential direction of the inter-dawn zone. For example, the central portion can extend over an area between 20% and 50% of the total extent along the circumferential direction of the inter-dawn zone.

[0014] It is understood that the edge zone extending between the two tangents from the skeleton line to the leading edge of two adjacent blades delimits the inter-blade zone of the upstream edge. Naturally, the inter-blade zone is the smallest circumferential portion extending between these two tangents. The skeleton line of a blade can also be viewed as the neutral axis of the blade. The skeleton line can be considered at the junction between the blade on the radially outer face and the radially inner face. The central portion is the portion centered on the midpoint of the inter-blade zone, that is, equidistant from each of the two adjacent blades. The central portion has a circumferential extent of approximately 50%, for example, approximately 25% of the total circumferential extent of the inter-blade zone.In other words, the central portion is 25%, for example 50%, of the total circumferential extent of the inter-blade area from each of two adjacent blades. There is a single groove outlet per inter-blade area, and this outlet extends wholly or partially into the central area. Such a configuration can be relatively simple to implement and can help guide the leakage flow between the stator blades, allowing some control of the leakage flow and significantly reducing its impact on the main flow.

[0015] In some embodiments, each groove may have a maximum circumferential width of between 5.00 mm (five millimeters) and 20.00 mm (twenty millimeters).

[0016] For example, the maximum circumferential width can be constant over the entire radial extent of the groove. This can facilitate the manufacturing and control of the grooves. As another example, the maximum circumferential width can vary along the radial direction, for example, increasing or decreasing. This can allow for greater flexibility in managing the straightening of the leakage flow.

[0017] In some embodiments, each groove may have an axial depth of between 5.00 mm (five millimeters) and 10.00 mm (ten millimeters).

[0018] Such an axial depth can be relatively simple to achieve, can preserve the mechanical strength of the platform while effectively helping to guide the leakage flow between the blades of the rectifier, allowing some control of the leakage flow and significantly reducing the impact of the leakage flow on the main flow.

[0019] In certain embodiments, the axial depth of each groove may be increasing, for example strictly increasing, from input to output.

[0020] The groove depth can increase in steps, or increase strictly from the inlet to the outlet. Such a configuration can be relatively simple to implement, can preserve the mechanical strength of the platform while effectively guiding the leakage flow between the stator blades, allowing for some control of the leakage flow and significantly reducing the impact of the leakage flow on the main flow.

[0021] In certain embodiments, considered in a plane perpendicular to the axis, each groove can be inclined with respect to a radial direction by an angle between 25° (twenty-five degree angle) and 45° (forty-five degree angle).

[0022] For example, each groove, for example each groove outlet, can be arranged circumferentially between an extrados of a first blade and an intrados of a second blade adjacent to the first blade, each groove being inclined towards the intrados of the second blade.

[0023] The angle considered is the smallest angle formed between the groove, or the central geometric axis of the groove, and the radius considered, for example, the radius passing through the midpoint of the groove inlet, for example, considered in the plane perpendicular to the axis comprising the central geometric axis of the groove. Such a configuration can be relatively simple to implement and effectively contribute to guiding the leakage flow between the stator blades, allowing some control of the leakage flow and significantly reducing the impact of the leakage flow on the main flow.

[0024] In some embodiments, each groove has two circumferentially opposed lateral walls, at least one lateral wall being inclined with respect to the axis.

[0025] Such an inclination of the walls can help guide a leakage flow with a certain circumferential component (or orientation). Such a configuration can be relatively simple to implement and effectively help guide the leakage flow between the blades of the rectifier, allowing some control of the leakage flow and significantly reducing the impact of the leakage flow on the main flow.

[0026] In certain embodiments, each groove can be arranged circumferentially between an extrados of a first blade and an intrados of a second blade adjacent to the first blade, only the lateral wall arranged on the side of the extrados of the first blade being inclined with respect to the axis (while the lateral wall arranged on the side of the intrados of the second blade can be parallel to the axis).

[0027] Such a configuration can be well suited to help guide a leakage flow with a strong circumferential component (or orientation). Such a configuration can be relatively simple to implement and contribute effectively effectively guide the leakage flow between the blades of the rectifier, allow some control of the leakage flow and significantly reduce the impact of the leakage flow on the main flow.

[0028] One embodiment relates to an aircraft turbomachine, for example an aircraft propulsion unit, for example a turbojet, comprising a compressor rectifier according to any one of the embodiments described in this exposition.

[0029] In some embodiments, the compressor may include a compressor wheel immediately upstream of the rectifier, the gases flowing from upstream to downstream within the aircraft turbomachine, the axial depth of each of the grooves being equal to one-third of an axial distance between the compressor wheel and the rectifier, plus or minus ten percent.

[0030] The axial distance between the moving wheel and the rectifier is sometimes called the "air gap" by those skilled in the art. Such a depth ensures a balance between efficiency and mechanical impact on the platform.

