Inter-turbine casing for a turbine engine, turbine engine assembly, turbine engine and method for positioning a splitter vane within an inter-turbine casing

Abstract

The invention relates to an inter-turbine casing (8) for a turbine engine of the turbine vane frame type, which casing acts as a turbine distributor in such a turbine engine, the inter-turbine casing comprising an inner shroud (9), an outer shroud (10), a plurality of arms (11) extending between the inner shroud and the outer shroud, and at least one set of N splitter vanes (12) positioned circumferentially between two successive arms. Each splitter vane has a mid-height axial chord shorter than the mid-height axial chord of the arms. The two successive arms define a reference position for each of the N splitter vanes. These reference positions are evenly spaced circumferentially between the two successive arms. At least one splitter vane (12-1) of the set of N splitter vanes is circumferentially offset from the reference position thereof. The invention further relates to a turbine engine assembly comprising the inter-turbine casing and a turbine (7) extending downstream of the inter-turbine casing (8) and comprising at least one movable vane (13) extending radially. The invention further relates to a turbine engine comprising the turbine engine assembly. The invention further discloses a method for positioning a splitter vane within an inter-turbine casing.

Description

INTER-TURBINE CASING FOR TURBOMACHINE, TURBOMACHINE ASSEMBLY, TURBOMACHINE AND METHOD FOR POSITIONING A SEPARATOR BLADE WITHIN AN INTER-TURBINE CASING Technical Field

[0001] This disclosure generally relates to an inter-turbine casing for a turbine vane frame (TVF) type turbomachine fulfilling the function of a turbine distributor in such a turbomachine. Prior art

[0002] In a manner known per se, a turbomachine comprises an inter-turbine casing arranged between the high-pressure turbine casing and the low-pressure turbine casing in a turbomachine. The inter-turbine casing comprises a fairing comprising an inner shroud and an outer shroud, which together delimit the flow path between the high-pressure turbine and the low-pressure turbine, as well as arms which extend radially between the inner shroud and the outer shroud.

[0003] The inter-turbine casing has both an aerodynamic function, as a stator allowing the incident flow from the high-pressure turbine to be deflected, a structural function to transmit mechanical forces between the inner shell and the outer shell, and an integration function for the passage of services. The fairing can therefore have a multi-profile configuration and include, in addition to the arms, splitter vanes (or "splitters" in English) which have a smaller chord compared to the arms, which are thicker and have a longer chord. The splitter vanes and the downstream part of the arms thus provide the aerodynamic function, while the arms provide the structural and integration functions.

[0004] However, the wakes generated by the blades of the inter-turbine casing (i.e. the thick arms and the separating vanes) are particularly wide and energetic. In particular, the significant profile differences between the arms and the separating vanes result in significant wake differences that lead to the appearance of significant flow distortions. downstream of the inter-turbine casing, these distortions being likely to disrupt the performance of the low pressure turbine (aerodynamic losses, appearance of separations), to impact its mechanical strength (periodic excitations) or even to pose problems of integration of the low pressure turbine (size of the low pressure turbine).

[0005] There is therefore a real need to remedy the aforementioned drawbacks, by proposing a solution making it possible to improve the aerodynamic efficiency and therefore the performance of the low-pressure turbine in a turbomachine comprising an inter-turbine casing. Statement of the invention

[0006] The present disclosure relates to an inter-turbine casing for a turbomachine, comprising an inner shroud, centered on a central axis, an outer shroud, coaxially surrounding the inner shroud, a plurality of arms, each arm extending between the inner shroud and the outer shroud, having a leading edge and a trailing edge and having an axial chord at mid-height, and at least one set of N separating vanes positioned circumferentially between two successive arms, each separating vane extending between the inner shroud and the outer shroud, having a leading edge and a trailing edge and having an axial chord at mid-height shorter than the axial chord at mid-height of the arms, in which said two successive arms define a reference position for each of said N separating vanes, these reference positions being regularly spaced circumferentially between the two successive arms,and wherein at least one separating blade of said set of N separating blades is circumferentially offset from its reference position.,

[0007] In other words, in the reference configuration, the azimuth Ai of the reference position of the i e separating blade of said set of N separating blades, that is to say its angular position around the central axis, can be written as follows Ai = A B I + (A B 2-ABI) / (N+1 ), with A B I the azimuth of the first arm and A B 2 the azimuth of the second arm delimiting said set of N separating blades.

