ROTOR BLADE FOR AN AIRCRAFT TURBOMACHINE

Abstract

Rotor blade (20) for an aircraft turbomachine (1), this blade (20) comprising a blade (21) connected by a platform (22) to a foot (23), the blade (20) comprising an internal cavity configured to allow the passage of a cooling airflow, the blade (20) comprising a body (36) made of a nickel-based alloy, characterized in that it comprises at least one insert (32) housed in a recess (33) of the body (31) and made of a ceramic-based material. Figure for the abbreviation: Figure 2

Description

Title of the invention: Rotor blade for an aircraft turbomachine Technical field of the invention

[0001] The present invention relates in particular to a rotor blade for an aircraft turbomachine, as well as a method for manufacturing this blade. Technical background

[0002] Aircraft turbomachine blades, particularly high-pressure turbine blades, are subjected to very high temperatures that can reduce their service life and degrade the turbomachine's performance. Indeed, the turbomachine turbines are arranged downstream of the turbomachine's combustion chamber, which ejects a hot gas stream. This gas is expanded by the turbines, causing them to rotate and thus operating the turbomachine. The high-pressure turbine, which is located directly at the combustion chamber outlet, is subjected to the highest temperatures.

[0003] To enable the turbine blades to withstand these severe thermal stresses, it is known to provide a cooling circuit in which relatively cooler air circulates, drawn from the compressors, which are located upstream of the combustion chamber. More specifically, each turbine blade comprises a blade with an intrados wall and an extrados wall, which are connected upstream by a leading edge and downstream by a trailing edge.

[0004] The cooling circuit generally comprises several cavities inside the blade, some of which are interconnected and supplied with cooling air from the blade root. Part of this cooling air is discharged through outlets located near the trailing edge. These outlets deliver jets of cooling air onto the blade surfaces.

[0005] The cavities of the cooling circuit are generally formed by at least one casting core which is used in a blade manufacturing process employing the lost-wax casting technique. The blade is generally made of a nickel-based metal alloy.

[0006] This process includes a first step of manufacturing one or more foundry cores. A foundry core is conventionally obtained by ceramic injection molding (the ceramic is then debound and sintered). These cores have shapes and dimensions that are predetermined according to the shapes and dimensions of the blade cooling circuit.

[0007] In another step of the process, wax is injected around the nucleus or nuclei. Once the wax has cooled, a model is obtained comprising the nuclei embedded in the wax. The model is arranged on a column with other similar models to form a cluster.

[0008] The process further comprises creating a shell from a refractory material around the cluster, which acts as a mold. The refractory material is, for example, a ceramic. The shell is created by immersing the cluster several times in a ceramic slip.

[0009] The process then consists of removing the wax from the shell by heating the shell and melting the wax. Molten metal is then poured or cast inside the shell to fill the cavities left by the wax removal in the patterns intended to form the metal parts, in this case the turbine blades. Indeed, prior to this metal pouring step, a wax removal step is carried out.

[0010] When the shell is cooled and solidified, a deshelling step destroys the shell and the cores in the metal parts (blade) so as to reveal the final blade and the cooling fluid circulation cavities.

[0011] A blade produced by this technique is formed from a single piece of nickel-based alloy. This material is known for its good mechanical strength within the operating temperature range of a high-pressure turbine (up to approximately 1200°C). More specifically, this material has good tensile and compressive strength that decreases progressively with temperature (from 800°C to 1200°C). This strength allows for use up to a temperature of approximately 1100°C.

[0012] Ceramic materials, for their part, have good resistance, particularly in compression at very high temperatures, up to about 1200°C.

[0013] During operation, the body of a blade undergoes thermal expansion which results in deformations. These deformations generate compression zones and tension zones in the blade.

[0014] There is a need to find a solution to improve the thermomechanical strength of a rotor blade so that it can withstand both tensile and compressive forces at high temperature (up to about 1050°C) and very high temperature (up to about 1200°C). Summary of the invention

[0015] The invention provides a rotor blade for an aircraft turbomachine, this blade comprising a blade connected by a platform to a foot, the blade comprising an internal cavity configured to allow the passage of a cooling airflow, the blade comprising a body made of a nickel-based alloy, characterized in that it includes at least one insert housed in a recess in the body and made of a ceramic-based material, said insert having a shape complementary to that of said recess.

