Inner shroud segment of aircraft turbine engine

Abstract

An inner shroud segment (1) for mounting in an aircraft turbine engine (50), the inner shroud being a rotating member rotating about a longitudinal axis (X) extending from upstream to downstream, the inner shroud segment (1) comprising a body (2) comprising an outer surface (Sext) internally defining a secondary flow channel (41), the body (2) comprising an upstream end (3) connected to a hub of an intermediate case of the aircraft turbine engine (50), the invention relates to an inner shroud section (1) for an intermediate casing, comprising a body (2) made of a composite material comprising a plurality of reinforcing fibers located in a matrix, the inner shroud section (1) comprising at least one hollow column (13) mounted at an upstream end (3) of the body (2), each hollow column (13) extending radially and configured to receive a connecting element for connecting the inner shroud section (1) to a hub of the intermediate casing.

Description

Technical Field

[0001] The present invention relates to the field of aircraft turbine engine casings, and more particularly to an inner shroud segment. Background Art

[0002] Known, reference Figure 1 An aircraft turbine engine 500, having a longitudinal axis X, includes, from upstream to downstream, a fan 200, a low-pressure compressor 220, a high-pressure compressor 230, a combustion chamber 240, a high-pressure turbine 250, and a low-pressure turbine 260. The compressors 220, 230, the combustion chamber 240, and the turbines 250, 260 collectively define a primary flow channel 400 for air circulation, which is externally bounded by a central casing 270. The aircraft turbine engine 500 also includes a secondary flow channel 410 for air circulation, which extends around the exterior of the primary flow channel 400 and is surrounded by the nacelle 300.

[0003] Known and Figure 1 As shown, an aircraft turbine engine 500 includes an intermediate casing 190, which includes an inner shroud 100 and an outer shroud 180, which respectively define the interior and exterior of a secondary flow channel 410. The inner shroud 100 and the outer shroud 180 are connected by a support arm 170, which extends radially in the secondary flow channel 410 downstream of a row of outlet guide vanes (OGVs) 330. The outer shroud 180 extends in the downstream extension of the fan casing 210 and in the upstream extension of the outer fan duct 320. The inner shroud 100 extends in the upstream extension of the inner fan duct 310. The inner fan duct 310 and the central casing 270 together define a core module 280 extending radially between the primary flow channel 400 and the secondary flow channel 410.

[0004] As is known, the inner shroud 100 is connected upstream to the intermediate casing hub 160. In practice, the inner shroud 100 is composed of a plurality of circumferentially distributed inner shroud segments, usually in the form of a one-piece titanium alloy component. Each inner shroud segment is connected to the hub 160 by a set of fastening screws, which are inserted into through holes in the inner shroud segment, which are formed by drilling through a thickened part of the titanium alloy. The high mechanical strength of the titanium alloy enables the inner shroud segments to efficiently transmit forces to the intermediate casing. Due to the high resistance of the titanium alloy to thermal deformation, the inner shroud segments also form a fire barrier. However, the disadvantage of such inner shroud segments is their high mass. This leads to an increase in the weight of the aircraft turbine engine 500 and therefore increases its energy consumption in flight.

[0005] The use of composite materials in one or more parts of an aircraft turbine engine case is known from application WO 2010007220 A2. Applications FR 3 108 679 A1, FR 2 992 353 A1 and FR 3 115 832 A1 are also known. Summary of the Invention

[0006] The present invention aims to reduce the mass of the inner shroud segment while maintaining its fire barrier properties and its ability to transmit mechanical forces to the intermediate casing.

[0007] The present invention relates to an inner shroud segment installed in an aircraft turbine engine. The inner shroud is a rotating member that rotates about a longitudinal axis extending from upstream to downstream. The inner shroud segment comprises a body having an outer surface that internally delimits a secondary flow channel of the aircraft turbine engine. The body comprises an upstream end connected to a hub of an intermediate casing of the aircraft turbine engine.

