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 moving blade, and in particular a blade of a rotor of a turbomachine turbine, in particular in the case where the blades are produced individually and fixed on a disk to form the rotor. Prior art

[0003] Turbomachines generally comprise moving blades which equip fan rotors, compressor rotors or even turbine rotors.

[0004] [Fig. 1] illustrates in particular a juxtaposition of such moving 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 which it equips, denoted C.

[0005] [Fig. 2] illustrates such a moving blade in more detail. Typically, each blade 10 of a turbomachine rotor, and in particular of a turbine or turbomachine compressor rotor, comprises an aerodynamic blade 20 connected by a platform 30 to a root 40, the root 40 making it possible to fix the blade 10 to a rotor disk. The set of platforms 30 forms the internal face of the vein in which the gas flow passes through the blades 20.

[0006] The roots 40 of the moving blades 10 are inserted into the annular disc of the rotor. In particular, the roots 40 of the moving blades are then received in grooves extending in a direction substantially parallel to the axis of rotation A of the rotor, also called cells. The cells are formed at the outer periphery of the disc and distributed regularly around the axis of rotation A of the rotor in the turbomachine. The disc further comprises teeth also formed at the outer periphery of the disc. There is an alternation at the periphery of the disc between the teeth and the cells.

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

[0008] The root 40 of a blade 10 may have a fir tree or dovetail shape. In the case of a dovetail blade root, its bulb comprises a single lobe. In the case of a fir tree blade root 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 collar 45, that is to say a part of smaller section or thickness.

[0009] Generally, the blade 10 generally comprises means for cooling by circulation of a fluid such as air in conduits formed from castings inside the blade 20 and the blade root 40. For this purpose, the root 40 is hollow to allow the blade 20 to be supplied with cooling air.

[0010] Conventionally, blades with fir-tree roots are used for relatively loaded turbines, i.e. for turbines with high speed or large vein section, because there are greater constraints due to the greater centrifugal force. This is particularly the case for high-pressure turbines.

[0011] Maintaining the blades on the rotors and sealing against the gas flow at the platforms are among the problems to be solved when designing these rotors and their blades.

[0012] Indeed, between each blade, clearances are necessary for the assembly of the rotor and to avoid contact during thermal expansion. These clearances cause leaks between the vein and the non-vein, which are harmful to the efficiency of the turbine and therefore to the performance of the turbomachine.

[0013] The objective of the present invention is thus to propose a moving blade making it possible to overcome at least some of the aforementioned drawbacks, and in particular making it possible to reduce leaks 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 substantially forming an angular sector of wall in a so-called axial direction, - blades, each extending from the platform in a so-called radial span direction substantially perpendicular to the axial direction, and - a single foot intended to cooperate with a rotor disc for fixing the blade, the single foot extending from the platform in 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 vanes is at least divided by two in the case of two-bladed vanes, reducing by a factor of two the leaks between two successive vanes, known as inter-vane leaks for a rotor, compared to a conventional rotor equipped with a single-bladed vane.

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

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

[0018] Although the root is solid, the number of attachments is reduced by at least two in the case of two-bladed blades, making it possible to reduce the mass of the rotor and consequently of the turbine which it equips and therefore of the associated turbomachine. The mass of the rotor thus minimized makes it possible to significantly reduce the overall mass balance of a turbine or a turbomachine which includes it.

[0019] The rotor blade for a turbine according to the invention may comprise one or more of the following characteristics, taken in isolation from one another or in combination with one another in all technically possible combinations: - the blade comprises at most two blades, a first blade and a second blade; - the foot is substantially centered in 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 their junction with the platform in a circumferential direction, the single foot having a dimension in the circumferential direction substantially equal to said given distance; - the foot extends continuously in 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 moving 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 previously.

[0021] Another subject 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 previously.

[0022] Another subject 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 previously. Brief description of the drawings

[0023] The present invention will be better understood and other details, characteristics and advantages of the present invention will appear more clearly on reading the description of a non-limiting example which follows, with reference to the appended drawings in which: - [Fig.l], already described, is a three-dimensional schematic view of a set of moving blades juxtaposed to each other to form a rotor according to the prior art; - [Fig.2], already described, is a three-dimensional schematic view of a moving blade of the assembly of [Fig.l] according to the prior art; - [Fig.3] schematically represents an axial sectional view of an example of a turbomachine to which the invention applies; - [Fig.4] illustrates a three-dimensional schematic view of a moving blade according to the invention in which the leading edges of the blades of the blade 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; - [Fig.7] illustrates a turbomachine rotor equipped with the moving blades of figures 4 to 6; and - [Fig.8] schematically represents an enlarged view of the periphery of an annular rotor disc configured to support blades according to the invention.