[0016] The invention thus proposes a rotor blade comprising a body made of a nickel-based alloy and one or more inserts in this body made of a ceramic-based material. Each of these materials operates within its preferred operating range. As mentioned above, the nickel-based alloy material has good tensile and compressive strength up to a temperature of 1050°C, i.e., medium to high temperature for an aircraft turbine. The ceramic-based material has good compressive strength up to a temperature of 1200°C, i.e., high temperature for an aircraft turbine.

[0017] It is thus understood that the ceramic-based inserts are located in the areas of the blade likely to be subjected to such temperature levels (up to 1200°C) during operation. Furthermore, these inserts are pre-stressed so that they always operate in compression, regardless of the initial load (tension or compression) of the area.

[0018] The invention therefore makes it possible to improve the thermal resistance performance of the blade, which can thus operate in a hotter environment. It is therefore possible to increase the temperature of the combustion gases in the turbine runner, to increase the temperature of the blade cooling air, and / or to reduce the blade cooling air flow rate, thereby optimizing the performance of the turbine and thus of the turbomachine.

[0019] The rotor blade according to the invention may comprise one or more of the following features, taken individually or in combination with each other:

[0020] — said body exerts a compressive prestress on said insert; this prestress can result from the blade manufacturing process and in particular from a phenomenon of material shrinkage or contraction at the end of the manufacturing process, during the solidification of the blade;

[0021] — said compressive prestress is maintained as long as the temperature the blade is less than a certain value, which corresponds for example to an operating temperature of the blade and the turbomachine;

[0022] — said compressive prestress decreases as the blade temperature increases and can cancel or disappear when the temperature of the dawn reaches the aforementioned value; this is due to the differential thermal expansions of the materials of the body and the insert;

[0023] - the blade includes an insert located at a free end of the blade, opposite the foot;

[0024] - the insert forms at least part of a bathtub at the free end of the blade, at base of which air outlet holes are formed;

[0025] - the awl includes an insert located in the platform;

[0026] - the blade includes an insert located in the blade;

[0027] - the insert is located at the trailing edge of the blade;

[0028] - the ceramic-based material is a ceramic or matrix composite ceramic.

[0029] - the insert has a general parallelepiped or patch shape with a length between 5 and 15mm and a width between 3 and 10mm.

[0030] — the nickel-based alloy is a nickel-based superalloy, possibly intended for single-crystal use;

[0031] — the nickel-based alloy can be part of the superalloy family commonly used for High Pressure Turbine blades: AMI, CMSX4SLS, René N5, PW1484, etc.;

[0032] — the ceramic material is silicon carbide based.

[0033] The present invention also relates to a turbine for an aircraft turbomachine, comprising at least one rotor blade as described above.

[0034] The present invention also relates to a method for manufacturing a rotor blade as described above, in which the method is based on the lost-wax casting principle and comprises the steps of:

[0035] a) manufacturing of the insert,

[0036] b) mounting the insert inside a shell and next to a core,

[0037] c) casting of the nickel-based alloy into the shell and around the core and the insert in such a way as to create the body of the awl,

[0038] d) removal of the core so as to free the blade cavity.

[0039] In this process, it is the solidification of the alloy that generates material shrinkage and can result in a compressive prestress of the body on the insert.

[0040] The present invention further relates to a method for manufacturing a rotor blade as described above, in which the method is based on the lost-wax casting principle and comprises the steps of:

[0041] a) manufacturing of the insert,

[0042] c) casting of the nickel-based alloy into a shell and around a core,

[0043] d) removal of the core so as to free the blade cavity as well as a recess designed to receive the insert, and

[0044] e) mounting the insert in the recess of the body by shrink fitting.

[0045] In this process, after shrink fitting, the parts return to a temperature such that the body can generate a compressive prestress on the insert. Shrink fitting can consist of heating the body and / or cooling the insert, to allow the insert to be mounted in the recess of the body.