[0008] The present invention is notable in that:

[0009] - the body is made of a composite material comprising a plurality of reinforcing fibers in a matrix, and

[0010] The inner shroud segment comprises at least one hollow post mounted at the upstream end of the body, each hollow post extending radially and configured to house a connecting element for connecting the inner shroud segment to the hub of the intermediate casing.

[0011] The use of composite materials reduces the mass of the inner shroud segments of the present invention. The use of composite materials is unprecedented for an inner shroud that must exhibit excellent fire resistance and, due to its numerous joints, is composed of multiple assembled inner shroud segments. Furthermore, the hollow columns of the present invention are inserts, which advantageously avoids the formation of additional thickenings in the main body, further contributing to weight reduction. This also allows the composite materials to be easily and evenly compressed during the main body manufacturing process.

[0012] According to one aspect of the invention, the upstream end of the body comprises at least one first through hole, and the hollow column is crimped into a corresponding first through hole. Advantageously, the connection obtained by this mounting method is able to resist any deformation of the composite material under the action of heat.

[0013] According to one aspect of the present invention, the upstream end of the main body comprises a first longitudinal wall and a second longitudinal wall radially spaced apart, the first longitudinal wall being provided with at least one first through-hole, the second longitudinal wall being provided with at least one second through-hole, and the hollow column connecting corresponding first through-holes with corresponding second through-holes. This geometric structure does not require any localized thickening and is advantageously suitable for main bodies formed of composite materials.

[0014] According to one aspect of the present invention, the upstream end of the main body includes a third wall connecting the first longitudinal wall and the second longitudinal wall, wherein the third wall is separated from each hollow column by a free volume. There is no filling material between the hollow column and the third wall, thereby reducing weight.

[0015] According to one aspect of the present invention, the main body comprises an inner surface opposite to the outer surface, and each first through hole opens into the inner surface. Advantageously, the crimping is performed on a side of the inner surface that is most likely to be subjected to mechanical and thermal deformation.

[0016] According to one aspect of the present invention, the upstream end of the main body includes a first inner longitudinal wall and a second outer longitudinal wall arranged radially apart, the first longitudinal wall is provided with at least one first through hole, the second longitudinal wall is provided with at least one second through hole, a hollow column connects the corresponding first through hole with the corresponding second through hole, and the upstream end of the main body includes a third wall, which connects the first longitudinal wall and the second longitudinal wall downstream of the at least one hollow column. This geometric structure does not require any local thickening portion and is advantageously suitable for a main body formed of a composite material. This geometric structure is particularly suitable for clamping a component with a larger thickness while ensuring good mechanical force transmission. The so-called larger thickness refers to the average thickness of the upstream end main body being greater than 1 cm, preferably greater than 1.5 cm.

[0017] According to one aspect of the present invention, the hollow columns are crimped into corresponding first through-holes. Advantageously, the connection achieved through this installation method is resistant to any deformation of the composite material caused by heat. The crimping is advantageously performed on the side of the inner surface most likely to be affected by mechanical and thermal deformation. The crimping also helps protect the turbine engine in the event of a fire by preventing heat from spreading. Compared to installation methods that use a clearance fit on the outside, or crimp or glue the hollow columns into the second through-holes in the second wall, this crimping method prevents flames from spreading through the first through-holes.

[0018] According to one aspect of the present invention, there is a gap between the hollow column and the corresponding second through hole. The existence of the gap can prevent the hollow column from being statically indeterminately installed on the main body.

[0019] According to one aspect of the present invention, the radial spacing between the first longitudinal wall and the second longitudinal wall is greater than four times the thickness of the first longitudinal wall, preferably greater than eight times the thickness of the first longitudinal wall. The spacing in question corresponds to the height of the hollow column and is preferably greater than 1 cm, more preferably greater than 1.5 cm. The inner shroud segment has a relatively large thickness, and the hollow column of the present invention can achieve clamping of this relatively thick portion, so that mechanical force can be efficiently transmitted while reducing the mass of the turbine engine. In the absence of the hollow column, in order to ensure the clamping of this relatively thick portion, a large number of composite material layers would be required, which is not feasible.

