Inner shroud sector for an aircraft turbine engine

Abstract

The invention relates to an inner shroud sector (1) configured to be mounted in an aircraft turbine engine, the inner shroud being a rotating part having a longitudinal axis (X) oriented from upstream to downstream, the inner shroud sector (1) comprising a main body (2) having an outer surface (Sext) configured to internally delimit a secondary flow path, the main body (2) comprising an upstream end (3) configured to be attached to a hub of an intermediate casing of the aircraft turbine engine, the main body (2) being made of a composite material comprising a plurality of reinforcing fibres in a matrix, the inner shroud sector (1) comprising at least one hollow column (13) mounted on the upstream end (3) of the main body (2), each hollow column (13) extending radially and being configured to accommodate an attachment element for attaching the inner shroud sector (1) to the hub of the intermediate casing.

Description

Inner shell sector for aircraft turbomachine

[0001] The present invention relates to the field of aircraft turbomachine casings and more specifically targets an inner shell sector.

[0002] In a known manner, with reference to the, an aircraft turbomachine 500, of longitudinal axis X, comprises 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 together define a primary vein 400 for circulating an air flow, delimited externally by a central casing 270. The aircraft turbomachine 500 also comprises a secondary vein 410 for circulating an air flow which extends externally around the primary vein 400 and is surrounded by a nacelle 300.

[0003] In a known manner and as illustrated in the, the aircraft turbomachine 500 comprises an intermediate casing 190 comprising an inner shroud 100 and an outer shroud 180 which respectively internally and externally delimit the secondary duct 410. The inner shroud 100 and the outer shroud 180 are connected by arms 170 extending radially in the secondary duct 410, downstream of a row of fixed blades 330 (“outlet guide vanes (OGV)”). The outer shroud 180 extends in the downstream extension of the fan casing 210 and in the upstream extension of an outer nacelle casing 320 (“outer fan duct”). The inner shroud 100 extends in the upstream extension of an inner nacelle casing 310 (“inner fan duct”).The inner nacelle casing 310 and the central casing 270 together delimit an inter-vein compartment 280 (“compartment core” in English) extending radially between the primary vein 400 and the secondary vein 410.

[0004] In a known manner, the inner ferrule 100 is fixed upstream to an intermediate casing hub 160. The inner ferrule 100 is in practice formed by several angular inner ferrule sectors, typically in the form of single-piece titanium parts. Each inner ferrule sector is fixed to the hub 160 by a set of fixing screws inserted into through-recesses of the inner ferrule sector, formed by drilling into an excess thickness of titanium. The high mechanical strength of titanium allows the inner ferrule sectors to effectively transmit forces into the intermediate casing. Such inner ferrule sectors also form a fire barrier, thanks to the high thermal resistance to deformation of titanium. However, such inner ferrule sectors have the disadvantage of having a high mass.This has the effect of making the aircraft turbomachine 500 heavier and therefore increasing its energy consumption in flight.

[0005] It is known from application WO2010007220A2 to use a composite material for one or more elements of an aircraft turbomachine casing. Applications FR3108679A1, FR2992353A1 and FR3115832A1 are also known.

[0006] The invention thus aims to reduce the mass of an inner shell sector while retaining its fire barrier properties and its capacity to transmit mechanical forces into the intermediate casing. PRESENTATION OF THE INVENTION

[0007] The invention relates to an inner shell sector configured to be mounted in an aircraft turbomachine, the inner shell being a part of revolution defined relative to a longitudinal axis oriented from upstream to downstream, the inner shell sector comprising a main body having an outer surface configured to internally delimit a secondary vein of the aircraft turbomachine, the main body comprising an upstream end configured to be fixed to a hub of an intermediate casing of the aircraft turbomachine.

[0008] The invention is remarkable in that:the main body is made of a composite material comprising a plurality of reinforcing fibers in a matrix, andthe inner shell sector comprises at least one hollow column mounted on the upstream end of the main body, each hollow column extending radially and being configured to receive a fixing element for fixing the inner shell sector to the hub of the intermediate casing.

[0009] Thanks to the composite material, the mass of an inner shell sector according to the invention is reduced. The use of composite material for such an inner shell, which must have fairly good fire resistance, and which has numerous fixing points due to its production in the form of sectors assembled together, is unprecedented. In addition, the hollow columns according to the invention are added parts that advantageously make it possible to avoid forming excess thicknesses in the main body, which also contributes to reducing the mass. This also allows easy and uniform compression of the composite material during the manufacture of the main body.