[0046] During step b) and / or c), the core (or also the insert when present) can / can be held in place in the shell by means of pins which pass through the shell and bear against the core (or the insert). Brief description of the figures

[0047] The invention will be better understood and other details, features and advantages of the invention will become more apparent upon reading the following description, given by way of non-limiting example and with reference to the accompanying drawings in which:

[0048] [Fig-1] Fig. 1 is a schematic axial and partial cross-sectional view of a aircraft turbomachine;

[0049] [Fig.2] [Fig.2] is a schematic perspective view of a rotor blade

[0050] [Fig.3] [Fig.3] is a schematic view of a free end of the blade of the [Fig.2] and shows hot areas (metal temperature) that are difficult to cool;

[0051] [Fig.4] [Fig.4] is a view similar to that of [Fig.2] and shows other areas with high metal temperature;

[0052] [Fig.5] [Fig.5] is a schematic cross-sectional view of the blade of [Fig.2] and shows its platform and high metal temperature zones;

[0053] [Fig.6] [Fig.6] is a very schematic view of a nickel-based alloy body and a ceramic-based insert during a sintering step; [Fig.7] [Fig.7] is a very schematic view of a nickel-based alloy body and a ceramic-based insert;

[0054] [Fig.8] [Fig.8] is a flowchart showing steps of a process according to the invention for manufacturing a rotor blade, according to two variants; and

[0055] [Fig.9] [Fig.9] is a schematic perspective view of a shell used in one of the variants of the manufacturing process of [Fig.8]. Detailed description of the invention

[0056] Figure 1 shows a partial axial cross-sectional view of a turbomachine 1 with longitudinal axis X to which the invention applies. The turbomachine 1 shown is a twin-spool, twin-flow turbomachine intended for mounting on an aircraft. Of course, the invention is not limited to this type of turbomachine.

[0057] This double-flow turbomachine 1 generally comprises a blower 2 mounted upstream of a gas generator 3.

[0058] In the present invention, and generally, the terms "upstream" and "downstream" are defined with respect to the gas flow in the turbomachine and here along the longitudinal axis X (and even from left to right in [Fig. 1]). The terms "axial" and "axially" are defined with respect to the longitudinal axis X. Similarly, the The terms "radial", "internal" and "external" are defined with respect to a radial axis Z perpendicular to the longitudinal axis X and with respect to the distance from the longitudinal axis X.

[0059] The gas generator 3 comprises, from upstream to downstream, a low pressure compressor 4a, a high pressure compressor 4b, a combustion chamber 5, a high pressure turbine 6a and a low pressure turbine 6b.

[0060] The blower 2, which is surrounded by a blower housing 7 carried by a nacelle 8, divides the air entering the turbomachine into a primary airflow which passes through the gas generator 3 and in particular in a primary channel 9, and into a secondary airflow which circulates around the gas generator in a secondary channel 10.

[0061] The secondary airflow is ejected by a secondary nozzle 11 terminating the nacelle while the primary airflow is ejected outside the turbomachine via an ejection nozzle 12 located downstream of the gas generator 3.

[0062] The high-pressure turbine 6a, like the low-pressure turbine 6b, comprises one or more stages. Each stage includes a stator blade mounted upstream of a moving blade. The stator blade comprises a plurality of stator or fixed blades, called distributors, which are distributed circumferentially around the longitudinal axis X. The moving blade comprises a plurality of moving blades which are also distributed circumferentially around a disk centered on the longitudinal axis X. The distributors deflect and accelerate the aerodynamic flow exiting the combustion chamber towards the moving blades so that the latter are driven into rotation.

[0063] With reference to [Fig. 2], each turbine blade (and here a movable high-pressure turbine blade 20) comprises a blade 21 rising radially from a platform 22. The latter is supported by a foot 23 which is designed to be inserted into one of the corresponding grooves in the turbine disk. Each blade 21 comprises an intrados wall 24 and an extrados wall 25 which are connected upstream by a leading edge 26 and downstream by a trailing edge 27.

[0064] Increased performance requirements and evolving aeronautical regulations are driving current engine manufacturers to design engines operating in increasingly demanding environments (temperature, pressure, rotational speed, emissions, etc.). This necessitates the development of "new generation" high-pressure turbine blades capable of withstanding these types of stresses.

[0065] Increasing the temperature of the gas driving the blade 20 improves the efficiency of the turbomachine 1. This temperature increasingly exceeds, by several hundred degrees, the melting point of the superalloy constituting the blade. The blade must therefore be cooled more and more effectively.

[0066] To this end, the blade 20 includes a cooling circuit 28 arranged inside the blade 21, designed to cool the blade walls 21 subjected to the high temperatures of the primary airflow leaving and passing through the combustion chamber 5. This cooling circuit 28 comprises several cavities (visible in the cross-section of [Fig. 5]) which communicate with each other to form a "trombone"-type conduit. This conduit includes several passages or reversals so that a cooling fluid, in this case cooling air, sweeps across the entire blade 21 from top to bottom along the radial axis.