[0020] According to one aspect of the present invention, the third wall and the at least one hollow column are separated by a free volume. There is no filling material between the hollow column and the third wall, thereby reducing weight.

[0021] According to one aspect of the present invention, the first longitudinal wall, the second longitudinal wall, and the third wall comprise multiple layers of composite material secured together, at least one of which is referred to as a fire-resistant layer, which extends continuously through the first longitudinal wall and the third wall. In a preferred embodiment, the fire-resistant layer defines the entire inner surface of the body. Thus, the continuity of the composite material layers creates a fire barrier on the side of the inner surface that is most susceptible to mechanical and thermal deformation.

[0022] According to one aspect of the present invention, at least one hollow column is made of metal, preferably stainless steel. Stainless steel is defined as steel with a carbon content of less than 1.2% and a chromium content of more than 10.5%. Advantageously, stainless steel has sufficient deformability to enable crimping.

[0023] The invention also relates to an intermediate casing for an aircraft turbine engine, comprising a hub and at least one inner shroud segment as described above, the intermediate casing extending along a longitudinal axis X and comprising at least one connecting element connecting the upstream end of the inner shroud segment to the hub, each connecting element extending into the hollow column of the inner shroud segment.

[0024] The invention also relates to an aircraft turbine engine comprising a secondary flow channel and an intermediate casing, the intermediate casing comprising a hub and at least one inner shroud segment as described above, the aircraft turbine engine extending along a longitudinal axis X of the inner shroud segment, the outer surface of the inner shroud segment internally delimiting the secondary flow channel, the aircraft turbine engine comprising at least one connecting element connecting the upstream end of the inner shroud segment to the hub of the intermediate casing, each connecting element extending into a hollow column of the inner shroud segment. BRIEF DESCRIPTION OF THE DRAWINGS

[0025] The invention will be better understood on reading the following description, given by way of example, and with reference to the following drawings, given by way of non-limiting example, in which like references indicate similar objects.

[0026] Figure 1 It is a longitudinal half-section schematic diagram of an aircraft turbine engine in the prior art.

[0027] Figure 2 It is a longitudinal half-section schematic diagram of an aircraft turbine engine equipped with an intermediate casing with a set of inner shroud segments according to one embodiment of the present invention.

[0028] Figure 3 yes Figure 2 A schematic perspective view of the intermediate casing is shown.

[0029] Figure 4 FIG. 1 is a schematic perspective view of an inner shroud segment according to an embodiment of the present invention as viewed from an upstream direction.

[0030] Figure 5An embodiment of the present invention Figure 4 A longitudinal cross-section showing the inner shroud segment connected to the hub of the intermediate casing.

[0031] Figure 6 is Figure 4 Schematic diagram of the operation of installing a hollow column on the main body of the inner shroud section.

[0032] It should be noted that the accompanying drawings illustrate the invention in detail in order to implement the invention and that said drawings can of course serve to better define the invention if necessary. DETAILED DESCRIPTION

[0033] refer to Figure 2 As previously mentioned, the aircraft turbine engine 50 generally extends along a longitudinal axis X extending from upstream to downstream. Hereinafter, the terms "upstream" and "downstream" are defined relative to the orientation of the longitudinal axis X. The terms "inner" and "outer" are defined relative to a radial axis extending perpendicular to the longitudinal axis X.

[0034] Still refer to Figure 2 An aircraft turbine engine 50 typically includes, from upstream to downstream, a fan 20, a low-pressure compressor 22, a high-pressure compressor 23, a combustor 24, a high-pressure turbine 25, and a low-pressure turbine 26. The compressors 22, 23, the combustor 24, and the turbines 25, 26 collectively define a primary flow passage 40 for air circulation, which is externally bounded by an intermediate casing 27. The aircraft turbine engine 50 typically also includes a secondary flow passage 41 for air circulation, which extends externally around the primary flow passage 40 and is surrounded by a nacelle 30. This type of aircraft turbine engine 50 is known as a ducted engine.