[0008] Thus, in such a configuration according to the invention, unlike the conventional configuration usually found in rotating machines, the separating blades are not all arranged regularly between the arms.

[0009] Indeed, the inventors found that this reference configuration in which the arms and the separating blades are regularly spaced around the central axis caused significant distortions in the flow rate and orientation of the flow passing through the inter-turbine casing. Conversely, the inventors determined that offsetting all or part of the separating blades relative to their commonly accepted reference positions made it possible to reduce the distortions of the flow, and therefore to increase the performance of the turbine downstream.

[0010] Thus, the invention is the result of technological research aimed at significantly improving the performance of turbomachines and, in this sense, contributes to reducing the environmental impact of the aeronautical sector.

[0011] In some embodiments, the plurality of arms comprises between 4 and 20 arms.

[0012] In some embodiments, the arms are distributed regularly around the central axis. Their distribution is therefore axisymmetric.

[0013] In some embodiments, all arms have the same profile.

[0014] In some embodiments, at least one arm is hollow, said at least one arm comprising a passage allowing the passage of a service of the turbomachine.

[0015] In some embodiments, said set of N separating vanes comprises between 1 and 4 separating vanes.

[0016] In some embodiments, the inter-turbine housing includes a set of separating vanes between each successive arm.

[0017] In some embodiments, each set of splitter vanes includes the same number of splitter vanes.

[0018] In some embodiments, the separating vanes are arranged in the same manner, relative to each other, in each set of separating vanes. Thus, the configuration of each set of separating vanes, i.e., the combination of the respective arrangements of each of their vanes, constitutes a single pattern which is repeated identically between arms.

[0019] In some embodiments, all of the separating vanes have the same profile.

[0020] In some embodiments, the axial chord at mid-height of the separating vanes is at least 2 times shorter, preferably at least 3 times shorter, than the axial chord at mid-height of the arms.

[0021] In some embodiments, the maximum thickness of the separating vanes is less than the maximum thickness of the arms.

[0022] In some embodiments, the maximum thickness of the separating vanes is at least 2 times shorter, preferably at least 3 times shorter, than the maximum thickness of the arms.

[0023] In some embodiments, the profile of at least a downstream portion of the separating blades is identical to the profile of a downstream portion of the arms. The aerodynamic behaviors of the arms and the separating blades are thus similar, at least near their trailing edges, which reduces flow distortions.

[0024] In some embodiments, the trailing edges of the splitter vanes are circumferentially aligned about the central axis with the trailing edges of the arms.

[0025] In some embodiments, the leading edges of the splitter vanes are disposed further downstream than the leading edges of the arms.

[0026] In some embodiments, at least one separating vane of said set of N separating vanes adjacent to an arm is circumferentially offset from its respective reference position. This is preferably the case for the two separating blades of the set adjacent to an arm. This reduces distortions caused by the difference in profile of the arms.

[0027] In some embodiments, all of the separating vanes of said set of N separating vanes are circumferentially offset from their respective reference positions.

[0028] In some embodiments, the circumferential offset of said at least one separating blade that is circumferentially offset from its reference position is less in absolute value than 0.25 x Aref, where Aref is the angular difference between two consecutive reference positions. Preferably, no separating blade has a circumferential offset from its reference position greater, in absolute value, than this ceiling. Indeed, the inventors have determined that the optimal offset zone lies within this range.

[0029] In some embodiments, at least one separating blade of said set of N separating blades has a thickness different from the other separating blades of said set of N separating blades. Consequently, in such a case, the stacking in trace BF (trailing edge) of the cross sections of the blade in question does not overlap with the stacking in trace BF of the cross sections of the arm.