[0010] According to one aspect of the invention, the upstream end of the main body comprises at least one first through hole, said at least one hollow column being mounted crimped in said at least one first through hole. Such an assembly advantageously makes it possible to obtain a fixing resistant to possible deformations of the composite material under the effect of heat.

[0011] According to one aspect of the invention, the upstream end of the main body comprises a first longitudinal wall, in which each first through-hole is formed, and a second longitudinal wall spaced radially from the first longitudinal wall and in which at least one second through-hole is formed, said at least one hollow column connecting said at least one first through-hole and said at least one second through-hole. Such a geometry does not require any local excess thickness and is advantageously suitable for a main body formed from a composite material.

[0012] According to one aspect of the invention, the upstream end of the main body comprises a third wall connecting the first longitudinal wall and the second longitudinal wall, the third wall being separated from each hollow column by a free volume. The absence of filling material between the hollow columns and the third wall allows a saving in mass.

[0013] According to one aspect of the invention, the main body comprises an inner surface opposite the outer surface, each first through hole opening onto the inner surface. The crimping is advantageously carried out on the side of the inner surface most likely to be subjected to mechanical and thermal deformations.

[0014] According to one aspect of the invention, the upstream end of the main body comprises a first inner longitudinal wall and a second outer longitudinal wall spaced radially apart from each other, at least one first through-hole being formed in the first longitudinal wall, at least one second hole being formed in the second longitudinal wall, said at least one hollow column connecting said at least one first through-hole and said at least one second through-hole, the upstream end of the main body comprising a third wall connecting the first longitudinal wall and the second longitudinal wall downstream of said at least one hollow column. Such a geometry does not require any local excess thickness and is advantageously suitable for a main body formed from a composite material. Such a geometry is particularly suitable for clamping a part of significant thickness while ensuring good transmission of mechanical forces.Significant thickness means an average thickness of the main body of the upstream end greater than 1 cm, preferably greater than 1.5 cm.

[0015] According to one aspect of the invention, said at least one hollow column is mounted crimped in said at least one first through hole. Such an assembly advantageously makes it possible to obtain a fixing resistant to possible deformations of the composite material under the effect of heat. The crimping is advantageously implemented on the side of the interior surface most likely to be subjected to mechanical and thermal deformations. In addition, the crimping contributes to protecting the turbomachine in the event of a fire by preventing the spread of heat. Indeed, the crimping makes it possible to prevent the spread of fire through the first through hole compared to an assembly with play or with crimping or gluing of the column at the second hole present in the second wall, on the external side.

[0016] According to one aspect of the invention, a clearance is present between said at least one hollow column and said at least one second through hole. The presence of clearance prevents hyperstatic mounting of the hollow columns on the main body.

[0017] According to one aspect of the invention, the first longitudinal wall is radially spaced from the second longitudinal wall by a distance greater than four times a thickness of the first longitudinal wall, preferably greater than eight times a thickness of the first longitudinal wall. The distance in question corresponds to the height of the hollow column and is preferably greater than 1 cm, preferably greater than 1.5 cm. The inner shell sector has a significant thickness for which the hollow column according to the invention allows clamping that effectively transmits mechanical forces while reducing the mass of the turbomachine. In the absence of a hollow column, a very large number of layers of composite material would be necessary to ensure the clamping of such a thickness, which is excluded.

[0018] According to one aspect of the invention, the third wall is separated from said at least one hollow column by a free volume. The absence of filling material between the hollow columns and the third wall allows a saving in mass.

[0019] According to one aspect of the invention, the first longitudinal wall, the second longitudinal wall and the third wall comprise a plurality of layers of composite material secured together, at least one layer called the fire-resistant layer extending continuously in the first longitudinal wall and the third wall. According to a preferred aspect, the fire-resistant layer defines the entire interior surface of the main body. The continuity of the layer of composite material thus forms a fire-resistant barrier on the side of the interior surface most likely to be subjected to mechanical and thermal deformations.

[0020] According to one aspect of the invention, said at least one hollow column is metallic, preferably made of stainless steel. Stainless steel is defined as a steel comprising less than 1.2% carbon and more than 10.5% chromium. Stainless steel is advantageously sufficiently deformable to allow crimping.