[0035] refer to Figure 2 and Figure 3 The aircraft turbine engine 50 comprises an intermediate casing 19 comprising an inner shroud 15 and an outer shroud 18 with a longitudinal axis X, said inner shroud and outer shroud respectively delimiting a secondary flow channel 41 externally and internally. The intermediate casing 19 also comprises a hub 16 to which the inner shroud 15 is connected upstream. The hub 16 extends internally along the longitudinal axis X relative to the inner shroud 15. Figure 3 In the example shown, the hub 16 externally delimits a primary flow channel 40. In this example, the intermediate casing 19 also comprises arms 17 connecting the inner shroud 15 and the outer shroud 18, said arms 17 extending radially in a secondary flow channel 41 downstream of a row of outlet guide vanes (OGVs) 33.

[0036] refer to Figure 2 and Figure 3The outer shroud 18 extends in the downstream extension of the fan housing 21 and in the upstream extension of the outer fan duct 32. The inner shroud 15 is arranged opposite the outer shroud 18 and extends in the upstream extension of the inner fan duct 31. The inner fan duct 31 and the center casing 27 generally define a core compartment 28 that extends radially between the primary flow channel 40 and the secondary flow channel 41.

[0037] like Figure 3 As shown, the inner shroud 15 is a rotating member that rotates about a longitudinal axis X in an aircraft turbine engine 50 , which in the following text refers indiscriminately to the axis of the turbine engine 50 and the axis of the inner shroud 15 , these two axes being coincident. Figure 3 As shown, the inner shroud 15 is assembled from a set of inner shroud segments 1 distributed circumferentially around the longitudinal axis X. In this example, the inner shroud segments 1 have a variable angular width.

[0038] According to the present invention and Figure 4 and Figure 5 As shown, each inner shroud segment 1 comprises:

[0039] - a body 2 made of a composite material comprising a plurality of reinforcing fibers in a matrix; and

[0040] One or more hollow posts 13 mounted on the upstream end 3 of the body 2 , each hollow post 13 extending radially and configured to house a connecting element 14 for connecting the inner shroud segment 1 to the hub 16 of the intermediate casing 19 .

[0041] Due to the use of composite materials, the inner shroud segment 1 of the present invention advantageously reduces mass compared to titanium alloy inner shroud segments according to the prior art. The use of composite materials in the inner shroud segment 1 is unprecedented and contradicts current usage, such as for acoustic panels connected to the fairing. Due to its location close to the main flow path in the engine, the inner shroud must be able to ensure mechanical strength in the event of a fire, which hinders the use of composite materials. In addition, the inner shroud is composed of several inner shroud segments connected by numerous connection points, some of which have a large clamping thickness, which also hinders the use of composite materials. As described below, the inner shroud segment 1 of the present invention forms a fire barrier and transmits mechanical forces to the intermediate casing.

[0042] like Figure 3 and Figure 4 As shown, the body 2 of the inner shroud segment 1 includes an outer surface Sext curved toward the longitudinal axis X and an inner surface Sint opposite the outer surface Sext. The outer surface Sext defines a secondary flow channel 41 internally. The inner surface Sint is configured to extend radially relative to the hub 16 at the upstream end 3 when the intermediate casing 19 is installed ( Figure 5 ).

[0043] like Figure 3 and Figure 4 As shown, the main body 2 includes an upstream end 3, on which one or more hollow posts 13 are disposed for securing the intermediate casing 19 to the hub 16. The number of hollow posts 13 depends on the angular width of the inner shroud segment 1, and preferably there are at least two. The hollow posts 13 are preferably distributed across the angular width of the upstream end 3. The following description of a single hollow post 13 and its attachment to the main body 2 applies to each hollow post 13.