[0031] In some embodiments, the compressor may be a high-pressure compressor, the aircraft turbomachine being further equipped with a low-pressure compressor separate from the high-pressure compressor.

[0032] The rectifier according to the present exposition is particularly well suited for high-pressure compressors, where the phenomenon of leakage flow is very likely to occur and negatively impact the main gas flow within the compressor. Brief description of the drawings

[0033] The purpose of this presentation and its advantages will be better understood upon reading the detailed description below of various embodiments given by way of non-limiting examples. This description refers to the attached figure pages, on which:

[0034] [Fig-1] [Fig.1] represents an aircraft turbomachine compressor of the state of the technique,

[0035] [Fig.2] Fig.2 represents an aircraft equipped with an aircraft turbomachine,

[0036] [Fig.3] The [Fig.3] schematically represents the aircraft turbomachine of the aircraft of the [Fig.2],

[0037] [Fig.4] The [Fig.4] schematically represents a part of the high-pressure compressor of the aircraft turbomachine of the [Fig.3],

[0038] [Fig.5] [Fig.5] represents the rectifier in perspective, seen along arrow V of [Fig.4],

[0039] [Fig.6] [Fig.6] represents the rectifier, seen according to arrow VI of [Fig.5],

[0040] [Fig.7] Fig.7 represents a variant of the rectifier,

[0041] [Fig.8] Fig.8 represents another variant of the rectifier,

[0042] [Fig.9] Fig.9 represents another variant of the rectifier,

[0043] [Fig. 10] Fig. 10 represents another variant of the rectifier,

[0044] [Fig. 11] Fig. 11 represents yet another variant of the rectifier, and

[0045] [Fig. 12] Fig. 12 schematically represents the guidance of a reflow circulation through the grooves of the rectifier according to the present exposition. Description of the implementation methods

[0046] Figure 2 represents an aircraft 100, in this example an airplane, equipped with two aircraft turbomachines 50, in this example two propulsion units 50, in this example two turbojet engines 50, namely one aircraft turbomachine 50 per wing 101, with only one aircraft turbomachine 50 and one wing 101 being shown in Figure 2. According to one variant, the aircraft 100 can be equipped with more than one aircraft turbomachine 50 per wing 101, each wing 101 being fitted with the same number of aircraft turbomachines 50.

[0047] Figure 3 shows a schematic cross-sectional view of the aircraft turbomachine 50, according to Figure 2. The aircraft turbomachine 50 comprises a fan 52, which may be shrouded or unshrouded, and a gas generator 54. In this example, the gas generator 54 comprises, from upstream to downstream, a compressor 54A (or compressor section 54A), a combustion chamber 54B, and a turbine 54C (or turbine section 54C). The fan 52 may be driven in rotation directly by a shaft of the gas generator 54, for example, a shaft of a low-pressure body, or via a speed reducer (not shown).

[0048] The gas generator 54 may be of the twin-spool type and comprise a low-pressure spool 60A and a high-pressure spool 60B. The low-pressure spool 60A may comprise a low-pressure compressor 62A rotationally coupled to a low-pressure turbine 66A via a low-pressure shaft 63A. The high-pressure spool 60B may comprise a high-pressure compressor 62B located downstream of the low-pressure compressor 62A and upstream of the combustion chamber 54B, and a high-pressure turbine 66B located downstream of the combustion chamber 54B and upstream of the low-pressure turbine 66A, and rotationally coupled to the high-pressure compressor 62B via a high-pressure shaft 63B. The compressor 54A of the gas generator 54 may comprise the low-pressure and high-pressure compressors 62A and 62B. The turbine 54C of the gas generator 30 can include the low and high pressure turbines 66A and 66B. The [Fig.3] is schematic, each compressor and each turbine being able to have one or more stages, each stage comprising a rotating wheel and a stator or rectifier.

[0049] Figure 4 partially represents the high-pressure compressor 62B. In particular, Figure 4 represents a stage 70 of the high-pressure compressor 62B, the stage 70 comprising a rotating wheel 72 upstream of an aircraft turbomachine compressor stator 10. Downstream of the stator 10, Figure 4 represents the rotating wheel 72 of the downstream stage. The vanes 12 of the stator 10 described below may have fixed or variable pitch.

