FR3083824A1 - PART FOR AN AIRCRAFT TURBOMACHINE ROTOR COMPRISING AN ELLIPTICALLY FULL OILING ORIFICE - Google Patents

Abstract

A part (50) for an aircraft turbomachine rotor comprises an annular wall (52) centered on an axis of rotation (54) of the part, as well as at least one deoiling orifice (40) passing through the annular wall (52 ) right through and having an edge (42) with an elliptical section defined by a major axis parallel to a circumferential direction (C) of the part and a minor axis included in a plane comprising the axis of rotation ( 54) of the part, such that a ratio of a length of the major axis of the ellipse divided by a length of the minor axis of the ellipse is between 1.45 and 1.75 inclusive. Such a shape of the deoiling orifice makes it possible to confer an increased service life on the annular wall (52) of the part.

Description

PART FOR AN AIRCRAFT TURBOMACHINE ROTOR COMPRISING AN ELLIPTICALLY DECOILED ORIFICE

DESCRIPTION

TECHNICAL AREA

The present invention relates to the field of turbomachines intended to equip aircraft and relates more particularly to a part intended to form part of a rotor of such a turbomachine, this part comprising an annular wall centered on an axis of rotation of the part, the annular wall comprising a deoiling orifice.

PRIOR STATE OF THE ART

Some of the discs forming the rotors of the compressors and turbines of aircraft turbomachines comprise an annular wall and one or more orifices called "deoiling" formed through the annular wall. Such an annular wall indeed delimits an enclosure in which prevails an oil mist originating from the lubrication of one or more bearings arranged within said enclosure. Each deoiling orifice allows the air loaded with oil to circulate from said enclosure to another space in which this air is recovered by a device provided for this purpose.

In known turbomachinery, the oil removal orifices are oblong holes (also called “slotted holes” in English terminology). Such an oblong hole is defined as being a hole having an elongated shape composed of a rectangular or square median part and two half-discs arranged symmetrically with respect to each other, on either side of the median part. , at the ends of the hole. As recalled in the applicant's prior document FR3028781, such a shape is in fact considered to be capable of providing a good lifetime for the annular wall.

However, the inventors have set themselves the objective of optimizing the shape of such deoiling orifices in order to further increase the life of the annular wall.

STATEMENT OF THE INVENTION

To this end, the invention proposes a part for the rotor of an aircraft turbomachine, comprising an annular wall centered on an axis of rotation of the part, as well as at least one deoiling orifice passing right through the annular wall, the part being characterized in that the deoiling orifice has an edge with an elliptical section defined by a large axis parallel to a circumferential direction of the part and a small axis included in a plane comprising the axis of rotation of the piece, and in which a ratio B / A, of a length B of the major axis of the ellipse divided by a length A of the minor axis of the ellipse, is between 1.45 and 1.75 inclusive.

The inventors have in fact determined that such a shape makes it possible to considerably increase the life of the annular wall in which the deoiling orifice or each deoiling orifice is formed. Such an improvement is reflected in particular by a reduction in the Von Mises constraint, as will appear more clearly in the following.

Preferably, the edge has chamfered angles.

In preferred embodiments of the invention, the part consists of a disc for an aircraft turbomachine compressor, in particular a one-piece bladed disc.

The invention also relates to a compressor for an aircraft turbomachine, comprising at least one part of the type described above.

Finally, the invention relates to a turbomachine for an aircraft, comprising at least one part of the type described above.

In preferred embodiments, the aircraft turbomachine comprises at least one compressor of the type described above.

In a preferred embodiment of the invention, the turbomachine is a double-body turbomachine, and the compressor is a high pressure compressor.