[0044] according to Figure 3 and Figure 4 In the preferred aspect shown, the upstream end 3 comprises a first longitudinal wall 6 and a second longitudinal wall 4 arranged radially apart, and for each hollow column 13, the first longitudinal wall 6 and the second longitudinal wall 4 respectively comprise a first through hole 8 and a second through hole 7. The second longitudinal wall 4 preferably extends in the upstream extension of the body 2. The first longitudinal wall 6 extends radially inwardly of the second longitudinal wall 4. Figure 4 As shown, the radial distance H between the second longitudinal wall 4 and the first longitudinal wall 6 corresponds to the height of the hollow pillar 13 and is preferably four times, and more preferably eight times, the thickness E of the first longitudinal wall 6 or the second longitudinal wall 4. The radial distance H is preferably greater than 1 cm, and more preferably greater than 1.5 cm. The first through-hole 8 and the second through-hole 7 corresponding to the same hollow pillar 13 are radially aligned along the radial axis Y, and the hollow pillar 13 extends along the radial axis Y in the through-holes 7 and 8.

[0045] according to Figure 3 and Figure 4 In the preferred aspect shown, the second longitudinal wall 4 is connected to the first longitudinal wall 6 by a third wall 5. The third wall 5 extends downstream of the longitudinal walls 4 and 6. In this example, the third wall 5 extends radially and forms a U-shape with the longitudinal walls 4 and 6, with the concave side facing upstream. Here, the longitudinal walls 4 and 6 extend at the upstream end 3 of the main body 2 and, together with the third wall 5, define a channel with a U-shaped cross-section and an upstream opening. Alternatively, in a plane defined by the longitudinal axis X and the radial axis, the third wall 5 extends in a direction forming an angle of 10° to 90° with the longitudinal axis X. In this alternative, the third wall 5 preferably extends in such a manner that the longitudinal length of the first longitudinal wall 6 is less than the longitudinal length of the second longitudinal wall 4.

[0046] As previously mentioned, the body 2 is made of a composite material comprising reinforcing fibers embedded in a matrix. The body 2 is preferably a one-piece structure. The reinforcing fibers, for example, are made of carbon to impart good mechanical strength to the body 2. The matrix is formed by heating and compressing a thermosetting material (such as a polyepoxide known as an epoxy resin). Such materials have a ductile structure that hardens to their final shape after heating and pressurizing.

[0047] like Figure 5 As shown, the body 2 comprises a composite combination of material layers Ca, Cb, Cc fixed together by heating and pressurizing to form a one-piece structure.

[0048] The main body 2 preferably comprises:

[0049] An inner layer Ca which delimits the entire inner surface Sint of the body 2 and extends continuously in particular in the first longitudinal wall 6 and in the third wall 5 of the upstream end 3 .

[0050] An outer layer Cc which delimits the entire outer surface Sext of the body 2 and extends continuously in particular in the second longitudinal wall 4 of the upstream end 3 .

[0051] an intermediate layer Cb which extends only in the upstream end 3 , radially between the outer layer Cc and the inner layer Ca, ie extends continuously in the first longitudinal wall 4 , the third wall 5 and the second longitudinal wall 6 .

[0052] In this example, the cross section of the inner layer Ca is generally "S"-shaped, the cross section of the outer layer Cc is generally linear, and the cross section of the middle layer Cb is generally "U"-shaped.

[0053] The continuous provision of the inner layer Ca, referred to as the fire-resistant layer, advantageously enables a fire barrier to be formed between the primary channel 40 and the secondary channel 41. The fire-resistant layer Ca helps to contain any fire and prevent the spread of heat. As will be seen later, due to the special crimping method used to install the hollow column 13, the fire barrier effect is also maintained at the first through-hole 8.