[0050] In this example, as shown in Figures 4, 5, and 6, the aircraft turbomachine compressor stator 10 has an X-axis (coinciding with the X-axis of the aircraft turbomachine 50) and comprises a plurality of blades 12 regularly distributed circumferentially (see circumferential direction C) around the X-axis. Each blade 12 has an airfoil 12A having a leading edge 12A1 and a trailing edge 12A2, an internal radial platform 14, and an optional external radial platform 16. The internal radial platform 14 has an internal radial face 14A, an external radial face 14B, and an upstream edge 14C disposed, relative to the airfoil 12A of each blade 12, on the side of the leading edge 12A1 of the blades 12. 14C extending radially (see radial direction R) between the radially internal face 14A and the radially external face 14B,the upstream edge 14C having as many grooves 18 as blades 12, the grooves 18 being regularly distributed around the axis X, each groove 18 having an inlet 18A formed in the radially internal face 14A and an outlet 18B formed in the radially external face 14B, a single outlet 18B of a single groove 18 being disposed between two adjacent blades 12, each groove 18 extending radially between the radially internal face 14A and the radially external face 14B and having a non-zero axial depth PA over the entire radial extent ER of the upstream edge 14C between the radially internal face 14A and the radially external face 14B. In the example shown in Figures 4, 5 and 6, the upstream edge 14C extends parallel to the radial direction R,but can, according to an unrepresented variant, be inclined with respect to the radial direction R. The compressor 62B may include a sealing system 19 which cooperates with the radially internal face 14A of the radially internal platform 14 of the rectifier 10 in order to limit leakage flow from downstream to upstream of the rectifier 10.

[0051] Figure 6 shows the external radial face 14B of the platform 14, viewed radially from the outside (according to arrow VI in Figure 5). An inter-blade zone (ILZ) is defined as a circumferential portion of the upstream edge 14C delimited by tangents T to a skeleton line LS at the leading edge 12A1 of two adjacent blades 12. The outlet 18B of each groove 18 can be arranged circumferentially, in whole or in part, within a central portion PC of an ILB, the central portion PC extending over 50% (represented by the double arrow associated with the The reference sign PC) of the total extent (represented by the double arrow associated with the reference sign ZIR) along the circumferential direction C of the inter-blade zone ZIR. The central portion PC is centered on the midpoint MI along the circumferential direction C of the inter-blade zone ZIR.

[0052] Each groove 18 can have an axial depth PA of between 5.00 mm and 10.00 mm. For example, with reference to [Fig.4], the compressor 62B includes a compressor wheel 72 immediately upstream of the rectifier 10, the gases flowing from upstream to downstream within the aircraft turbomachine 50, the axial depth PA of each of the grooves 18 can be equal to one-third of an axial distance D between the compressor wheel 72 and the rectifier 10, plus or minus ten percent.

[0053] In the example shown in Figures 4, 5 and 6, the axial depth PA of each of the grooves can be constant over the entire radial extent ER of the groove 18 (or the upstream edge 14C). According to a variant shown in [Fig. 7], the axial depth PA of each groove 18 can be increasing, in this example strictly increasing, from the inlet 18A to the outlet 18B of the grooves 18.

[0054] In the examples shown in Figures 4 to 7, each of the grooves 18 can extend parallel to the radial direction R. According to a variant shown in Figure 8, each groove 18 can be inclined with respect to the radial direction R by an angle Q between 25° and 45°, considered in a plane perpendicular to the axis X (in the example of Figure 8, the plane in which the angle Q is represented and comprising the two axes delimiting the angle Q). For example, each groove 18, for example the outlet 18B of each groove 18, can be arranged circumferentially between an extrados 12B of a first blade 12-1 and an intrados 12C of a second blade 12-2 adjacent to the first blade 12-1, each groove 18 being inclined towards the intrados 12C of the second blade 12-2.

[0055] In the examples shown in Figures 4 to 8, the two circumferentially opposing side walls 18C1 and 18C2 of each groove 18 may extend parallel to the X-axis. According to a variant shown in [Fig. 9], at least one side wall, in this example both side walls 18C1 and 18C2, may be inclined with respect to the X-axis. For example, the walls 18C1 and 18C2 may be flat, as shown in [Fig. 9], and form an angle of between 5° (five degrees) and 50° (fifty degrees) with the X-axis. According to a variant not shown, the walls 18C1 and / or 18C2 may have a curved profile from the upstream edge 14C to the bottom of the groove 18, the curved profile being either concave or convex.

[0056] According to another variant shown in [Fig. 10], each groove 18 can be arranged circumferentially between an extrados 12B of a first blade 12-1 and an intrados 12C of a second blade 12-2 adjacent to the first blade 12-1, only the lateral wall 18C1 arranged on the side of the extrados 12B of the first blade 12-1 can be inclined with respect to the X axis while the lateral wall 18C2 arranged on the side of the intrados 12C of the second blade 12-2 can be parallel to the X axis.

[0057] In the examples shown in Figures 4 to 10, the connections between the different faces of each groove 18 and between the faces of each groove and the faces of the platform 14 are schematically angular. According to one embodiment, all or part of these connections may be rounded or filleted. According to one embodiment shown in [Fig. 11], the connection RI between the bottom of each groove 18 and the radially internal face 14A of the platform 14 and the connection R2 between the bottom of each groove 18 and the radially external face 14B of the platform 14 are rounded.

