TURBOMACHINE ROTOR BLADE AND ASSOCIATED TURBOMACHINE ROTOR

Abstract

The invention relates to a rotor blade (110) for a turbine, particularly an aircraft turbomachine, comprising: - a platform (130) forming substantially an angular sector of wall along an axial direction (A), - blades (120), each extending from the platform (130) along a radial span direction (R) substantially perpendicular to the axial direction, and - a single foot (140) intended to cooperate with a rotor disk for attaching the blade, the single foot extending from the platform (130) in a direction opposite to the radial direction (R), the single foot being solid in the shape of a fir tree. Figure for the abstract: Figure 4

Description

Title of the invention: Blade for a turbomachine rotor and associated turbomachine rotor technical field

[0001] The technical field of the invention is that of blades and in particular aircraft turbomachine blades.

[0002] The present invention relates to a movable blade, and in particular a blade of a rotor of a turbomachine turbine, especially in the case where the blades are made individually and fixed on a disk to form the rotor. Previous technique

[0003] Turbomachinery generally includes moving blades that equip fan rotors, compressor rotors or turbine rotors.

[0004] Figure 1 illustrates in particular a juxtaposition of such movable blades for the purpose of forming such a rotor or moving turbine wheel of a turbomachine. Such a rotor has a substantially annular shape around an axis of rotation corresponding to the longitudinal axis of the turbomachine it equips, denoted C.

[0005] Figure 2 illustrates such a movable blade in more detail. Typically, each turbine rotor blade 10, and in particular that of a turbine or compressor turbine rotor, comprises an aerodynamic blade 20 connected by a platform 30 to a foot 40, the foot 40 enabling the blade 10 to be fixed to a rotor disc. The assembly of platforms 30 forms the inner face of the duct through which the gas flow passes over the blades 20.

[0006] The feet 40 of the movable blades 10 are inserted into the annular disk of the rotor. In particular, the feet 40 of the movable blades are received in grooves extending in a direction substantially parallel to the axis of rotation A of the rotor, also called recesses. The recesses are formed on the outer periphery of the disk and distributed regularly around the axis of rotation A of the rotor in the turbomachine. The disk further has teeth also formed on its outer periphery. There is an alternation between the teeth and the recesses at the periphery of the disk.

[0007] The foot 40 of a blade comprises two parts: a radially internal part called the bulb 42 and a radially external part called the stilt 43. The bulb 42 of the foot is connected by its stilt 43 to the blade platform 10.

[0008] The foot 40 of a blade 10 can be fir-shaped or dovetail-shaped. In the case of a dovetail-shaped blade foot, its bulb comprises a single lobe. In the case of a fir-shaped blade foot 40 as illustrated in [Fig. 2], its bulb 42 comprises several lobes 44, generally two or three. Each lobe 44 is connected to another lobe or at Péchasse 43 of the foot by a col 45, that is to say a part of smaller section or thickness.

[0009] Generally, the blade 10 includes cooling means for circulating a fluid such as air through cast conduits inside the blade 20 and the blade foot 40. For this purpose, the foot 40 is hollow to allow the blade 20 to be supplied with cooling air.

[0010] Traditionally, blades with fir wood feet are used for relatively loaded turbines, i.e., for high-speed turbines or turbines with a large flow cross-section, because there are greater stresses due to the higher centrifugal force. This is particularly the case for high-pressure turbines.

[0011] The retention of the blades on the rotors and the sealing against the gas flow at the level of the platforms are among the problems to be solved during the design of these rotors and their blades.

[0012] Indeed, clearances are necessary between each blade for rotor assembly and to prevent contact during thermal expansion. These clearances cause leakage between the turbine runner and the non-runner area, which is detrimental to turbine efficiency and therefore to the performance of the turbomachine.

[0013] The objective of the present invention is thus to provide a movable blade that overcomes at least some of the aforementioned drawbacks, and in particular reduces leakage between two successive blades in order to improve the performance of the turbine equipped with such blades. Summary of the invention

[0014] To this end, the invention relates to a rotor blade for a turbine, in particular for an aircraft turbomachine, comprising: - a platform forming essentially an angular sector of wall along a so-called axial direction, - blades, each extending from the platform along a radial span direction substantially perpendicular to the axial direction, and - a single foot intended to cooperate with a rotor disc for blade attachment, the single foot extending from the platform along a direction opposite to the radial direction, the single foot is in the shape of a solid fir tree.