[0054] In this example, the body 2 comprises only three layers Ca, Cb, Cc. Alternatively, the body may further comprise other layers sandwiched between the layers Ca, Cb, Cc, preferably arranged in such a way as to maintain a substantially constant thickness.

[0055] like Figure 5 and Figure 6 As shown, each hollow column 13 is an insert in the form of a through-hole connecting hole, which is radially arranged between the first through hole 8 and the second through hole 7 of the upstream end 3 along the radial axis Y. The hollow column 13 has a radial opening 13b for accommodating a connecting element 14 such as a screw. Figure 6 As shown, the hollow column 13 further includes a first end 13 c that cooperates with the first through hole 7 and a second end 13 a that cooperates with the second through hole 8 .

[0056] According to a preferred aspect, the first end 13c is crimped into the first through-hole 8, i.e., inserted into the first through-hole 8 by deformation of the material. This connection forms a fire barrier at the first through-hole 8, acting as an extension of the fire-resistant layer Ca. Compared to conventional installation methods, particularly those with gaps, crimping prevents the spread of fire through the first through-hole 8. The crimping means that even if the composite material of the body 2 deforms slightly under the influence of heat, the seal (the seal between the hollow column 13 and the fire-resistant layer Ca) is maintained at each first through-hole 8.

[0057] Preferably, if Figure 5 and Figure 6 As shown, the first end 13c and the first through-hole 8 include mutually cooperating frustoconical portions that increase in size from the outside to the inside along the radial axis Y. The hollow post 13 is inserted from the outside inward, and the frustoconical portion of the first end 13c is then deformed to mate with the first through-hole 8 (i.e., it is pressed against the first longitudinal wall 6). Also preferably, the first end 13c includes an abutment portion that is configured to abut radially from the outside against the first longitudinal wall 6. The diameter of the abutment portion is greater than the diameter of the first through-hole 8. The frustoconical portion of the end 13c and the abutment portion together provide a radial barrier to the outward and inward translation of the hollow post 13, respectively.

[0058] Alternatively, the first end 13c is connected to the first through hole 8 by gluing or riveting.

[0059] According to a preferred aspect, the second end 13a extends freely in the second through hole 7, that is, it is installed in the second through hole 7 in a clearance fit manner. This avoids an over-static installation of the hollow column 13 on the body 2 and avoids mechanical stress that may shorten the service life. Figure 6 As shown, the cross section of the second through hole 7 is larger than the outer cross section of the hollow column 13 , so that the hollow column 13 can be radially inserted through the end 13 a to be mounted on the main body 2 .

[0060] Hollow column 13 is preferably made of metal to ensure good mechanical strength, preferably stainless steel. Advantageously, stainless steel is deformable, making it easier to press-fit, unlike alloys such as titanium. Stainless steel is defined as steel with a carbon content of less than 1.2% and a chromium content of greater than 10.5%. Hollow column 13 is preferably a one-piece structure.

[0061] according to Figure 4 and Figure 5In the preferred aspect shown, the hollow column 13 is separated from the third wall 5 by a free volume V. Due to the press-fit installation, no filler material is required between the hollow column 13 and the third wall 5. Besides facilitating the choice of composite materials, this also helps reduce the weight of the inner shroud segment 1. Preferably, the radial width of the U-shaped channel defined by the three walls 4, 5, and 6 is at least equal to its longitudinal depth, preferably at least 1 cm, and more preferably at least 1.5 cm. For this geometry, the use of filler or reinforcement materials is impractical and ineffective.

[0062] The hollow column 13 mounted on the main body 2 advantageously simplifies the geometric structure of the main body 2, in particular maintains a constant thickness, thereby allowing the main body 2 to be made of composite materials. In addition, the hollow column 13 avoids forming a thickened portion in the third wall 5 to form a mortise by drilling as in the prior art.