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

[0025] Referring firstly to [Fig. 3], there is shown an aircraft turbomachine 50 to which the invention applies. This is a double-flow, double-spool turbojet. 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 fan 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 in general, the terms "upstream" and "downstream" are defined in relation to a general direction of circulation of the fluids inside the turbomachine, and here along the longitudinal axis C, that is to say from left to right with reference to [Fig.3].

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

[0029] In the description, the expressions "internal" or "interior" and "external" or "exterior" are used without limitation in reference to the radial distance from the longitudinal axis C around which the turbomachine extends, the expression "internal" defining a zone radially closer to the longitudinal axis of the nacelle, as opposed to the expression "external". Furthermore, in the description and the claims, the terminology axial, radial and transverse will be adopted without limitation in reference to the trihedron A, R, T indicated 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 bladed wheels, each rotor corresponding to a stage of the turbine. A so-called staged turbine is a turbine comprising several stages. Between each pair of consecutive rotors, the turbine comprises stator blades forming distributors.

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

[0033] Figures 4 to 6 represent, in different orientations, a moving 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 towards the inside, blades 120 constituting the aerodynamic part of the blade, a platform 130 and a single blade root 140. By single blade root 140, it is meant that the blade 110 comprises a root common to the blades 120 of the blade unlike the mobile blade 10 of the prior art illustrated in [Fig.2].

[0035] According to a preferred embodiment of the invention, the moving blade 110 comprises at most two blades as shown in FIGS. 4 to 6: a first blade 120A and a second blade 120B. Hereinafter, such a moving 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 extrados or extrados wall 122 having a generally convex outer surface and an intrados or intrados wall 123 having a generally concave inner surface. The extrados 122 and intrados 123 walls are spaced from each other on either side of a center line of the blade profile. The extrados 122 and intrados 123 walls 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 edges BA and trailing edges 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 directions A and radial directions R, between the first blade 120A and the second blade 120B so as to balance the constraints to retain the blades under centrifugal forces.

[0040] The root 140 of the moving 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 root 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 example illustrated, the bulb 142 comprises two lobes. Each lobe 144 is connected to another lobe or to the Péchasse 143 of the foot by a neck 145, that is to say a part of smaller section or thickness.

[0042] The stilt 143 usually has a small thickness in 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 their junction with the platform 130 in the circumferential direction T. The single root 140 has a dimension in the circumferential direction substantially equal to said given distance D.

[0044] More precisely, the bulb 142 of the root extends circumferentially, that is to say in 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 close to the extrados wall 122A of the first blade 120A and the second end 147 of the bulb is arranged close to the intrados wall 123B of the second blade 120B. Thus, the distance in 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 in the circumferential direction T between the first end 146 and the second end 147, is sized to withstand additional stresses compared to a single-bladed vane of the prior art and as illustrated in [Fig. 2] due to the offset of the root 140 relative to the blades 120A and 120B. According to one example, the circumferential dimension The differential of the bulb of the root of a two-bladed blade is at least double that of the root of a single-bladed blade.

[0046] The plane P may 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 root 140 is solid. In other words, it is devoid of air passage channels or orifices, in particular orifices provided in the bottom 148 of the root, that is to say the radially internal surface of the root.

[0048] The root 140 is configured to be inserted into an annular disc of the rotor to allow the blades to be fixed to the disc. An example of such a disc 200 is illustrated in [Fig. 8]. In particular, the blade root 140 is shaped to be received in a groove 202 of a shape complementary to the root and extending in a direction substantially parallel to the axis of rotation of the rotor. The grooves 202 of a rotor disc 200 are also called cells. In a known manner, the cells are formed at the outer periphery of the disc and distributed regularly around the axis of rotation of the rotor in the turbomachine. The disc 202 further comprises teeth 204 also formed at the outer periphery of the disc to cooperate with the blade roots. There is an alternation at the periphery of the disc 200 between the teeth 204 and the cells 202.

