FR3118891A1 - Manufacture of a turbine injector by laser fusion on a powder bed - Google Patents

Abstract

The present invention relates to a method of manufacturing a turbomachine turbine injector by laser powder bed fusion comprising the following steps: S1: manufacturing an annular flange (202) coaxial with a normal manufacturing axis (X-X) to a plate (3) on which the powder bed is placed, S2: simultaneously with step S1, the manufacture of a sacrificial support (4) arranged radially with respect to the annular flange (202), then S3: the fabricating a wall (204) extending radially from an axial surface of the annular flange (202), said wall (204) being integral with the annular flange (202) and with the sacrificial support (4). Figure for the abstract: Fig. 1

Description

Manufacture of a turbine injector by laser fusion on a powder bed

FIELD OF THE INVENTION

The present invention relates to the manufacture of a mechanical part by laser fusion on a powder bed.

More specifically, the present invention relates to the manufacture of a turbomachine turbine injector by laser melting on a powder bed.

STATE OF THE ART

Documents FR 2 841 591 and FR 2 937 371 are known examples of a turbomachine high-pressure turbine injector. The injector is configured in particular to allow the ventilation of a disc of the high pressure turbine.

Such an injector is generally manufactured by assembling several distinct parts, manufactured separately, which is long and costly. Furthermore, the known injectors have too great a mass.

It was therefore envisaged to manufacture a one-piece injector by additive manufacturing, and more particularly by laser fusion on a powder bed. Such a process makes it possible to reduce the mass, the time and the manufacturing cost of the injector. However, the manufacture of a turbine injector by laser fusion on a powder bed presents difficulties.

One of them comes from the fact that an injector generally comprises an annular flange coaxial with the axis of the turbomachine and a wall extending radially from an axial surface of the annular flange. However, during manufacture by laser fusion on a powder bed, the axis of manufacture of the injector, that is to say the axis normal to the plate intended to receive the bed of powder, is generally coincident with the shaft of the turbomachine.

To manufacture the injector by laser fusion on a powder bed, it would therefore be necessary to manufacture a wall extending along the same plane as the powder bed. Concretely, this would require manufacturing by fusion an element having an extended surface and resting only on powder. Then, a large area of this element would not adhere to the last fabricated layer and would therefore be susceptible to significant deformation. This deformed element would also be carried away by the powder scraper during the following passages.

There is therefore a need to overcome the disadvantages of the state of the art in this respect.

An object of the invention is to manufacture by laser fusion on a powder bed an injector comprising an annular flange coaxial with a manufacturing axis normal to a plate on which the powder bed is placed, and a wall extending radially from an axial surface of the annular flange.

It is for this purpose proposed, according to a first aspect of the invention, a method for manufacturing a turbomachine turbine injector by laser fusion on a powder bed comprising the following steps:
S1: manufacture of an annular flange coaxial with a manufacturing axis normal to a plate on which the powder bed is placed,
S2: simultaneously with step S1, the manufacture of a sacrificial support arranged radially with respect to the annular flange, then
S3: the manufacture of a wall extending radially from an axial surface of the annular flange, said wall being integral with the annular flange and with the sacrificial support.

Thanks to the presence of the sacrificial support, the wall does not collapse under its own weight since, during the deposition of the layer of powder intended to form the wall, the latter rests on the annular flange and the sacrificial support. In addition, during manufacturing, the presence of the sacrificial support makes it possible to limit the deformations of the injector and facilitates the evacuation of the heat generated by the laser fusion.

Advantageously, but optionally, the method according to the invention may comprise at least one of the following characteristics, taken alone or in combination:
- the sacrificial support is annular and coaxial with the manufacturing axis,
- the sacrificial support comprises a first end and a second end, the wall being integral with the second end and a thickness of the sacrificial support being variable between the first end and the second end,
- the sacrificial support comprises a ferrule which extends at a distance from the annular flange, the ferrule being connected to the wall by a junction,
- the junction has a weakness in order to facilitate removal of the sacrificial support from the injector once the manufacture of the wall has been completed,
- during step S2, a first end of the sacrificial support, opposite the junction, is crenellated in order to facilitate the circulation of the powder towards the outside of the injector,
- the sacrificial support further comprises a honeycomb structure extending from the axial surface of the annular flange,
- the sacrificial support extends from the axial surface of the annular flange and comprises a honeycomb structure, and
- The method further comprises a step S4 of removing the sacrificial support from the injector once step S3 has been completed, preferably by machining and/or by trimming the sacrificial support.