[0063] like Figure 6 As shown, the hollow post 13 is radially inserted into the second through-hole 7 and then connected to the first through-hole 8 by means of crimping to form the inner shroud segment 1. The inner shroud segments 1, which together define the inner shroud 15, are then installed in the aircraft turbine engine 50 via the connecting element 14. The connecting element 14 is radially inserted from the outside into the hollow post 13 of the intermediate casing 19 and the hub 16 to connect the inner shroud segment 1 to the hub 16.

Claims

1. An inner shroud segment (1) to be installed in an aircraft turbine engine (50), the inner shroud being a rotating member defined relative to a longitudinal axis (X) extending from upstream to downstream, the inner shroud segment (1) comprising a body (2), the body (2) comprising an outer surface (Sext) internally delimiting a secondary flow channel (41) of the aircraft turbine engine (50), the body (2) comprising an upstream end (3) connected to a hub (16) of an intermediate casing (19) of the aircraft turbine engine (50), the body (2) being made of a composite material, the inner shroud segment (1) being characterized by: - the composite material of the body (2) comprises a plurality of reinforcing fibers in a matrix, and - the inner shroud segment (1) comprises at least one hollow post (13) mounted at the upstream end (3) of the body (2), each hollow post (13) extending radially and configured to receive a connecting element (14) for connecting the inner shroud segment (1) to the hub (16) of the intermediate casing (19).

2. The inner shroud segment (1) according to claim 1, characterized in that The upstream end (3) of the main body (2) comprises a first inner longitudinal wall (6) and a second outer longitudinal wall (4) arranged radially spaced apart, the first longitudinal wall (6) being provided with at least one first through hole (8), the second longitudinal wall (4) being provided with at least one second through hole (7), a hollow column (13) connecting the corresponding first through hole (8) and the corresponding second through hole (7), and the upstream end (3) of the main body (2) comprises a third wall (5), the third wall connecting the first longitudinal wall (6) and the second longitudinal wall (4) downstream of the at least one hollow column (13).

3. The inner shroud segment (1) according to claim 2, characterized in that The hollow column (13) is pressed into the corresponding first through hole (8).

4. The inner shroud segment (1) according to claim 2 or 3, characterized in that There is a gap between the hollow column (13) and the corresponding second through hole (7).

5. The inner shroud segment (1) according to any one of claims 2 to 4, characterized in that The radial spacing (H) between the first longitudinal wall (6) and the second longitudinal wall (4) is greater than four times the thickness (E) of the first longitudinal wall (6), preferably greater than eight times the thickness (E) of the first longitudinal wall (6).

6. The inner shroud segment (1) according to one of claims 2 to 5, characterized in that The third wall (5) is separated from the at least one hollow column (13) by a free volume (V).

7. The inner shroud segment (1) according to one of claims 2 to 6, characterized in that The first longitudinal wall (6), the second longitudinal wall (4) and the third wall (5) comprise a plurality of composite material layers (Ca, Cb, Cc) fixed together, at least one of which is a fire-resistant layer (Ca), which extends continuously in the first longitudinal wall (6) and the third wall (5).

8. The inner shroud segment (1) according to claim 7, characterized in that The refractory layer (Ca) defines the entire inner surface (Sint) of the body (2).

9. The inner shroud segment (1) according to one of claims 1 to 8, characterized in that The at least one hollow column (13) is made of metal, preferably stainless steel.

10. An aircraft turbine engine (50) comprising a secondary flow channel (41) and an intermediate casing (19), said intermediate casing comprising a hub (16) and at least one inner shroud segment (1) according to one of claims 1 to 9, characterized in that The aircraft turbine engine (50) extends along the longitudinal axis X of the inner shroud, the outer surface (Sext) of the inner shroud segment (1) delimiting an internal secondary flow channel (41), the aircraft turbine engine (50) comprising at least one connecting element (14) for connecting the upstream end (3) of the inner shroud segment (1) to the hub (16) of the intermediate casing (19), each connecting element (14) extending in a hollow column (13) of the inner shroud segment (1).

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