[0015] The invention thus makes it possible to reduce leaks at the rotor level since for a given number of blades, the number of blades is at least divided by two in the case of two-bladed blades reducing by a factor of two the leaks between two successive blades, known as inter-blade leaks, for a rotor compared to a conventional rotor equipped with a single-bladed blade.

[0016] In addition, the number of interfaces with the rotor disk is also reduced.

[0017] In addition, a solid foot improves the mechanical strength of the blade, allowing it to resist more effectively the stresses due to centrifugal force.

[0018] Although the base is solid, the number of attachments is reduced by at least half in the case of two-bladed blades, thus reducing the mass of the rotor and consequently of the turbine it equips, and therefore of the associated turbomachine. The minimized rotor mass significantly reduces the overall mass balance of a turbine or turbomachine that incorporates it.

[0019] The turbine rotor blade according to the invention may comprise one or more of the following features, taken individually or in combination with each other according to all technically possible combinations: c - the blade comprises at most two blades, a first blade and a second blade; - the foot is substantially centered along a circumferential direction, perpendicular to the axial and radial directions, between the first blade and the second blade; - the first blade and the second blade are spaced apart from each other by a given distance at the level of their junction with the platform along a circumferential direction, the single foot having a dimension along the circumferential direction substantially equal to said given distance; - the foot extends continuously along the axial direction between an upstream face and a downstream face, the upstream face and the downstream face of the foot being parallel to each other; - each of the blades is hollow and has an internal cavity and the platform has an internal surface comprising through holes, each associated with a blade, each hole opening into the internal cavity of the associated blade; - the movable blade is intended to equip a high-pressure turbine.

[0020] The invention also relates to a turbine rotor for a turbomachine, in particular for an aircraft turbomachine, comprising blades according to the invention and as described above.

[0021] Another object of the invention relates to a staged turbine for a turbomachine, in particular for an aircraft turbomachine, comprising at least one turbine rotor according to the invention and as described above.

[0022] Another object of the invention relates to a turbomachine, in particular an aircraft turbomachine, comprising at least one staged turbine according to the invention and as described above. Brief description of the drawings

[0023] The present invention will be better understood and other details, features and advantages of the present invention will become more apparent upon reading the description of a non-limiting example that follows, with reference to the accompanying drawings in which: - the [Fig.l], already described, is a three-dimensional schematic view of a set of movable blades juxtaposed to each other to form a rotor according to the prior art; - the [Fig.2], already described, is a three-dimensional schematic view of a moving blade of the whole of the [Fig.1] according to the prior art; - [Fig.3] schematically represents an axial cross-sectional view of an example of a turbomachine to which the invention applies; - [Fig.4] illustrates a three-dimensional schematic view of a movable blade according to the invention in which the leading edges of the blade blades are in the foreground; - [Fig.5] is a three-dimensional schematic view of the moving blade of [Fig.4] in which the trailing edges of the blades of the blade are in the foreground; - [Fig.6] is a three-dimensional schematic view of the moving blade of [Fig.4] in which the lower surface of the blade platform is visible; - Figure 7 illustrates a turbomachine rotor equipped with the movable blades shown in Figures 4 to 6; and - [Fig.8] schematically represents an enlarged view of the periphery of an annular rotor disk configured to support blades according to the invention.

[0024] Elements having the same functions in the different implementations have the same references in the figures. Description of the implementation methods

[0025] With reference first to [Fig. 3], an aircraft turbomachine 50 is shown, to which the invention applies. This is a twin-spool, turbofan engine. However, it could be a turbomachine of another type, for example a turboprop, without departing from the scope of the invention.

[0026] The turbomachine 50 has a longitudinal axis denoted C around which its various components extend. It comprises, from upstream to downstream, a blower 52, a low-pressure compressor 54, a high-pressure compressor 55, a combustion chamber 56, a high-pressure turbine 57 and a low-pressure turbine 58.

[0027] In the present invention, and generally, the terms "upstream" and "downstream" are defined with respect to a general direction of fluid flow inside the turbomachine, and here along the longitudinal axis C, i.e. from left to right with reference to [Fig.3].

[0028] Conventionally, after passing through the blower 52, the air F splits into a central primary flow P and a secondary flow S that surrounds the primary flow. The primary flow P flows into a main gas circulation channel 60a passing through the compressors 54, 55, the combustion chamber 56, and the turbines 57, 58. The flow secondary S flows into a secondary vein 60b delimited radially outwards by an engine casing, surrounded by a nacelle 62.