[0067] The foot 23 includes a supply channel for the circuit 28, which is connected to a cooling air intake system upstream of the combustion chamber 5, such as on the low-pressure compressor 4a. The cooling circuit 28 includes outlet ports 29 arranged in the vicinity of the trailing edge 27, the leading edge 26, and the free end 34 of the blade 21 opposite the foot 23.

[0068] In this way, the RF cooling air which circulates from the base 23 of the blade 20, passes through the cavities inside the blade 21 and opens into the outlet ports 29.

[0069] Fig. 3 shows the free end of the blade 21, which is opposite the foot 23. This free end includes a peripheral rim 30 which extends along the intrados and extrados walls 24, 25, from the leading edge 26 to the trailing edge 27. This free end thus forms a "bathtub" at the bottom of which cooling air outlets 29 are also provided.

[0070] During operation, this tub is subjected to high local temperatures.

[0071] Figures 4 and 5 show that areas of high local temperature ZI can also appear at the trailing edge 27 of the blade 21 and at the platform 22 of the blade 20. The platform 22 is particularly affected by the increase in gas temperatures in the runner. Indeed, with current technologies, its cooling is either nonexistent, inefficient, or very restrictive for design and production. Furthermore, the platform 22 of a blade 20 is typically subjected to strong out-of-phase cyclic loading cycles (hot compression / cold tension) leading to cracks in the surface.

[0072] One solution for withstanding increased wall temperatures is to use ceramic-based materials, which have very high-temperature mechanical properties (>1200°C), unlike nickel-based alloys, which reach their limits. Ceramics have the secondary advantage of being significantly lighter (density < 4000 kg / m³). Conversely, they are potentially more complex to process, and above all, their tensile strength is low across the entire temperature range. For these two reasons, it is not possible to produce an identical blade. to those classically made from nickel-based alloy, using the lost-wax casting process.

[0073] Knowing that the mechanical characteristics of these ceramic-based materials are better in compression (by a factor of approximately 2), the present invention proposes to produce a blade made of two materials: - a body 31 made of nickel-based alloy, which preferably operates in tension at medium to high temperatures (< 1050°C), and - one or more inserts 32 made of ceramic-based material, which preferably operate in compression at very high temperature (> 1200°C); this material may also have better resistance to corrosion and in particular to oxidation.

[0074] The hottest areas can thus include inserts 32 made of ceramic material.

[0075] In the case of [Fig. 3], the bathtub or part of it may be formed by an insert. This insert 32 may, for example, include part or all of the rim 35, and part or all of the outlet holes 29 of the bathtub.

[0076] In the case of [Fig.4], the trailing edge of the blade may include recesses 38 in which ceramic inserts are housed.

[0077] In the case of [Fig.5], the blade platform may also include recesses 38 in which ceramic inserts are housed.

[0078] The insert 32 or each insert preferably has a general parallelepiped shape or a patch with a length between 5 and 15mm and a width between 3 and 10mm. Its thickness can be between 0.3mm and 1.5mm.

[0079] Insert 32 can be replaced, if necessary, during a maintenance operation, by example in case of wear. This replacement may also be planned when the cooling circuit of the blade 20 is redesigned and for example when the outlet ports 29 of the tub have to be positioned differently.

[0080] The blade body 31 is preferably made of nickel-based superalloy, and for example of Inconel®.

[0081] The ceramic-based material is a ceramic or a ceramic matrix composite. For example, silicon carbide.

[0082] The awl 20 according to the invention is preferably made by the lost-wax manufacturing technique which includes steps schematically illustrated in [Fig.8].

[0083] The method for manufacturing the blade according to the invention includes a step of:

[0084] a) manufacturing the insert 32.

[0085] This insert 32 can then be housed in a recess 33 in the body 31 of the blade 20, after the manufacture of the latter, or can be incorporated into the manufacturing shell of the blade 20.

[0086] There are thus two variants of the manufacturing process. According to a first variant, after manufacturing the insert in step a), the process comprises the following steps:

[0087] b) mounting the insert 32 inside a shell and next to a core,

[0088] c) casting of the nickel-based alloy into the shell and around the core and insert 32 so as to produce body 31 of the blade 20, and

[0089] d) removal of the core so as to free the blade cavity.