[0021] The invention also relates to an intermediate casing of an aircraft turbomachine comprising a hub and at least one inner shell sector as described previously, the intermediate casing extending along the longitudinal axis X, the intermediate casing comprising at least one element for fixing the upstream end of the inner shell sector to the hub, each fixing element extending in a hollow column of the inner shell sector.

[0022] The invention also relates to an aircraft turbomachine comprising a secondary vein and an intermediate casing comprising a hub and at least one inner shell sector as described previously, the aircraft turbomachine extending along the longitudinal axis X of the inner shell sector, the outer surface of the inner shell sector internally delimiting the secondary vein, the aircraft turbomachine comprising at least one element for fixing the upstream end of the inner shell sector to the hub of the intermediate casing, each fixing element extending in a hollow column of the inner shell sector. PRESENTATION OF FIGURES

[0023] The invention will be better understood upon reading the following description, given by way of example, and referring to the following figures, given by way of non-limiting examples, in which identical references are given to similar objects.

[0024] This is a schematic representation in longitudinal half-section of an aircraft turbomachine according to the prior art.

[0025] This is a schematic representation in longitudinal half-section of an aircraft turbomachine equipped with an intermediate casing with a set of inner shell sectors according to one embodiment of the invention.

[0026] This is a schematic perspective representation of the intermediate casing of the.

[0027] This is a schematic perspective representation from upstream of the inner shell sector according to one embodiment of the invention.

[0028] This is a schematic representation in longitudinal section of the inner shell sector of the fixed to the hub of the intermediate casing according to one embodiment of the invention.

[0029] This is a schematic representation of an operation of mounting a hollow column on the main body of the inner shell sector of the.

[0030] It should be noted that the figures set out the invention in detail to implement the invention, said figures can of course be used to better define the invention if necessary. DETAILED DESCRIPTION OF THE INVENTION

[0031] With reference to the and as described in the preamble, an aircraft turbomachine 50 conventionally extends along a longitudinal axis X oriented from upstream to downstream. Subsequently, the terms “upstream” and “downstream” are defined with respect to the orientation of the longitudinal axis X. The terms “inner” and “outer” are defined with respect to a radial axis extending orthogonally to the longitudinal axis X.

[0032] Still with reference to the, the aircraft turbomachine 50 conventionally comprises from upstream to downstream a fan 20, a low-pressure compressor 22, a high-pressure compressor 23, a combustion chamber 24, a high-pressure turbine 25 and a low-pressure turbine 26. The compressors 22, 23, the combustion chamber 24 and the turbines 25, 26 together define a primary vein 40 for circulating an air flow, delimited externally by a central casing 27. The aircraft turbomachine 50 also conventionally comprises a secondary vein 41 for circulating an air flow which extends externally around the primary vein 40 and is surrounded by a nacelle 30. Such an aircraft turbomachine 50 is said to be a double-flow turbomachine.

[0033] With reference to Figures 2 and 3, the aircraft turbomachine 50 comprises an intermediate casing 19 comprising an inner shroud 15 and an outer shroud 18 of longitudinal axis X which respectively internally and externally delimit the secondary flow path 41. The intermediate casing 19 also comprises a hub 16 on which the inner shroud 15 is fixed upstream. The hub 16 extends along the longitudinal axis X internally relative to the inner shroud 15. The hub 16 externally delimits the primary flow path 40 in the example of 1a. The intermediate casing 19 also comprises in this example arms 17 connecting the inner shroud 15 and the outer shroud 18, which extend radially in the secondary flow path 41, downstream of a row of fixed blades 33 (“outlet guide vanes (OGV)” in English).

[0034] With reference to Figures 2 and 3, the outer shroud 18 extends in the downstream extension of a fan casing 21 and, here, in the upstream extension of an outer nacelle casing 32 (“outer fan duct” in English). The inner shroud 15 extends opposite the outer shroud 18, in the upstream extension of an inner nacelle casing 31 (“inner fan duct” in English). The inner nacelle casing 31 and the central casing 27 typically together delimit an inter-vein compartment 28 (“compartment core” in English) extending radially between the primary vein 40 and the secondary vein 41.

[0035] As illustrated in the, the inner shroud 15 is a part of revolution defined with respect to the longitudinal axis X in the aircraft turbomachine 50, the longitudinal axis X subsequently designating indifferently the axis of the turbomachine 50 and the axis of the inner shroud 15 which are merged. As illustrated in the, the inner shroud 15 is formed by a set of inner shroud sectors 1 distributed angularly around the longitudinal axis X. The inner shroud sectors 1 extend over a variable angular width in this example.