BRIEF DESCRIPTION OF THE DRAWINGS

The invention will be better understood, and other details, advantages and characteristics thereof will appear on reading the following description given by way of non-limiting example and with reference to the appended drawings in which:

- Figure 1 is a schematic view in axial section of a turbomachine;

- Figure 2 is a schematic view in axial section of a bladed disc of a compressor of the turbomachine of Figure 1, constituting a part according to a preferred embodiment of the invention;

- Figure 3 is a schematic view in axial section of a portion of a pin belonging to the bladed disc of Figure 2;

- Figure 4 is a top view of part of the pin of Figure 3, illustrating a deoiling orifice formed in the pin;

- Figure 5 is a graph showing the evolution of the Von Mises o _VM stress (on the ordinate) as a function of a B / A ratio of dimensions of the oil removal orifice;

FIG. 6 is a table recording the test parameters which made it possible to obtain the graph of FIG. 5.

Throughout these figures, identical references may designate identical or analogous elements.

DETAILED DESCRIPTION OF PREFERRED EMBODIMENTS

FIG. 1 illustrates a turbomachine 10 for aircraft, generally comprising a fan 12 intended for the suction of an air flow dividing downstream of the fan into a primary flow circulating in a primary flow flow channel , hereinafter called the primary vein PF, within a core of the turbomachine, and a secondary flow bypassing this core in a secondary flow flow channel, hereinafter called the secondary vein SF.

The turbomachine is for example of the double flow and double body type. The heart of the turbomachine thus generally comprises a low pressure compressor 14, a high pressure compressor 16, a combustion chamber 18, a high pressure turbine 20 and a low pressure turbine 22.

The respective rotors of the high pressure compressor and the high pressure turbine are connected by a shaft called "high pressure shaft", while the respective rotors of the low pressure compressor and the low pressure turbine are connected by a shaft called "low pressure shaft" In a well-known manner.

The turbomachine is faired by a nacelle 24 surrounding the secondary stream SF. Furthermore, the rotors of the turbomachine are rotatably mounted about a longitudinal axis 28 of the turbomachine.

Throughout this description, the axial direction X is the direction of the longitudinal axis 28. The radial direction R is at all points a direction orthogonal to the longitudinal axis 28 and passing through the latter, and the circumferential direction or tangential C is at any point a direction orthogonal to the radial direction R and to the longitudinal axis 28. The terms "internal" and "external" refer respectively to a relative proximity, and a relative distance, of an element relative to to axis 28. Finally, the “upstream” and “downstream” directions are defined by reference to the general direction of the flow of gases in the primary PF and secondary SF streams of the turbomachine, in the axial direction X.

The high pressure compressor 16 typically comprises several successive stages, each comprising a fixed blade, called a rectifier, belonging to a stator of the turbomachine, and a bladed disc connected to the high pressure shaft and therefore belonging to the rotor of the high pressure compressor 16.

The bladed disc of at least one stage of the high pressure compressor 16 is preferably a one-piece bladed disc, also called "DAM", or even "blisk" in English terminology.

FIG. 2 shows a radially internal part of such a bladed disc 30, and in particular makes it possible to see a pin 32, that is to say an annular wall extending downstream from a side downstream 34 of the bladed disc 30, also partially visible in FIGS. 3 and 4.

The journal 32 comprises for example a cylindrical portion 36, one upstream end of which is connected to the downstream side 34 of the bladed disc 30, and one downstream end of which is provided with a flange forming an annular flange 38.

The annular flange 38 comprises for example through orifices 39 (FIG. 3) intended for the passage of fixing members intended to connect the annular flange 38 to another part of the rotor of the high pressure compressor.

The pin 32 has, in its cylindrical portion 36, at least one oil removal orifice 40 (Figures 2-4).

In the more general terminology of the invention, the bladed disc 30 constitutes a part 50 for the rotor of an aircraft turbomachine, and the cylindrical portion 36 of the journal 32 constitutes an annular wall 52, centered on an axis of rotation 54 of the part , which corresponds to the longitudinal axis 28 of the turbomachine. The axial X, radial R, and circumferential C directions of the turbomachine therefore constitute respectively axial, radial, and circumferential directions of the part 50.