[0030] In certain embodiments, at least one separating blade of said set of N separating blades has a different geometry, in particular by having a different stacking law, from the other separating blades of said set of N separating blades. Similarly, in such a case, the stacking in trace BF (trailing edge) of the cross sections of the blade in question does not overlap with the stacking in trace BF of the cross sections of the arm.

[0031] The present disclosure also relates to a turbomachine assembly, comprising an inter-turbine casing according to any one of the preceding embodiments, and a turbine extending downstream of the inter-turbine casing and comprising at least one radially extending moving blade.

[0032] The present disclosure also relates to a turbomachine, comprising a turbomachine assembly according to any one of the preceding embodiments.

[0033] The present disclosure also relates to a method for positioning a separating blade within an inter-turbine casing according to any one of the preceding embodiments, comprising the following steps: providing an inter-turbine casing, positioning separating blades within the inter-turbine casing in their respective reference positions; evaluating a parameter of the flow passing through the inter-turbine casing on either side of a given separating blade and determining a reference distortion of this parameter; moving said given separating blade into one or more azimuthal positions circumferentially offset relative to its reference position; and for each of these offset azimuthal positions: - evaluate said parameter of the flow on either side of the given separating blade (12-1), - determine the distortion (Dd') of this parameter, and - compare the distortion after displacement (Dd') with the reference distortion (Dd), position the separating blade within the inter-turbine casing at an offset azimuthal position which reduces the distortion of said flow parameter.

[0034] It is understood here that the steps of positioning, displacement and evaluation of the parameter can be carried out experimentally or numerically using a numerical simulation. An offset position is then retained for the given separating blade which makes it possible to reduce, or even minimize, the distortion of the flow parameter.

[0035] In some embodiments, the flow parameter is evaluated as a function of azimuthal position in a downstream radial plane.

[0036] In some embodiments, the step of evaluating the post-displacement flow parameter is performed for several different offset positions of the given splitter vane within a scanning range.

[0037] In some embodiments, the scanning range has an amplitude greater than 0.1 x Aref and less than 0.25 x Aref, where Aref is the angular deviation between two consecutive reference positions.

[0038] In certain embodiments, these steps are repeated successively for each separating blade of said set of N separating blades.

[0039] In some embodiments, the flow parameter being evaluated is the flow rate or the angle of the flow relative to the central axis.

[0040] In this disclosure, the terms "longitudinal", "transverse", "lower", "upper" and their derivatives are defined in relation to the main direction of the blades; the terms "axial", "radial", "tangential", "inner", "outer" and their derivatives are defined in relation to the central axis of the inter-turbine casing, i.e. the main axis of the turbomachine; "axial plane" means a plane passing through the main axis of the turbomachine and "radial plane" means a plane perpendicular to this main axis; the terms "upstream" and "downstream" are defined in relation to the circulation of air in the turbomachine; finally, the terms "front" and "rear" are understood to be circumferentially in the direction when proceeding in a clockwise direction.

[0041] The above-mentioned features and advantages, as well as others, will become apparent from the following detailed description of examples of embodiments of the inter-turbine casing and the proposed method. This detailed description refers to the attached drawings. Brief description of the drawings

[0042] The attached drawings are schematic and are intended primarily to illustrate the principles of the presentation.

[0043] In these drawings, from one figure to another, identical elements (or parts of elements) are identified by the same reference signs.

[0044] [Fig. 1] Figure 1 is an axial sectional plan of a turbomachine according to the invention.

[0045] [Fig. 2] Figure 2 is a sectional view of a turbomachine assembly comprising an example of an inter-turbine casing and a turbine.

[0046] [Fig. 3] Figure 3 is a partial perspective view of the exemplary turbomachine assembly of Figure 2.

[0047] [Fig. 4] Figure 4 is a graph representing the flow rate as a function of azimuthal position for a reference configuration and an offset configuration.

[0048] [Fig. 5] Figure 5 is a graph of flow angle versus azimuthal position for the reference configuration and the offset configuration of Figure 4. Description of the embodiments

[0049] In order to make the description more concrete, an example of an inter-turbine casing is described in detail below, with reference to the attached drawings. It is recalled that the invention is not limited to this example.