[0036] According to the invention and as illustrated in Figures 4 and 5, each inner shell sector 1 comprises: a main body 2 made of a composite material comprising reinforcing fibers in a matrix, and one or more hollow columns 13 mounted on an upstream end 3 of the main body 2, each hollow column 13 extending radially and being configured to receive a fixing element 14 for fixing the inner shell sector 1 to the hub 16 of the intermediate casing 19.

[0037] Thanks to the composite material, the inner shell sectors 1 according to the invention advantageously have a reduced mass compared to the titanium inner shell sectors according to the prior art. The use of composite material in an inner shell sector 1 is unprecedented and goes against its current use, such as in acoustic panels fixed on a hood. Indeed, the shell must be able, due to its position close to the primary vein in the engine, to guarantee mechanical strength in the event of a fire, which discourages the use of a composite material. In addition, the shell is formed of several sectors connected by numerous fixing points, some of which have a significant clamping thickness, which also discourages the use of a composite material.As will be seen later, an inner shell sector 1 according to the invention makes it possible to form a fire barrier and to transmit mechanical forces into the intermediate casing.

[0038] As illustrated in Figures 3 and 4, the main body 2 of an inner shell sector 1 comprises an outer surface Sext curved towards the longitudinal axis X and an inner surface Sint opposite the outer surface Sext. The outer surface Sext internally delimits the secondary vein 41. The inner surface Sint is configured to extend radially against the hub 16 at the upstream end 3 when the intermediate casing 19 is mounted ().

[0039] As illustrated in Figures 3 and 4, the main body 2 comprises an upstream end 3 on which one or more hollow columns 13 are mounted to ensure fixing to the hub 16 of the intermediate casing 19. The number of hollow columns 13 depends on the angular width of the inner shell sector 1, preferably at least equal to two. The hollow columns 13 are preferably distributed over the angular width of the upstream end 3. A single hollow column 13 and its mounting in the main body 2 are described below, this description being valid for each hollow column 13.

[0040] According to a preferred aspect illustrated in Figures 3 and 4, the upstream end 3 comprises a first longitudinal wall 6 and a second longitudinal wall 4 spaced radially apart and comprising respectively a first through hole 8 and a second through hole 7 for each hollow column 13. The second longitudinal wall 4 preferably extends in the upstream extension of the main body 2. The first longitudinal wall 6 extends radially inward relative to the second longitudinal wall 4. As illustrated in the, the radial distance H separating the longitudinal walls 4, 6, corresponding to the height of the hollow column 13, is preferably greater than four times, preferably eight times, a thickness E of the first longitudinal wall 6 or of the second longitudinal wall 4. The radial distance H is preferably greater than 1 cm, preferably greater than 1.5 cm.The first through hole 8 and the second through hole 7 associated with the same hollow column 13 are aligned radially along a radial axis Y, the hollow column 13 extending along the radial axis Y in the through holes 7, 8.

[0041] According to a preferred aspect illustrated in Figures 3 and 4, the longitudinal walls 4, 6 are connected together by a third wall 5. The third wall 5 extends downstream of the longitudinal walls 4, 6. In this example, the third wall 5 extends radially and forms with the longitudinal walls 4, 6 a U whose concavity is turned upstream. Here, the longitudinal walls 4 and 6 extend along the entire upstream end 3 of the main body 2, and delimit, with the third wall 5, a channel, in this case with a U-shaped section, open on the upstream side. Alternatively, the third wall 5 extends in a direction forming, in a plane defined by the longitudinal axis X and a radial axis, an angle of between 10° and 90° relative to the longitudinal axis X. In this alternative, the third wall 5 preferably extends in such a way that the first longitudinal wall 6 comprises a longitudinal length less than that of the second longitudinal wall 4.

[0042] As described above, the main body 2 is made of a composite material in the form of reinforcing fibers trapped in a matrix. The main body 2 preferably has a single-piece structure. The reinforcing fibers are, for example, made of carbon, to give the main body 2 good mechanical strength. The matrix is ​​obtained by heating and compressing a thermosetting material, for example a polyepoxide, known as “epoxy”. Such a material has a malleable structure configured to harden after heating and pressurizing to obtain its final shape.

