• 专利附录
  • 专利说明
  • 权利要求
  • 类似技术
  • 同族专利
  • 引用文献
  • 法律状态
  • 优先权
    • 专利摘要:
    The main object of the invention is an intermediate case hub for an aircraft turbojet, comprising: an outer shell (14) intended to externally delimit a secondary flow space of a secondary gas flow and internally an inter zone veins, the outer ferrule (14) being provided with a secondary orifice (29), and a discharge valve comprising a discharge duct (30) of composite material, located in the inter-vein zone, characterized in that the discharge (30) is attached to the outer shroud (14) at the secondary orifice (29), at least one air and fire seal (33) being disposed between the discharge conduit (30). ) and the outer shell (14), and in that the discharge pipe (30) of composite material comprises a composite wall (30a, 30b) draped, consisting of a plurality of folds impregnated with resin.
    公开号:FR3036136A1
    申请号:FR1554378
    申请日:2015-05-15
    公开日:2016-11-18
    发明作者:Forian Benjamin Kevin Lacroix；Cyrille Francois Antoine Mathias；Idaline Francoise Chantal Texier；Maxime Marie Desiree Blaise
    申请人:Safran SA；SNECMA SAS；
    IPC主号:
    专利说明:

    [0001] Description: The present invention relates to the field of aircraft turbojet engines, and more particularly to the general field of double-body and dual-flow turbojets. BACKGROUND OF THE INVENTION The invention thus relates to an intermediate casing hub for an aircraft turbojet engine, in particular of the type comprising at least two mechanically independent bodies. In a double-body turbojet engine, the term "intermediate casing" is usually used to designate a casing whose hub is substantially arranged between a low-pressure compressor casing and a high-pressure compressor casing.
    [0002] The present invention relates more particularly to an intermediate casing hub of the type comprising discharge valves (also known by the acronym VBV for "Variable Bleed Valves"). The discharge valves are intended to regulate the inlet flow of the high-pressure compressor, in particular in order to limit the risks of pumping the low-pressure compressor, by allowing part of the air to escape from the annular space of the compressor. flow of the primary flow. In addition, in case of accidental penetration into this flow space, water, especially in the form of rain or hail, or various debris, which are likely to affect the operation of the turbojet engine, these discharge valves allow recovering this water or debris which is centrifuged in the aforementioned flow space and ejecting it outwardly of the latter. In the case of turbojet engines, these relief valves are thus configured to allow the passage of air, water or debris from the flow space of the primary flow to an annular flow space of a flow. secondary stream. To do this, the discharge valves comprise in particular conduits for discharging the primary flow to the secondary flow connecting ports respectively communicating with the primary flow and the secondary flow. Thus, more specifically, the invention relates to an intermediate case hub for an aircraft turbojet engine comprising an air and fire seal at a discharge duct provided with a draped composite wall. a plurality of plies impregnated with resin, an intermediate casing comprising such a hub, and an aircraft turbojet engine comprising such an intermediate casing. STATE OF THE ART In the field of aircraft turbojet engines, the use of composite materials is becoming more and more frequent. In particular, many aeronautical parts are now made, at least in part, from organic matrix composite materials (generally designated by the acronym CMO for "Organic Matrix Composite"). Thus, it may be envisaged to make the discharge ducts, which equip the previously described discharge valves with an intermediate casing hub for an aircraft turbojet engine, from composite materials, and in particular organic matrix composites (CMOs). ). However, composite materials are particularly sensitive to the risk of fire, and in particular organic matrix composites (CMO) because the organic resin forming the matrix is combustible. However, these composite materials are often used in turbojet equipment located in areas at risk of fire. In particular, the discharge ducts of the discharge valves are typically located in a fire hazard zone, namely the inter-vein zone of the turbojet engine, so that it is necessary to prevent any fresh air supply of a contained in this zone. By way of illustration of the technical context of the invention, FIG. 1 partially shows, in axial section, an example of a hub 10 of an intermediate casing 11 for