[0050] A dual-flow turbomachine 1 (typically, an aircraft turbomachine 1) generally comprises, from upstream to downstream in the direction of gas flow, a fan 2, an annular primary flow vein I and an annular secondary flow vein II. The air mass sucked in by the fan 2 is thus divided into a primary flow, which circulates in the primary flow vein I, and into a secondary flow, which is concentric with the primary flow and circulates in the secondary flow vein II.

[0051] The primary flow stream I passes through a primary body comprising one or more compressor stages, for example a low pressure compressor 3 and a high pressure compressor 4, a combustion chamber 5, one or more turbine stages, for example a high pressure turbine 6 and a low pressure turbine 7 separated by an inter-turbine casing 8, and a gas exhaust nozzle.

[0052] In the present application, upstream and downstream are defined relative to the normal flow direction of the gases in the turbomachine 1. Furthermore, the axis of rotation of the rotor of the low-pressure turbine 7, which coincides with the extension axis of the turbomachine 1, is called the X axis. An axial direction corresponds to the direction of the X axis, a radial direction is a direction perpendicular to this X axis and passing through it. Furthermore, a circumferential direction corresponds to a direction perpendicular to the X axis and not passing through it.

[0053] The inter-turbine casing 8 comprises an inner shroud 9 and an outer shroud 10 substantially coaxial with the axis X, and a fixed blading comprising a plurality of arms 11 together forming a crown. The arms 11 extend from the inner shroud 9 to the outer shroud 10 and may be substantially radial with respect to the axis X. Each arm 11 comprises a structural part, configured to transmit mechanical forces between the inner shroud 9 and the outer shroud 10, which is housed in an aerodynamically shaped wall.

[0054] The aerodynamic wall thus makes it possible to straighten the flow at the outlet of the high-pressure turbine 6 and to improve the supply of the low-pressure turbine 7, which is located immediately downstream of the inter-turbine casing 8.

[0055] The inter-turbine casing 8 may in particular comprise between four and twenty arms 11. In this example, the inter-turbine casing 8 comprises ten arms 11. Preferably, these arms 11 are regularly spaced around the axis X. The inter-turbine casing 8 also comprises a plurality of separating vanes 12 positioned circumferentially between the arms 11: a set of one or more separating vanes 12 is thus provided between each arm 11. It will be noted that, conventionally, the chord of each separating vane 12 is shorter than the chord of each arm 11. The trailing edge 12f of each separating vane 12 is circumferentially aligned with the trailing edge 11f of each arm 11; consequently, the leading edges 11a of the arms 11 are positioned further upstream than the leading edges 12a of the separating blades 12. In addition, preferably, the downstream part of the arms 11 has a profile corresponding to at least a downstream part of the separating blades 12.

[0056] The inter-turbine casing 8 may comprise between one and four separating blades 12 between successive arms 11. Preferably, each set of blades separator 12 comprises the same number of blades 12. In the present example, three separator blades 12 are thus provided between each arm 11.

[0057] The low-pressure turbine 7 comprises, in a manner known per se, a plurality of turbine stages each comprising at least one moving blade 13. The low-pressure turbine 7 may comprise at least three turbine stages, for example between three and five turbine stages in the case of a turbomachine whose fan 2 is driven via a reduction mechanism. Since the low-pressure turbine 7 is conventional, it will not be detailed further here.

[0058] Each arm 11 has a mid-height axial chord 14. The mid-height axial chord 14 of a given arm 11 is defined from a mid-height line 15, which corresponds to the (fictitious) line included in a plane perpendicular to the X axis and which includes all the points located midway between the inner shell 9 and the outer shell 10 and midway between the intrados 11i and the extrados 11e of the arm 11. The mid-height line 15 of a given arm 11 extends axially from the trailing edge of the row of fixed blades of the last stage of the high-pressure turbine 6 immediately upstream of the inter-turbine casing 8 to the leading edge 12a of the moving blade 12 of the low-pressure turbine 7 located immediately downstream of the inter-turbine casing 8.The axial chord at mid-height 13 then corresponds to the length of the straight line segment connecting points A and B, which correspond respectively to the projection on the X axis of point A' located at the intersection between the mid-height line 15 and the leading edge 11 a of the arm 11 , and of point B' located at the intersection between the mid-height line 15 and the trailing edge 11 f of the arm 11 .