[0043] As illustrated in the, the main body 2 comprises a set of layers Ca, Cb, Cc of composite material secured together by heating and compression, and forming the single-piece structure. The main body 2 preferably comprises:An inner layer Ca defining the entire inner surface Sint of the main body 2, in particular extending continuously in the first longitudinal wall 6 and in the third wall 5 of the upstream end 3,An outer layer Cc defining the entire outer surface Sext of the main body 2, in particular extending continuously in the second longitudinal wall 4 of the upstream end 3,An intermediate layer Cb extending only in the upstream end 3, radially between the outer layer Cc and the inner layer Ca, namely continuously in the first longitudinal wall 4, in the third wall 5 and in the second longitudinal wall 6.

[0044] In this example, the inner layer Ca has a substantially “S” shaped section, the outer layer Cc has a substantially straight section and the intermediate layer Cb has a substantially “U” shaped section.

[0045] The continuous arrangement of the inner layer Ca, called the fire-resistant layer Ca, advantageously makes it possible to form a fire barrier between the primary vein 40 and the secondary vein 41. The fire-resistant layer Ca contributes in particular to confining any possible outbreak of fire and preventing the spread of heat. The fire barrier is also ensured at the first through holes 8 thanks to a particular assembly by crimping the hollow columns 13, as will be seen later.

[0046] In this example, the main body 2 comprises only the three layers Ca, Cb, Cc. Alternatively, the body also comprises other layers sandwiched between the layers Ca, Cb, Cc, preferably in such a way as to maintain a substantially constant thickness.

[0047] As illustrated in Figures 5 and 7, each hollow column 13 is an insert in the form of a through-fixing housing mounted radially in the upstream end 3 between the first through-hole 8 and the second through-hole 7 along a radial axis Y. The hollow column 13 comprises a radial opening 13b, configured to receive a fixing element 14 such as a screw. As illustrated in the, the hollow column 13 also comprises a first end 13c cooperating with the first through-hole 7 and a second end 13a cooperating with the second through-hole 8.

[0048] According to a preferred aspect, the first end 13c is mounted crimped in the first through hole 8, namely inserted into the first through hole 8 by material deformation. Such a fixing forms a fire barrier at the first through holes 8, in the extension of the fire-resistant layer Ca. Indeed, the crimping makes it possible to prevent the propagation of fire through the first through hole 8 compared to a conventional assembly, in particular with play. The crimping makes it possible to maintain a form of sealing at the level of each first through hole 8 (sealing between the column 13 and the fire-resistant layer Ca) even if the composite material of the main body 2 deforms slightly under the effect of heat.

[0049] Preferably, as illustrated in Figures 5 and 7, the first end 13c and the first through hole 8 comprise a frustoconical section increasing from the outside towards the inside along the radial axis Y which cooperate together. The hollow column 13 is inserted from the outside towards the inside, then the frustoconical section of the first end 13c is deformed to allow its cooperation with the first through hole 8 (ie: to crimp it on the first longitudinal wall 6). Also preferably, the first end 13c comprises a stop section configured to come into abutment radially from the outside against the first longitudinal wall 6. The stop section has a diameter greater than that of the first through hole 8. The frustoconical section and the stop section of the end 13c together ensure the radial locking in translation respectively towards the outside and towards the inside of the hollow column 13.

[0050] Alternatively, the first end 13c is secured by gluing or riveting in the first through hole 8.

[0051] According to a preferred aspect, the second end 13a extends freely in the second through hole 7, i.e. is mounted with play in the second through hole 7. This avoids hyperstatic mounting of the hollow column 13 on the main body 2 and limits mechanical stresses likely to reduce the service life. As illustrated in the, the second through hole 7 comprises a section greater than the external section of the hollow column 13 so that the hollow column 13 is inserted radially via the end 13a to be mounted on the main body 2.

[0052] The hollow column 13 is preferably made of metal to ensure good mechanical strength, preferably stainless steel. Stainless steel is advantageously deformable, which allows for easy crimping, unlike titanium, for example. Stainless steel is defined as a steel containing less than 1.2% carbon and more than 10.5% chromium. The hollow column 13 is preferably made of a single piece.

[0053] According to a preferred aspect illustrated in Figures 4 and 5, the hollow column 13 is separated from the third wall 5 by a free volume V. Thanks to the crimped assembly, no filling material is necessary between the hollow column 13 and the third wall 5. This contributes to reducing the mass of the inner shell sector 1, in a manner complementary to the choice of the composite material. Preferably, the U-shaped channel, delimited by the three walls 4, 5, 6, has a radial width at least equal to its longitudinal depth, preferably at least equal to 1 cm, preferably at least equal to 1.5 cm. For such a geometry, the use of a filling or reinforcing material would be inconvenient and ineffective.