The part 50 may nevertheless be another type of part of a turbomachine rotor, without departing from the scope of the invention.

According to the invention, the oil removal orifice 40 has an edge 42 with an elliptical section (FIG. 4) defined by a major axis 44 parallel to the circumferential direction C, and a minor axis 46 included in a plane P comprising the axis of rotation 54 of the part.

The section considered is of course in a plane orthogonal to a central axis 48 of the oil removal orifice 40 (Figures 2 and 4).

In addition, a ratio B / A, in which B is the length of the major axis 44 of the ellipse and A is the length of the minor axis 46 of the ellipse, is between 1.45 and 1.75 inclusive (figure 4).

The inventors have in fact determined that such a shape makes it possible to considerably increase the life of the annular wall 52 in which the deoiling orifice 40 is formed.

Such an improvement is reflected in particular by a reduction in the Von Mises stress, as shown in FIG. 5.

FIG. 5 is a graph showing, for a particular application, the evolution of the Von Mises o _VM constraint in megapascal (on the ordinate) as a function of the B / A ratio for a series of eleven tests for different shapes of the orifice deoiling (curve 60), the main parameters of which are recorded in the table in FIG. 6, and for a part of the prior art comprising an oblong deoiling orifice (right 62). The table in FIG. 6 also includes a column in which the values of the difference I (in percent) are recorded between the B / A ratio for each of the eleven tests and the B / A ratio for the piece of the prior art .

In a preferred embodiment, the deoiling orifice 40 has a shape such that the B / A ratio is equal to 1.58.

For this purpose, the deoiling orifice 40 advantageously has a shape such that the length A of the minor axis 46 is equal to 3 mm and the length B of the major axis 44 is equal to 4.75 mm, in the case of a annular wall having an outside diameter of 310.3 mm and a thickness of approximately 3 mm near the deoiling orifice 40.

In this case, the central axis 48 of the oil removal orifice 40 is preferably the space of the annular flange 38, with a distance E equal to 12.45 mm (FIG. 3).

Furthermore, each deoiling orifice 40 is generally the subject of a process known as “angle breaking”, consisting of machining the edges situated at the intersection of the orifice and, respectively, of the internal surfaces 52A and external 52B (FIG. 2) of the annular wall 52 in which the orifice is formed. The purpose of such a method is in particular to limit the stress concentration phenomena at the edges of the deoiling orifice 40, and to improve the aerodynamic characteristics of the air flow through the orifice and along the annular wall 52.

The machining preferably consists in forming a chamfer inclined relative to the central axis 48 of the deoiling orifice 40, for example at an angle of approximately 45 degrees.

Claims (8)

1. Part (50) for an aircraft turbomachine rotor, comprising an annular wall (52) centered on an axis of rotation (54) of the part, as well as at least one deoiling orifice (40) passing through the annular wall (52) right through, the part being characterized in that the deoiling orifice (40) has an edge (42) of elliptical cross section defined by a major axis (44) parallel to a circumferential direction ( C) of the part and a minor axis (46) included in a plane (P) comprising the axis of rotation (54) of the part, and in which a ratio B / A, of length B of the major axis ( 44) of the ellipse divided by a length A of the minor axis (46) of the ellipse, is between 1.45 and 1.75 inclusive. 2. Part according to claim 1, wherein the edge (42) has chamfered angles. 3. Part according to claim 1 or 2, consisting of a disc (30) for a compressor (16) of an aircraft turbomachine (10). 4. Part according to claim 3, the disc (30) being a one-piece bladed disc. 5. Compressor (16) for an aircraft turbomachine (10), comprising at least one part (50) according to any one of claims 1 to 4. 6. Turbomachine (10) for aircraft, comprising at least one part (50) according to any one of claims 1 to 4. 7. A turbomachine (10) for an aircraft, comprising at least one compressor (16) according to claim 5. 8. The turbomachine according to claim 7, the turbomachine being a double-body turbomachine, and said compressor (16) being a high pressure compressor.