According to a second aspect of the invention, there is proposed a turbomachine turbine injector manufactured by means of a method as described above.

DESCRIPTION OF FIGURES

Other characteristics, objects and advantages of the invention will emerge from the description which follows, which is purely illustrative and not limiting, and which must be read in conjunction with the appended drawings in which:

There illustrates a schematic sectional view of a turbomachine.

There illustrates a schematic sectional view of part of a turbomachine.

There illustrates a perspective view of one embodiment of a disassembled turbine injector.

There illustrates a cross-sectional and perspective view of one embodiment of an assembled turbine injector.

There illustrates a longitudinal sectional view of the .

There illustrates a perspective view of part of one embodiment of a turbine injector.

There illustrates a perspective view of one embodiment of a turbine injector.

There illustrates a cross-sectional and perspective view of the turbine injector illustrated in the .

There is a flowchart of steps of a manufacturing method according to an embodiment of the invention.

There illustrates part of a turbine injector on a plate during a manufacturing process according to one embodiment of the invention.

There illustrates a turbine injector during a manufacturing process according to an embodiment of the invention.

There illustrates part of a turbine injector on a plate during a manufacturing process according to one embodiment of the invention.

There illustrates part of a turbine injector on a plate during a manufacturing process according to one embodiment of the invention.

In all the figures, similar elements bear identical references.

DETAILED DESCRIPTION OF THE INVENTION

Turbomachinery

With reference to the , a turbomachine 1 has a longitudinal axis XX and comprises a fan 10, a compressor section 12, a combustion chamber 14 and a turbine section 16 which are capable of being rotated about the longitudinal axis XX with respect to to a casing 18 of the turbomachine 1.

In operation, the fan 10 draws in a flow of air, a portion of which is successively compressed within the compressor section 12, ignited within the combustion chamber 14, and expanded within the turbine section 16 before to be ejected from the turbomachine 1. In this way, the turbomachine 1 generates thrust. This thrust can also be put to the benefit of an aircraft (not shown) on which the turbomachine 1 is attached and fixed.

In the present text, the upstream and the downstream are defined with respect to the direction of normal flow of the air through the turbomachine 1 in operation. Likewise, an axial direction corresponds to the direction of the longitudinal axis X-X, a radial direction refers to a direction which is perpendicular to this longitudinal axis X-X passes through the latter, and a circumferential, or tangential, direction corresponds to the direction of a flat, closed curved line, all points of which are equidistant from the longitudinal axis X-X. Furthermore, and unless otherwise specified, the terms "internal (or interior)" and "external (or exterior)", respectively, are used in reference to a radial direction so that the internal part or surface (i.e. radially internal) of an element is closer to the longitudinal axis X-X than the external (i.e. radially external) part or surface of the same element. Finally, an "annular" element comprises at least a ring portion, that is to say it has a circumferential dimension and a radial dimension, the circumferential dimension being strictly greater than the radial dimension. In a particular embodiment, an element can be annular all around an axis, i.e. it forms a complete, closed ring, centered on the axis.

Injector

With reference to the , the turbomachine 1 comprises an injector 2 arranged:
- downstream of the combustion chamber 14 and,
- upstream of turbine section 16.

The turbine section 16 comprises a disc 160 centered on the longitudinal axis X-X and blades 162 fixedly mounted on an outer end of the disc 160, the blades 162 being subjected to the flow of air exiting the combustion chamber 14 .

Cooling channels (not shown) are formed inside each blade 162. Air is routed into the cooling channels from the compressor section 12. This air passes through an annular bypass space 140 arranged radially inside the combustion chamber 14, then by an annular cavity 164 delimited by an upstream surface of the disc 160 and a downstream surface of a sealing flange 166.

As seen on the , the air passes from the annular bypass space 140 to the annular cavity 164 via the injector 2 mounted fixed relative to the housing 18.

Referring to Figures 3 to 5, in one embodiment, the injector 2 comprises a first annular crown 20 and a second annular crown 22, coaxial with each other and centered on the longitudinal axis X-X.