[0058] The arrows in [Fig. 12] represent the flow of the leakage EF between the blades 12, resulting from the guidance produced by the grooves 18. Thanks to the grooves 18, the leakage EF is deflected from the blades 12, avoids the leading edges 12A1, remains mostly distant from the intrados and extrados 12B and 12C and is guided between the blades 12. For example, the grooves 18 are configured to guide the leakage EF parallel to the chords C of the blades 12, at the level of the platform 14 (see [Fig. 10], the chord C of a blade 12 being a geometric segment connecting the leading edge 12A1 to the trailing edge 12A2).

[0059] Although the present invention has been described with reference to specific embodiments, it is evident that modifications and changes can be made to these examples without departing from the general scope of the invention as defined by the claims. In particular, individual features of the various embodiments illustrated / mentioned can be combined in additional embodiments. Therefore, the description and drawings should be considered in an illustrative rather than restrictive sense. In particular, the groove configurations illustrated in Figures 4 to 12 can be combined in whole or in part.

Claims

Claims

1. A rectifier (10) for an aircraft turbomachine compressor, the rectifier (10) having an axis (X) and comprising a plurality of blades (12) regularly distributed circumferentially around the axis (X), each blade (12) having an aerodynamic profile (12A) having a leading edge (12A1) and a trailing edge (12A2), a radially internal platform (14) having a radially internal face (14A), a radially external face (14B) and an upstream edge (14C) arranged, relative to the aerodynamic profile (12A) of each blade (12), on the side of the leading edge (12A1) of the blades (12), the upstream edge (14C) extending radially between the radially internal face (14A) and the radially external face (14B), the upstream edge (14C) having as many grooves (18) than blades (12), the grooves (18) being regularly distributed around the axis (X),each groove (18) having an inlet (18A) formed in the radially inner face (14A) and an outlet (18B) formed in the radially outer face (14B), a single outlet (18B) of a single groove (18) being arranged between two adjacent blades (12), each groove (18) extending radially between the radially inner face (14A) and the radially outer face (14B) and having a non-zero axial depth (PA) over the entire radial extent (ER) of the upstream edge (14C) between the radially inner face (14A) and the radially outer face (14B).,

2. Aircraft turbomachine compressor rectifier (10) according to claim 1, wherein an inter-blade zone (ZIR) is defined as a circumferential portion of the upstream edge (14C) delimited by the tangents (T) to a skeleton line (LS) at the leading edge (12A1) of two adjacent blades (12), the outlet (18B) of each groove (18) being arranged circumferentially in whole or in part in a central portion (PC) of an inter-blade zone (ZIR), the central portion (PC) extending over approximately 50%, for example over approximately 25%, of the total extent in the circumferential direction (C) of the inter-blade zone (ZIR).

3. Aircraft turbomachine compressor rectifier (10) according to claim 1 or 2, in which each groove (18) has an axial depth (PA) of between 5.00 mm and 10.00 mm.

4. Aircraft turbomachine compressor rectifier (10) according to claim 3, wherein the axial depth (PA) of each groove (18) is increasing, for example strictly increasing, from the input (18A) to the output (18B).

5. Aircraft turbomachine compressor rectifier (10) according to any one of claims 1 to 4, in which, considered in a plane perpendicular to the axis (X), each groove (18) is inclined relative to a radial direction (R) by an angle (@) between 25° and / ICO

6. 45. Aircraft turbomachine compressor rectifier (10) according to any one of claims 1 to 5, in which each groove (18) has two circumferentially opposite side walls (18C1, 18C2), at least one side wall (18C1, 18C2) being inclined relative to the axis (X).

7. Aircraft turbomachine compressor rectifier (10) according to claim 6, wherein each groove (18) is arranged circumferentially between an extrados (12B) of a first blade (12-1) and an intrados (12C) of a second blade (12-2) adjacent to the first blade (12-1), only the side wall (18C1) arranged on the side of the extrados (12B) of the first blade (12-1) being inclined relative to the axis (X).

8. An aircraft turbomachine (50) comprising a compressor (62B), the compressor (62B) comprising a compressor rectifier (10) according to any one of claims 1 to 7.

9. An aircraft turbomachine (50) according to claim 8, wherein the compressor (62B) comprises a moving compressor wheel (72) immediately upstream of the stator (10), the gases flowing from upstream to downstream within the aircraft turbomachine (50), the axial depth (PA) of each of the grooves (18) being equal to one third of an axial distance (D) between the moving compressor wheel (72) and the stator (10), plus or minus ten percent.

10. Aircraft turbomachine (50) according to claim 8 or 9, wherein the compressor is a high pressure compressor (62B), the aircraft turbomachine (50) being further provided with a low pressure compressor (62A) separate from the high pressure compressor (62B).

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