[0059] Similarly, each separating blade 12 has a mid-height axial chord measured by projecting onto the X axis the points located at the intersection between the mid-height line 15 and the leading edge 12a of the separating blade 12, and at the intersection between the mid-height line 15 and the trailing edge 12f of the separating blade 12.

[0060] Finally, the arm-blade distance 16 corresponds to the axial distance, measured at mid-height of the arm 11, between the trailing edge 11f of the arm 11 and the leading edge 13a of a moving blade 13 immediately downstream of the arm 11, that is to say to the length of the straight line segment connecting the points B (defined above) and C, where the point C corresponds to the projection on the X axis of the point C' located at the intersection between the mid-height line 15 and the leading edge 13a of the moving blade 13. It will be noted here that the arm-blade distance 16 can be measured for any moving blade 13 of the most upstream stage of the low-pressure turbine 7 (generally designated as the "first stage" of the low-pressure turbine), insofar as the moving blades 13 are symmetrical in revolution around the X axis so that, whatever the moving blade 13 selected in this stage, the projection on the X axis of the point at the intersection between the mid-height line 15 and the leading edge 13a of the moving blade 13 is identical and coincides with point C.

[0061] From then on, we define the downstream plane TVF 17, which is the plane perpendicular to the X axis at 50% of the arm-blade distance 16, that is to say halfway between points B and C.

[0062] An example of a method for positioning the separating vanes 12 will now be described.

[0063] This method aims to adjust the azimuth, i.e. the angular position around the X axis, of each separating blade 12. Preferably, all the sets of separating blades 12 have the same configuration, i.e. the same combination of azimuths for each of their separating blades 12: in other words, the i e separating blade 12 of a set is always located at the same angular distance from the arm 11 preceding it. Therefore, in this example, the method focuses on adjusting the azimuth of the separating blades 12 of a single given set and then the combination of azimuths obtained for this set will be reproduced identically for the other sets.

[0064] That being said, a reference position is first calculated for each separating blade 12 of the set considered. This reference position corresponds to the position that the separating blade 12 would have in the classic configuration of the equal distribution around the X axis, that is to say the reference configuration in which the gap between the separating blades 12 or between a separating blade 12 and its adjacent arm 1 1 is always equal. In the present example, with ten arms 1 1 and three separating blades 12 per arm 1 1 , it can be calculated that this regular gap is equal to 360 ° / (1 Ck (3 + 1 )) = 9 ° . Therefore, if the arm 1 1 preceding the set of separating blades 12 considered in the clockwise direction has azimuth 0°, the reference position of Ia1 ere separating blade 12-1 will be 9°, the reference position of the ? me separating blade 12-2 will be 18°, and the reference position of the 3 eme separating blade 12-3 will be 27°.

[0065] A numerical simulation, or an experiment on a test bench, is then carried out in the reference configuration in order to measure the flow rate of the flow passing through the inter-turbine casing 8 as a function of the azimuthal position in the downstream TVF plane 17, at least along the angular sector separating the two arms 11 framing the set of separating blades 12 considered. A curve 21 is then obtained as represented on the graph in Figure 4 (this graph is represented here in a partial manner, centered on the first blade 12-1 of the set considered since it is the first blade that we will seek to position).