The first annular crown 20 comprises a first external annular flange 200 and a first internal annular flange 202 interconnected by a first wall 204. A plurality of inlet orifices 2000 are formed in the first external annular flange 200, the orifices of entry 2000 being distributed circumferentially all around the longitudinal axis XX. In addition, the first annular crown 20 comprises a plurality of first fins 208 each projecting from a downstream radial surface of the first wall 204. As visible on the , each first fin 208 has a first intrados 2080 facing an inlet 2000 and a first extrados 2082 facing the longitudinal axis XX, a first chord 2084 of the first fin 208 being inclined with respect to a radial direction . The first chord 2084 of a first fin 208 is a straight line connecting a first leading edge 2086 to a first trailing edge 2088 of the first fin 208.

The second annular crown 22 comprises a second internal annular flange 220 and a second wall 222 extending radially from an external axial surface of the second internal annular flange 220. Furthermore, the second annular crown 22 comprises a plurality of second fins 224 forming each projecting from an upstream radial surface of the second wall 222. As visible on the , each second fin 224 has a second intrados 2240 facing an inlet 2000 once the first annular crown 20 and the second annular crown 22 are assembled on one another, and a second extrados 2242 facing the longitudinal axis XX, a second chord 2244 of the second fin 224 being inclined with respect to a radial direction. The second chord 2244 of a second fin 224 is a straight line connecting a second leading edge 2246 to a second trailing edge 2248 of the second fin 224.

Referring to Figures 4 and 5, once the first annular crown 20 and the second annular crown 22 assembled one on the other so that the downstream axial surface of the first wall 204 is facing the axial surface upstream of the second wall 222, the first fins 208 and the second fins 224 are aligned in a circumferential direction. In this way, channels 24 curved in a tangential direction make it possible to connect the inlet orifices 2000 to the radially internal part of the injector 2. As visible on the , a third annular crown 26 having a third radial wall 260 is attached to the assembly formed by the first annular crown 20 and the second annular crown 24, thus forming the last wall delimiting the channels 24. In other words, the third annular crown 26 closes the channels 24.

In operation, the air coming from the annular bypass space 140 enters through the inlets 2000, is accelerated during its passage through the channels 24 before emerging with a maximized tangential speed into the annular cavity 164. This makes it possible to limit the heating of the disc 160 by maximizing the efficiency of its cooling.

Figures 6 to 8 illustrate other embodiments of the injector 2 capable of being produced by additive manufacturing, and more particularly by laser fusion on a powder bed. These embodiments of the injector 2 have a reduced number of parts. Therefore, injector 2 is faster and cheaper to manufacture. Also, injector 2 is lighter.

With reference to the , in one embodiment, the third annular ring is removed and the first annular ring 20 comprises all the fins 208 projecting from a downstream axial surface of the first wall 204. The second annular ring 22 then makes it possible to close the channels 24 This is why the structure of the second annular crown 22 of this embodiment is simplified compared to the structure of the second annular crown 24 of the embodiment illustrated in FIGS. injector manufacturing cost 2.

Referring to Figures 7 and 8 and 10 to 13, in one embodiment, the second annular crown and the third annular crown are eliminated, the channels 24 being manufactured in one piece with the first external annular flange 200, the first wall 204 and the first internal annular flange 202. The channels 24 thus open at the level of a plurality of outlet orifices 2020 made in the internal annular flange 202 and distributed circumferentially all around the longitudinal axis X-X. Preferably, the inlet orifices 2000 each have a triangular section in a tangential plane, while the outlet orifices 2020 have an oblong section in a tangential plane. In any case, additive manufacturing, and more particularly laser fusion on a powder bed, makes it possible, in this embodiment, to produce various geometries of channels 24 which are not accessible when the injector 2 comprises more than an annular crown 20, 22, 26.

Manufacturing process

With reference to FIGS. 9 to 13, a method E for manufacturing the injector 2 is, according to one mode of implementation, carried out by laser fusion on a powder bed.

This mode of implementation comprises a preliminary step S0 of supplying a plate 3 intended to receive the bed of powder, that is to say the layers of powder successively fused by laser to manufacture the injector 2, and the injector 2 itself as it is manufactured.

The manufacturing axis X-X is chosen normal to the plate 3, that is to say orthogonal to the plane in which the plate 3 extends. The layers of powder successively arranged on the plate 3 also extend in a plane orthogonal to manufacturing axis X-X. Furthermore, as can be seen in FIGS. 10, 12 and 13, once the manufacture of the injector 2 has been completed, the manufacturing axis X-X coincides with the longitudinal axis X-X of the turbomachine 1. In other words, the injector 2 is coaxial with the manufacturing axis X-X.