[0029] In the description, the terms "internal" or "inside" and "external" or "outside" are used by way of non-limiting agreement with reference to the radial distance from the longitudinal axis C around which the turbomachine extends, the term "internal" defining an area radially closer to the longitudinal axis of the nacelle, as opposed to the term "external". Furthermore, in the description and the claims, the terminology axial, radial, and transverse shall be adopted by way of non-limiting agreement with reference to the trihedral axis A, R, T shown in the figures, the axial axis A being parallel to the longitudinal axis of the turbomachine.

[0030] The longitudinal axis C is the axis of rotation of the moving elements of the turbomachine 50, and in particular, of the turbines 57, 58.

[0031] Each turbine 57, 58, and in particular the high-pressure turbine 57, comprises one or more rotors or rotating blades, each rotor corresponding to a stage of the turbine. A turbine is called a staged turbine. Between each pair of consecutive rotors, the turbine includes stator blades forming distributors.

[0032] Each rotor is formed of an annular disk extending circumferentially around an axis, coinciding with the longitudinal axis C of the turbomachine, on the periphery of which movable blades are fixed. The movable blades are regularly distributed around the disk.

[0033] Figures 4 to 6 represent, in different orientations, a movable blade according to the invention intended to equip any one of the turbines 57 and 58, in particular the high-pressure turbine 57.

[0034] According to the invention, the blade 110 comprises, in a radial direction denoted R, going from the outside to the inside, blades 120 constituting the aerodynamic part of the blade, a platform 130 and a single blade foot 140. By single blade foot 140, it is understood that the blade 110 comprises a foot common to the blades 120 of the blade unlike the movable blade 10 of the prior art illustrated in [Fig.2].

[0035] According to a preferred embodiment of the invention, the movable blade 110 comprises at most two blades as shown in Figures 4 to 6: a first blade 120A and a second blade 120B. Hereafter, such a movable blade with two blades will be called a two-bladed blade.

[0036] Each blade 120 extends from the platform 130 in a span direction to a free end 121, also called the apex. The span direction is perpendicular to the axial axis A and substantially parallel to the radial axis R.

[0037] Each blade 120 comprises an upper surface (extrados) or upper surface (extrados) 122 having a generally convex outer surface and an lower surface (intrados) or lower surface (intrados) 123 having a generally concave inner surface. The upper surfaces (extrados) 122 and lower surfaces (intrados) 123 are spaced apart on either side of a median line in the blade profile. The upper surfaces (extrados) 122 and lower surfaces (intrados) 123 are connected at their upstream ends by a leading edge (BA) and at their downstream ends by a trailing edge (BF) of the gases flowing into the turbine.

[0038] The extrados 122 and intrados 123 walls as well as the leading edge BA and trailing edge BF have curved shapes which extend substantially parallel to the span direction of the blade.

[0039] The foot 140 is substantially centered along the circumferential direction T, perpendicular to the axial direction A and radial direction R, between the first blade 120A and the second blade 120B so as to balance the stresses to retain the blades under centrifugal forces.

[0040] The foot 140 of the movable blade 110 according to the invention comprises a radially internal part called bulb 142 and a radially external part called stilt 143. The bulb 142 of the foot is connected by its stilt 143 to the platform 130 of the blade 110.

[0041] The foot 140 has a fir tree shape. Thus, its bulb 142 comprises several lobes 144, generally two or three. In the illustrated example, the bulb 142 comprises two lobes. Each lobe 144 is connected to another lobe or to the stem 143 by a collar 145, that is to say, a part of smaller cross-section or thickness.

[0042] The stilt 143 usually has a low thickness along the circumferential direction T.

[0043] The first blade 120A and the second blade 120B are spaced apart from each other by a given distance D at the level of their junction with the platform 130 along the circumferential direction T. The single foot 140 has a dimension along the circumferential direction substantially equal to said given distance D.

[0044] More precisely, the bulb 142 of the foot extends circumferentially, that is, along the direction T, on either side of a plane P between a first end 146 and a second end 147. The first end 146 of the bulb is arranged near the upper surface 122A of the first blade 120A and the second end 147 of the bulb is arranged near the lower surface 123B of the second blade 120B. Thus, the distance along the circumferential direction T between the first end 146 and the second end 147 is substantially equal to the given distance D between the two blades.