[0066] On this curve 21 , it should be understood that the minimum 21 a corresponds to the reduction of the flow in the extension of the arm 11 , while the minimum 21 b corresponds to the reduction of the flow in the extension of the first blade 12-1 , which makes it possible to identify the latter. It is also noted that the presence of the arm 11 and the blade 12-1 causes flow peaks 21 c and 21 d respectively on their intrados side. In addition to these extrema 21a-21d directly caused by the presence of the arms 11 and the separating blades 12, we note the presence of a local maximum 21 e followed by a local minimum 21 f between the arm 11 and the first blade 12-1 , as well as the presence, in a similar manner, of a local maximum 21 g followed by a local minimum 21 h in front of the first blade 12-1 in the clockwise direction.

[0067] The flow distortion Dd between the front and the rear of the blade 12-1, constituting a reference distortion, is then measured between the maximum flow rate 21 e behind the first blade 12-1 in the clockwise direction and the minimum flow rate 21 h in front of the first blade 12-1 in the clockwise direction.

[0068] The first blade 12-1 is then moved forward and / or backward from its reference position, preferably in a range deviating by a maximum of 0.25 times the normal angular deviation, i.e. 0.25x9° = 225° in the present example. Thus, in the present example, several positions can be tested for the first blade 12-1 between the azimuth 6.75° and 11.25°.

[0069] For each position tested, a numerical simulation or an experiment on a test bench is again carried out in order to measure the flow rate of the flow passing through the inter-turbine casing 8 as a function of the azimuthal position in the downstream plane TVF 17, at least on either side of the first blade 12-1. Curve 22 is then obtained as shown in the graph in Figure 4.

[0070] We then find extrema 22a-22h similar to those of curve 21 of the reference configuration but whose positions are shifted. In particular, we note that the minimum 22b and the maximum 22d are shifted forward in the clockwise direction, which reveals that blade 12-1 has been moved forward in the clockwise direction relative to its reference position.

[0071] In a similar way to what was done for the reference configuration, we then measure in this offset configuration the flow distortion Dd' between the maximum flow 22e behind the first blade 12-1 and the minimum flow 22h in front of the first blade 12-1. We then note that the offset of the blade 12-1 has made it possible to significantly reduce this flow distortion Dd'.

[0072] It is then possible to retain the tested position which makes it possible to minimize the flow distortion Dd'. Then, once the optimal position of the first blade 12-1 has been determined, the same steps can be carried out to determine the optimal position of the second blade 12-2 by focusing on the flow distortion between the rear and the front of the second blade 12-2. And so on for all the separating blades 12 in the set.

[0073] Alternatively, or in addition, it is also possible to carry out these positionings based on another parameter of the flow and in particular the angle of the flow relative to the X axis as measured at the level of the downstream plane TVF 17, as shown in Figure 5.

[0074] Thus, curve 31 of Figure 5 represents the flow angle curve as a function of the azimuthal position in the downstream TVF 17 plane in the reference configuration, while curve 32 represents the flow angle curve in the offset configuration.

[0075] On these curves, minima 31a and 32a allow arm 11 to be located while minima 31b, 32b allow the first blade 12-1 to be located. flow angle distortion Da, Da' is measured between the maximum 31 e, 32 e behind the first blade 12-1 and the maximum 31 h, 32 h in front of the first blade 12-1, in a clockwise direction. It can thus be noted that the offset of the first blade 12-1 has made it possible to significantly reduce the angle distortion Da' compared to the angle distortion Da in the reference configuration.

[0076] In an alternative embodiment, in addition to a circumferential offset of at least one separating blade, it is also possible to modify the thickness of a separating blade. This modification can concern both a circumferentially offset blade and a blade that has not been circumferentially offset.

[0077] In another embodiment, in addition to a circumferential offset of at least one separating blade, it is also possible to modify the stacking law of a separating blade. This modification can concern both a circumferentially offset blade and a blade that has not been circumferentially offset.

[0078] Although the present invention has been described with reference to specific exemplary embodiments, it is obvious that modifications and changes may be made to these examples without departing from the general scope of the invention as defined by the claims. In particular, individual features of the various illustrated / mentioned embodiments may be combined in additional embodiments. Therefore, the description and drawings should be considered in an illustrative rather than restrictive sense.

[0079] It is also obvious that all the characteristics described with reference to a method are transposable, alone or in combination, to a device, and conversely, all the characteristics described with reference to a device are transposable, alone or in combination, to a method.