Method E comprises a step S1 of manufacturing an annular flange 202 coaxial with the manufacturing axis X-X, typically the first internal annular flange 202, as seen in FIGS. 10 to 13. Preferably, the annular flange 202 manufactured during of step S1 is annular all around the manufacturing axis X-X.

Simultaneously with step S1, the method comprises a step S2 of manufacturing a sacrificial support 40, 42 arranged radially with respect to the annular flange 202 manufactured during step S1. In other words, the sacrificial support 40, 42 occupies a space adjacent to the annular flange 202 in a radial direction, the space being able to be radially external to the annular flange 202, as seen in FIGS. 10 to 13, or be radially internal to the annular flange 202.

In one embodiment, a plurality of sacrificial supports 40, 42 can be manufactured during step S2, typically by being distributed circumferentially around the manufacturing axis X-X. The sacrificial supports 40, 42 may be arranged equidistant from each other, or at different distances from each other, in a circumferential direction. Furthermore, each sacrificial support 40, 42 can be arranged at a distance from the X-X manufacturing axis identical to the distance from the X-X manufacturing axis at which another sacrificial support 40, 42 is arranged. Alternatively, the distances of the sacrificial supports 40, 42 to the manufacturing axis X-X can be different from each other. The geometry and the positioning of the sacrificial support(s) 40, 42 are preferably determined according to the manufacturing constraints, typically the time and cost of manufacturing, and/or the structural reinforcement and/or of the heat evacuation that it is sought to confer on the sacrificial support(s) 40, 42 during the manufacture of the injector 2.

Preferably, the sacrificial support 40, 42 is annular and coaxial with the manufacturing axis X-X. For example, during step S2, a plurality of sacrificial supports 40, 42 forming ring portions, distributed circumferentially around the manufacturing axis X-X, at equal distance from each other, or not, can be manufactured. Alternatively, as seen in Figures 10 to 13, the sacrificial support 40, 42 is annular all around the manufacturing axis X-X, and coaxial with the latter. The manufacture of an annular sacrificial support 40, 42, and more particularly annular all around the axis of manufacture X-X, improves the structural reinforcement and the thermal evacuation of the sacrificial support 40, 42 during the manufacture of the injector 2.

In one embodiment illustrated in Figures 10 to 13, the sacrificial support 40, 42 comprises a ferrule 40 which extends away from the annular flange 202. The ferrule 40 constitutes a solid section of the sacrificial support 40, 42 , as opposed to the honeycomb structure 42 described below. Figures 10 to 13 even illustrate two ferrules 40 located at equal radial distance from each other and from the annular flange 202. The manufacture of a ferrule 40 offers increased structural reinforcement to the injector 2 during manufacture. , because the ferrule 40 has a density substantially identical to the density of the injector. In addition, the presence of a ferrule 40 facilitates the manufacture of the wall 204, as will be detailed below, because the ferrule 40 offers increased stability during the deposition of the layer of powder intended to form the wall 204.

In one embodiment, the sacrificial support 40, 42 extends from an axial surface of the annular flange 202 and includes a honeycomb structure 42.

Referring to Figures 10-13, in one embodiment, the sacrificial support 40, 42 includes a ferrule 40 which extends away from the annular flange 202 and a honeycomb structure 42 extending from an axial surface of the annular flange 202, preferably the axial surface of the annular flange open to the radial space, internal or external, where the ferrule 40 extends.

Whatever the mode of implementation, the honeycomb structure 42 has a lower density than the density of the annular flange 202 and/or of the sacrificial support 40, 42. The honeycomb structure 42 facilitates heat evacuation within the injector 2 during manufacture. In addition, the honeycomb structure 42 is easier, less time-consuming and less expensive to manufacture than the ferrule 40. Consequently, in this mode of implementation, the respective function of the ferrule 40, which provides structural reinforcement of the injector 2 during manufacture, and the honeycomb structure 42, which facilitates heat evacuation within the injector 2 during manufacture, are complementary.

Once steps S1 and S2 have been completed, the method comprises a step S3 of manufacturing a wall 204 extending radially from an axial surface of the annular flange 202 and being integral with the annular flange 202 and with the sacrificial support 40, 42. With reference to FIGS. 6 and 10 to 13, this wall 204 corresponds to the first wall 204 of the first crown 20 of the injector 2. Thanks to the presence of the sacrificial support 40, 42 the successive layers of powder forming, in a plane orthogonal to the manufacturing axis X-X, the wall 204 does not deform nor is carried away during the scraping of the powder.