[0045] The circumferential dimension of the bulb, i.e. the distance along the circumferential direction T between the first end 146 and the second end 147, is dimensioned to withstand additional stresses compared to a blade single-bladed of the prior art and as illustrated in [Fig.2] due to the offset of the foot 140 relative to the blades 120A and 120B. According to one example, the circumferential dimension of the bulb of the foot of a two-bladed blade is at least double that of the foot of a single-bladed blade.

[0046] The plane P can be a plane of symmetry of the foot, such as in the example illustrated in figures 4 to 6.

[0047] According to the invention, the blade foot 140 is solid. In other words, it is devoid of air passage channels or orifices, in particular of orifices provided in the bottom 148 of the foot, that is to say the radially internal surface of the foot.

[0048] The foot 140 is configured to be inserted into an annular disk of the rotor to allow the blades to be attached to the disk. An example of such a disk 200 is shown in [Fig. 8]. In particular, the blade foot 140 is shaped to be received in a groove 202 complementary in shape to the foot and extending in a direction substantially parallel to the axis of rotation of the rotor. The grooves 202 of a rotor disk 200 are also called recesses. As is known, the recesses are formed on the outer periphery of the disk and are regularly distributed around the axis of rotation of the rotor in the turbomachine. The disk 202 further has teeth 204 also formed on the outer periphery of the disk to cooperate with the blade feet. There is an alternation on the periphery of the disk 200 between the teeth 204 and the recesses 202.

[0049] The platform 130 extends transversely, that is, in a direction substantially perpendicular to the radial axis R. It has a substantially rectangular shape around the blades of the blade. More precisely, the platform 130 extends circumferentially, that is, along the direction T, between a first edge 131 arranged beyond the lower surface 123B of the second blade 120B and a second edge 132 arranged beyond the upper surface 122A of the first blade 120A.

[0050] In addition, along the axial direction A, the platform 130 further extends between an upstream edge 133 arranged upstream of the leading edges BA of the blades 120A, 120B and a downstream edge 134 arranged downstream of the trailing edges BF of the blades 120A, 120B.

[0051] The platform 130 has an axial dimension, that is to say the distance along the axial direction A between the upstream edge 133 and the downstream edge 134.

[0052] The platform 130 has a circumferential dimension, that is to say the distance along the circumferential direction T between the first edge 131 and the second edge 132. The foot 140 is substantially centered along the circumferential direction T between the first edge 131 and the second edge 132 of the platform so as to be substantially centered between the first blade 120A and the second blade 120B so as to balance the stresses to retain the blades under centrifugal forces.

[0053] The stilt 143 usually has a low thickness along a circumferential direction T while the platform 130 also extends on both sides of the stilt 143 along this same circumferential direction T.

[0054] The platform 130 has a radially external surface 135 which radially delimits the main vein 60a inwards.

[0055] With reference to Figures 5 and 6, the platform 130 comprises two opposing axial ends, each forming a spoiler: an upstream spoiler 136 located between the upstream edge 133 and the leading edges BA of the blades 120 and a downstream spoiler 137 arranged between the downstream edge 134 and the trailing edges BF of the blades 120.

[0056] In the example illustrated in Figures 4 to 6, which are not limiting, the blade 110 further comprises two stiffeners 161, 162 respectively associated with each of the two spoilers 136, 137. Each stiffener 161, 162 extends radially inward from its associated spoiler, over a circumferential length substantially identical to that of the spoiler, and over an axial thickness small compared to the circumferential length. Each stiffener 161, 162 is therefore also located in projection on either side of the strut 143 along the circumferential direction T, at the axial ends of this strut.

[0057] Furthermore, the foot extends continuously along the axial direction A, between an upstream face 163 and a downstream face 164. The first stiffener 161 extends in the same plane as the upstream face 163 of the foot, while the second stiffener 162 extends in the same plane as the downstream face 164 of the foot. The upstream and downstream faces of the foot are spaced apart by a substantially constant distance. The upstream and downstream faces of the foot are preferably parallel to each other.

[0058] Preferably, the blade 110 is a single, one-piece casting.

[0059] Preferably, each blade 120 is hollow and has an internal cavity 170 delimited by the extrados wall 122 and intrados wall 123 and the leading edge BA and trailing edge BF of the blade.

[0060] The blade further advantageously comprises a blade cooling system, in particular by circulating cooling air in the internal cavity 170. This cooling air is taken from one of the compressors 54, 55 of the turbomachine. Internal ducts through which circulates cooling air taken from one of the compressors 54, 55 of the turbomachine.