[0054] Such a hollow column 13 mounted in an attached manner on the main body 2 advantageously makes it possible to simplify the geometry of the main body 2, in particular with a constant thickness, in order to be able to form the main body 2 in composite material. In addition, the hollow column 13 makes it possible to avoid forming an excess thickness in the third wall 5 in order to form housings by drilling as in the prior art.

[0055] As illustrated in the, the hollow columns 13 are inserted radially into the second through holes 7 and then fixed, here by crimping, in the first through holes 8 to form the inner shell sector 1. The inner shell sectors 1, together defining the inner shell 15, are then mounted in the aircraft turbomachine 50 using the fixing elements 14. The fixing elements 14 are inserted radially from the outside into the hollow columns 13 and into the hub 16 of the intermediate casing 19, to fix the inner shell sectors 1 on the hub 16.

Claims

Inner shell sector (1) configured to be mounted in an aircraft turbomachine (50), the inner shell being a part of revolution defined with respect to a longitudinal axis (X) oriented from upstream to downstream, the inner shell sector (1) comprising a main body (2) having an outer surface (Sext) configured to internally delimit a secondary vein (41) of the aircraft turbomachine (50), the main body (2) comprising an upstream end (3) configured to be fixed on a hub (16) of an intermediate casing (19) of the aircraft turbomachine (50), the main body (2) being made of a composite material, the inner shell sector (1) being characterized in that: the composite material of the main body (2) comprises a plurality of reinforcing fibers in a matrix, and the inner shell sector (1) comprises at least one hollow column (13) mounted on the upstream end (3) of the main body (2),each hollow column (13) extending radially and being configured to receive a fixing element (14) for fixing the inner shell sector (1) on the hub (16) of the intermediate casing (19)., Inner shell sector (1) according to claim 1, wherein the upstream end (3) of the main body (2) comprises a first inner longitudinal wall (6) and a second outer longitudinal wall (4) spaced radially from each other, at least one first through hole (8) being formed in the first longitudinal wall (6), at least one second through hole (7) being formed in the second longitudinal wall (4), said at least one hollow column (13) connecting said at least one first through hole (8) and said at least one second through hole (7), the upstream end (3) of the main body (2) comprising a third wall (5) connecting the first longitudinal wall (6) and the second longitudinal wall (4) downstream of said at least one hollow column (13). Inner ferrule sector (1) according to claim 2, wherein said at least one hollow column (13) is mounted crimped in said at least one first through hole (8). Inner ferrule sector (1) according to one of claims 2 and 3, in which a clearance is present between said at least one hollow column (13) and said at least one second through hole (7). Inner shell sector (1) according to one of claims 2 to 4, in which the first longitudinal wall (6) is radially spaced from the second longitudinal wall (4) by a distance (H) greater than four times a thickness (E) of the first longitudinal wall (6), preferably greater than eight times a thickness (E) of the first longitudinal wall (6). Inner shell sector (1) according to one of claims 2 to 5, wherein the third wall (5) is separated from said at least one hollow column (13) by a free volume (V). Inner shell sector (1) according to one of claims 2 to 6, in which the first longitudinal wall (6), the second longitudinal wall (4) and the third wall (5) comprise a plurality of layers (Ca, Cb, Cc) of composite material secured together, at least one layer called fire-resistant layer (Ca) extending continuously in the first longitudinal wall (6) and the third wall (5). Inner shell sector (1) according to claim 7, wherein the fire-resistant layer (Ca) defines the entire inner surface (Sint) of the main body (2). Inner shell sector (1) according to one of claims 1 to 8, wherein said at least one hollow column (13) is metallic, preferably stainless steel. Aircraft turbomachine (50) comprising a secondary vein (41) and an intermediate casing (19) comprising a hub (16) and at least one inner shell sector (1) according to one of claims 1 to 9, the aircraft turbomachine (50) extending along the longitudinal axis X of the inner shell, the outer surface (Sext) of the inner shell sector (1) internally delimiting the secondary vein (41), the aircraft turbomachine (50) comprising at least one fixing element (14) of the upstream end (3) of the inner shell sector (1) on the hub (16) of the intermediate casing (19), each fixing element (14) extending in a hollow column (13) of the inner shell sector (1).

Images