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

[0050] Furthermore, in 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 in 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 in the circumferential direction T between the first edge 131 and the second edge 132. The foot 140 is substantially centered in 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 constraints to retain the blades under centrifugal forces.

[0053] The stilt 143 usually has a small thickness in a circumferential direction T while the platform 130 also extends on either side of the Péchasse 143 in this same circumferential direction T.

[0054] The platform 130 has a radially outer surface 135 which delimits ra- dialy inwards the main vein 60a.

[0055] With reference to Figures 5 and 6, the platform 130 comprises two opposite 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 non-limiting figures 4 to 6, the blade 110 further comprises two stiffeners 161, 162 respectively associated with each of the two spoilers 136, 137. Each stiffener 161, 162 extending radially inwards from its associated spoiler, over a circumferential length substantially identical to that of the spoiler, and over a small axial thickness compared to the circumferential length. Each stiffener 161, 162 is therefore also located projecting on either side of the Péchasse 143 in the circumferential direction T, at the axial ends of this stilt.

[0057] Furthermore, the foot extends continuously in 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 face and the downstream face of the foot are spaced apart from each other by a substantially constant distance. The upstream face and the downstream face of the foot are preferably parallel to each other.

[0058] Preferably, the blade 110 is a single, single-piece piece from a foundry.

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

[0060] The blade further advantageously comprises a system for cooling the blade, 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 comprises through orifices 180, an inlet 182 of which is provided 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 precisely, the platform 130 comprises at least one first through orifice 180A opening into the internal cavity of the first blade 120A and at least one second through orifice 180B opening into the internal cavity of the second blade 120B.

[0062] In a known manner, each blade 120 has through holes and slots through which the cooling air is evacuated. For information purposes, the leading edge BA is curved 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 shape tapered, and it has a series of cooling slots 174. The intrados wall 123, which is subject to significant heating in operation, also has a series of through holes 176 distributed at several locations on this intrados.

[0063] [Fig. 7] represents a turbomachine rotor 190, and in particular a turbine rotor, equipped with the moving blades according to the invention, and in particular the two-bladed blades in the illustrated example, and as described previously. In [Fig. 7], the disk has not been 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 with each other 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 which can generate leaks.

[0064] On the contrary, for a “conventional” rotor of the prior art, such as that shown in [Fig. 1], comprising the same number of blades, i.e. 60, there are 60 feet and therefore 60 interfaces with the rotor disc and 60 interfaces between the blades which can generate leaks, i.e. double that of the invention.

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

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

[0067] Although the root is solid, the number of attachments is reduced by at least two in the case of two-bladed blades, making it possible to reduce the mass of the rotor and consequently of the turbine which it equips and therefore of the associated turbomachine. The mass of the rotor thus minimized makes it possible to significantly reduce the overall mass balance of a turbine or a turbomachine which includes it.

[0068] It is 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

Claims

1. Rotor blade (110) for a turbine, in particular an aircraft turbomachine, comprising: - a platform (130) substantially forming an angular wall sector in a so-called axial direction (A), - blades (120), each extending from the platform (130) in a so-called radial span direction (R) substantially perpendicular to the axial direction, and - a single root (140) intended to cooperate with a rotor disk for fixing the blade, the single root extending from the platform (130) in a direction opposite to the radial direction (R), the single root is in the shape of a solid fir tree.

2. A 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, in which the root (140) is substantially centered in a circumferential direction (T), perpendicular to the axial (A) and radial (R) directions, between the first blade (120A) and the second blade (120B).

4. A 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 their junction with the platform (130) in a circumferential direction (T), the single root (140) having a dimension in the circumferential direction substantially equal to said given distance (D).

5. A blade according to any preceding claim, wherein the root (140) extends continuously in the axial direction between an upstream face (163) and a downstream face (164), the upstream face and the downstream face of the root being parallel to each other.

6. A blade according to any preceding claim, wherein each of the blades is hollow and has an internal cavity and the platform has an inner surface comprising through-holes, each associated with a blade, each hole opening into the internal cavity of the associated blade.

7. A blade according to any one of the preceding claims, wherein the moving blade is intended to equip a high pressure turbine.

8. Turbine rotor for a turbomachine, in particular for an aircraft turbomachine, comprising blades according to any one of the claims-

9. preceding indications. Staged turbine for a turbomachine, in particular for an aircraft turbomachine, comprising at least one turbine rotor according to the claim

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

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