Claims

Claims

1. Inter-turbine casing for a turbomachine, comprising an inner shroud (9), centered on a central axis (X), an outer shroud (10), coaxially surrounding the inner shroud (9), a plurality of arms (11), each arm (11) extending between the inner shroud (9) and the outer shroud (10), having a leading edge (11a) and a trailing edge (11f) and having a mid-height axial chord (14), and at least one set of N separating vanes (12) positioned circumferentially between two successive arms (11), each separating vane (12) extending between the inner shroud (9) and the outer shroud (10), having a leading edge (12a) and a trailing edge (12f) and having a mid-height axial chord shorter than the mid-height axial chord (14) of the arms (11), in which said two successive arms (11) define a reference position for each of said N separating blades (12),these reference positions being regularly spaced circumferentially between the two successive arms (11), and in which at least one separating blade (12-1) of said set of N separating blades (12) is circumferentially offset relative to its reference position.

2. Inter-turbine casing according to claim 1, in which the arms (11) are distributed regularly around the central axis (X).

3. Inter-turbine casing according to claim 1, in which at least one arm (11) is hollow, said at least one arm (11) comprising a passage allowing the passage of a service of the turbomachine (1).

4. An inter-turbine casing according to any one of claims 1 to 3, comprising a set of separating vanes (12) between each successive arm (11), wherein each set of separating vanes (12) comprises the same number of separating vanes (12), and wherein the separating vanes (12) are arranged in the same manner, relative to each other, in each set of separating vanes (12).

5. Inter-turbine casing according to any one of claims 1 to 4, in which the profile of at least one downstream portion of the separating vanes (12) is identical to the profile of a downstream portion of the arms (11), and in which the trailing edges (12f) of the separating vanes (12) are circumferentially aligned around the central axis (X) with the trailing edges (llf) of the arms (11).

6. An inter-turbine casing according to any one of claims 1 to 5, wherein all of the separating vanes (12) of said set of N separating vanes (12) are circumferentially offset from their respective reference positions.

7. Inter-turbine casing according to any one of claims 1 to 6, wherein the circumferential offset of said at least one separating blade (12-1) which is circumferentially offset from its reference position is less in absolute value than 0.25 x Aref, where Aref is the angular difference between two consecutive reference positions.

8. Inter-turbine casing according to any one of claims 1 to 7, in which at least one separating blade of said set of N separating blades (12) has a thickness different from the other separating blades of said set of N separating blades (12).

9. Inter-turbine casing according to any one of claims 1 to 7, in which at least one separating blade of said set of N separating blades (12) has a different geometry, in particular by having a different stacking law, from the other separating blades of said set of N separating blades (12).

10. A turbomachine assembly, comprising an inter-turbine casing (8) according to any one of the preceding claims, and a turbine (7) extending downstream of the inter-turbine casing (8) and comprising at least one radially extending moving blade (13).

11. An assembly according to claim 10, wherein the turbine (7) comprises a moving wheel, carrying a plurality of moving blades (13), located immediately downstream of the inter-turbine casing (8).

12. A turbomachine, comprising a turbomachine assembly (7-8) according to claim 10 or 11.

13. A method for positioning a separating blade within an inter-turbine casing according to any one of claims 1 to 9, comprising the following steps: providing an inter-turbine casing; positioning separating blades (12) within the inter-turbine casing in their respective reference positions; evaluating a parameter of the flow passing through the inter-turbine casing (8) on either side of a given separating blade (12-1) and determining a reference distortion (Dd) of this parameter; moving said given separating blade (12-1) into one or more azimuthal positions circumferentially offset from its reference position; and for each of these offset azimuthal positions: - evaluate said parameter of the flow on either side of the given separating blade (12-1), - determine the distortion (Dd') of this parameter, and - compare the distortion after displacement (Dd 7) with the reference distortion (Dd), position the separating blade within the inter-turbine casing at an offset azimuthal position which reduces the distortion of said flow parameter.

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