[0090] In order to keep the insert 32 and the core in position in the shell, retaining pins 34 can pass through the shell and bear against the insert 31 and the core (see [Fig.9]).

[0091] During casting, the molten metal can be at a temperature of approximately 1500°C.

[0092] It is understood that the insert 32 is trapped in the body 31 of dawn 20 after Cooling and solidification of the body. During the cooling phase, the contracting metal traps the ceramic insert 32 and compresses it. This compression ensures the insert's stability. Indeed, ceramics exhibit superior mechanical properties under compression. This results in the configuration shown in [Fig. 7], in which the insert 32 is held tightly by the body 31.

[0093] According to a second embodiment, after manufacturing the insert 32 in step a), the process comprises the steps of:

[0094] c) casting the nickel-based alloy into a shell and around a core,

[0095] d) removal of the core so as to free the blade cavity 20 as well as the recess 3 designed to receive insert 32, and

[0096] e) mounting of the insert 32 in the recess 33 of the body by shrink fitting.

[0097] Shrink fitting can be carried out hot or cold (see [Fig. 6]). In the case of shrink fitting with In the case of cold shrink fitting, the insert 32 is cooled to a very low temperature to contract it and facilitate its assembly in the recess 33 of the body 31. In the case of hot shrink fitting, the body 31 is heated to a very high temperature to expand it and facilitate its assembly in the recess 33 of the body 31.

[0098] Regardless of the manufacturing variant of the blade 20, during operation (when hot), the ceramic insert 32 is always held in place by the metal body 31, but the latter transmits little or no tension to it. The insert 32 is therefore under compression or slight tension. During operation (when cold), the ceramic insert 32 is under compression due to the pressure of the surrounding metal.

[0099] In both embodiments of the process, the insert has a shape complementary to that of the recess. At the end of the process, the body can exert a compressive prestress on the insert, which disappears or can disappear completely in operation due to differential thermal expansion between the two materials of the insert and the body.

[0100] The weak areas of the blade 20 (hottest areas) can thus be made of a suitable material, and the blade 20 can operate in a hotter environment (increased temperature of the combustion gases and / or of the cooling air temperature), where it is possible to reduce the cooling flow of the blade 20, in order to improve the performance of the turbomachine 1.

Claims

Claims

1. Rotor blade (20) for an aircraft turbomachine (1), this blade (20) comprising a blade (21) connected by a platform (22) to a root (23), the blade (20) comprising an internal cavity configured to allow the passage of a cooling air flow, the blade (20) comprising a body (36) made of a nickel-based alloy, and at least one insert (32) housed in a recess (33) of the body (31) and made of a ceramic-based material, said insert having a shape complementary to that of said recess, characterized in that said at least one insert is shrunk into the recess so that said at least one insert is prestressed in the recess.

2. Rotor blade (20) according to claim 1, in which it comprises an insert (32) located at a free end of the blade (21), opposite the root (23).

3. Rotor blade (20) according to claim 2, in which the insert (32) forms at least part of a tub at the free end of the blade (21), at the bottom of which air outlet orifices (29) are formed.

4. Rotor blade (20) according to one of the preceding claims, in which it comprises an insert (32) located in the platform (22).

5. Rotor blade (20) according to one of the preceding claims, in which it comprises an insert (32) located in the blade (21).

6. A rotor blade (20) according to claim 5, wherein the insert (32) is located at a trailing edge (27) of the blade (21).

7. A rotor blade (20) according to any preceding claim, wherein the ceramic-based material is a ceramic or a ceramic matrix composite.

8. Rotor blade (20) according to one of the preceding claims, in which the insert (32) has a generally parallelepiped or patch shape with a length of between 5 and 15mm and a width of between 3 and 10mm.

9. Turbine (6a) for an aircraft turbomachine (1), comprising at least one rotor blade (20) according to one of the preceding claims.

10. A method of manufacturing a rotor blade (20) according to one of claims 1 to 8, wherein the method is based on the principle of lost wax manufacturing and comprises the steps of: a) manufacturing the insert (32), b) casting the nickel-based alloy into a shell and around a core, b) removing the core so as to release the blade cavity as well as a recess (33) intended to receive the insert (32), and d) mounting the insert in the recess (33) of the body (31) by shrinking.

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