As visible in Figures 10 to 13, in one embodiment, the sacrificial support 40, 42 comprises a first end 401 and a second end 402, the wall 204 being integral with the second end 402. During this implementation, the sacrificial support 40, 42 is manufactured so that a thickness of the sacrificial support 40, 42 is variable between the first end 401 and the second end 402. The thickness constitutes a radial dimension of the sacrificial support 40 , 42. For example, the thickness of the sacrificial support 40, 42 is greater at its second end 402, so as to provide a larger support surface to the layers of powder during their deposition.

In one embodiment of method E, the ferrule 40 is connected to the wall 204 by a junction 43. Preferably, the junction 43 has a weakness in order to facilitate removal of the sacrificial support 40, 42 from the injector 2 once the manufacture of the wall 204 is complete. In this way, the injector 2 can be manufactured in a reduced time. Referring to Figures 12 and 13, In an advantageous variant of this implementation, during step S2, a first end 41 of the sacrificial support, opposite to the junction 43, is crenellated in order to facilitate circulation of the powder towards the outside of the injector 2, marked by arrows in FIGS. 12 and 13. This is particularly advantageous when depowdering the injector 2 subsequent to its manufacture by laser fusion on a powder bed.

With reference to the , in one embodiment, the method E further comprises a step S4 of removing the sacrificial support 40, 42 from the injector 2 once the step S3 has been completed. This is the reason why the support 40, 42 is qualified as “sacrificial”. Advantageously, step S4 is performed by machining and/or by trimming the sacrificial support 40, 42.

Of course, subsequent manufacturing steps to steps S1, S2, and S3 can be provided, and whether these steps are implemented previously or subsequently to step S4. These manufacturing steps by laser melting on a powder bed involve in particular the manufacture of the channels 24, the continuation of the manufacture of the annular flange 202 and/or the manufacture of another annular flange 200, typically the external annular flange 200.

Claims (10)

Method (E) for manufacturing an injector (2) of a turbine (16) of a turbomachine (1) by laser melting on a powder bed, comprising the following steps:
S1: the manufacture of an annular flange (202) coaxial with a manufacturing axis (XX) normal to a plate (3) on which the powder bed is placed,
S2: simultaneously with step S1, the manufacture of a sacrificial support (40, 42) arranged radially with respect to the annular flange (202), then
S3: fabricating a wall (204) extending radially from an axial surface of the annular flange (202), the wall (204) being integral with the annular flange (202) and with the sacrificial support (40 , 42). Method (E) according to claim 1, wherein the sacrificial support (40, 42) is annular and coaxial with the manufacturing axis (X-X). Method (E) according to one of claims 1 and 2, in which the sacrificial support (40, 42) comprises a first end (401) and a second end (402), the wall (204) being integral with the second end (402) and a thickness of the sacrificial support (40, 42) being variable between the first end (401) and the second end (402). Method (E) according to one of Claims 1 to 3, in which the sacrificial support (40, 42) comprises a ferrule (40) which extends at a distance from the annular flange (202), the ferrule (40) being connected to the wall (204) by a junction (43). Method (E) according to claim 4, wherein the junction (43) has a weakness in order to facilitate removal of the sacrificial support (40, 42) from the injector (2) once the manufacture of the wall (204) has been completed. . Method (E) according to one of Claims 4 and 5, in which, during step S2, a first end (401) of the sacrificial support (40, 42), opposite the junction (43), is crenelated in order to facilitate a circulation of the powder towards the outside of the injector (2). Method (E) according to one of claims 4 to 6, wherein the sacrificial support (40, 42) further comprises a honeycomb structure (42) extending from the axial surface of the annular flange (202). Method (E) according to one of claims 1 to 3, wherein the sacrificial support (40, 42) extends from the axial surface of the annular flange (202) and comprises a honeycomb structure (42). Method (E) according to one of claims 1 to 8, further comprising a step S4 of removing the sacrificial support (40, 42) from the injector (2) once step S3 has been completed, preferably by machining and /or by trimming the sacrificial support (40, 42). Injector (2) of turbine (16) of turbomachine (1) manufactured by means of a method (E) according to one of claims 1 to 9.