[0061] The platform 130 has through-holes 180, one of which has an inlet 182 on a radially inner surface 138 opposite the radially outer surface 134, and an outlet opening into the internal cavity 170 of the blade. More specifically, the platform 130 has at least one first through-hole 180A opening into the internal cavity of the first blade 120A and at least one second through-hole 180B opening into the internal cavity of the second blade 120B.

[0062] As is known, each blade 120 has through holes and slots through which cooling air is expelled. By way of example, the leading edge BA is convex and provided with a series of cooling holes 172 passing through the wall of each blade in this area. The trailing edge BF of the blades has a tapered shape and includes a series of cooling slots 174. The lower surface 123, which is subjected to significant heating during operation, also has a series of through holes 176 distributed at several locations on this lower surface.

[0063] Figure 7 represents a turbomachine rotor 190, and in particular a turbine rotor, equipped with the movable blades according to the invention, and in particular two-bladed blades in the illustrated example, as described above. In Figure 7, the disk is not shown in order to clearly visualize the assembly / juxtaposition of the two-bladed blades 110 according to the invention and the reduced number of feet compared to the number of blades. Thus, in this example, 30 two-bladed blades are juxtaposed for a total of 60 blades. There are therefore only 30 feet 140 in this rotor and consequently only 30 interfaces with the rotor disk and 30 interfaces between blades that could generate leaks.

[0064] On the contrary, for a "classic" rotor of the prior art, such as that shown in [Fig.1], having the same number of blades, i.e. 60, there are 60 feet and therefore 60 interfaces with the rotor disk and 60 interfaces between the blades that can generate leaks, i.e. twice that of the invention.

[0065] The invention as described above thus makes it possible to reduce leaks at the rotor level since for a given number of blades, the number of blades is at least divided by two in the case of two-bladed blades reducing by at least a factor of two the leaks between two successive blades, known as inter-blade leaks, for a rotor compared to a conventional rotor equipped with a single-bladed blade.

[0066] In addition, a solid foot improves the mechanical strength of the blade, allowing it to resist more effectively the stresses due to centrifugal force.

[0067] Although the base is solid, the number of attachments is reduced by at least half in the case of two-bladed blades, thus reducing the mass of the rotor and consequently of the turbine it equips, and therefore of the associated turbomachine. The minimized rotor mass significantly reduces the overall mass balance of a turbine or turbomachine that incorporates it.

[0068] It should be noted that the examples illustrated in the figures are in no way limiting; the invention can also be applied to moving blades of low-pressure turbines or to turbines for different types of engines.

Claims

Demands

1. A rotor blade (110) for a turbine, in particular an aircraft turbomachine, comprising: - a platform (130) forming substantially an angular sector of wall along an axial direction (A), - blades (120), each extending from the platform (130) along a radial span direction (R) substantially perpendicular to the axial direction, and - a single foot (140) intended to cooperate with a rotor disk for the attachment of the blade, the single foot extending from the platform (130) along a direction opposite to the radial direction (R), the single foot is in the form of a solid fir tree, in which each of the blades is hollow and has an internal cavity and the platform has an internal surface comprising through holes, each associated with a blade, each hole opening into the internal cavity of the associated blade.

2. Blade according to claim 1, comprising at most two blades, a first blade (120A) and a second blade (120B).

3. Blade according to claim 2, wherein the foot (140) is substantially centered along a circumferential direction (T), perpendicular to the axial (A) and radial (R) directions, between the first blade (120A) and the second blade (120B).

4. Blade according to claim 2 or 3, wherein the first blade (120A) and the second blade (120B) are spaced apart from each other by a given distance (D) at the level of their junction with the platform (130) along a circumferential direction (T), the single foot (140) having a dimension along the circumferential direction substantially equal to said given distance (D).

5. Blade according to any one of the preceding claims, wherein the foot (140) extends continuously along the axial direction between an upstream face (163) and a downstream face (164), the upstream face and the downstream face of the foot being parallel to each other.

6. Blade according to any one of the preceding claims, wherein the movable blade is intended to equip a high-pressure turbine.

7. 11 Turbine rotor for turbomachine, in particular for aircraft turbomachine, comprising blades according to any one of the preceding claims.

8. Staged turbine for turbomachine, in particular for aircraft turbomachine, comprising at least one turbine rotor according to the preceding claim.

9. Turbomachine, in particular aircraft turbomachine, comprising at least one staged turbine according to the preceding claim.

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