an aircraft turbojet engine 12 having a double body and a double flow. a known type. The hub 10 of the intermediate casing 11 usually comprises two coaxial annular ferrules, respectively internal 13 and external 14, mutually connected by two transverse flanges, namely an upstream transverse flange 15 and a downstream transverse flange 16. The upstream transverse flange 15 is arranged downstream of a low-pressure compressor 17 of the turbojet engine 12, while the downstream transverse flange 16 is arranged upstream of a high-pressure compressor 18 of this turbojet engine 12. This high-pressure compressor 18 generally comprises a succession of rotors and variable-speed stators, to control the flow of air passing through it. Furthermore, between the inner and outer shells 13 and 14 between the upstream and downstream transverse flanges 15, intermediate spaces 19 are provided distributed around the axis of the hub 10, coinciding with the axis of rotation T of the turbojet engine 12. The intermediate spaces 19 are upstream of an inter-vein zone ZC. In addition, the inner shell 13 delimits an annular primary flow space 20 of a primary flow of the turbojet engine 12. Furthermore, the inner shell 13 has air passages 21, called primary orifices in the following, each of which is closed by the pivoting valve 22 of a corresponding discharge valve 23, 20 for the regulation of the flow of the high-pressure compressor 18, and where appropriate, the evacuation of air, water or debris as explained before. Such a discharge valve 23 usually takes the form of a door 24, which comprises the pivoting valve 22 at its radially inner end and which is pivotally mounted about an axis Y so that in the closed position of the primary orifices 25 21, the valve 22 extends the inner ferrule 13 of the intermediate casing 11 substantially continuously to reduce the risk of aerodynamic disturbances of the primary flow by this valve 22, and that in the open position of said primary orifices 21, the valve 22 protrudes radially inwardly relative to the inner shell 13 and thus forms a bail for a portion of the primary flow 30 in the space 20. The door 24 comprises a conduit 25 through which air scoop transits, this conduit 26 ending downstream on an outlet port 26 opening into the corresponding intermediate space 19. The patent application FR 2 961 251 A1 of the Applicant further describes a Another example of a discharge valve of an aircraft turbojet intermediate case hub.
    [0003] Moreover, the outer shell 14 delimits an annular secondary flow space 27 of a secondary flow F2 of the turbojet engine 12, and is connected to structural arms 28, relatively spaced apart from each other, passing through this space 27. , the outer shell 14 has air passages 29, called secondary orifices in the following, and arranged downstream of the downstream transverse flange 16. Otherwise, in this example of FIG. air, water or debris is through the outer shell 14. However, alternatively (not shown), when for example the outer shell 14 carries guide vanes relatively close to each other, the latter In this case, it may be desirable to allow this evacuation further downstream, through the annular wall of an extension of the hub of the intermediate casing. to say the wall annular of a structural part which is sometimes used to support at its downstream end thrust reverser elements such as fairing panels. When the variable-pitch stators of the high-pressure compressor 18 are in a position reducing the flow rate of air entering this compressor, an excess of air in the secondary flow space can then be discharged through the secondary orifices 29, thus avoiding pumping phenomena which can lead to deterioration or even complete destruction of the low-pressure compressor 17. In addition, as previously explained, discharge ducts 25 each extend between a respective inlet opening 31 opening into the intermediate space 19 and a corresponding secondary orifice 29. Inside these discharge ducts 30 circulates a discharge flow FD, originating from the primary flow, in the direction of the secondary flow F2. The inlet orifice 31 is generally arranged flush with the surface of the downstream transverse flange 16 giving onto the intermediate space 19. The secondary orifice 29 comprises, in turn, a control gate 32, fixed to the discharge duct 30 at its output, to be able to control the discharge flow FD during its rejection in the secondary flow F2. In each intermediate space 19, the outlet orifice 26 of the primary duct 25 and the inlet orifice 31 of the discharge duct 30 are arranged facing each other.
    [0004] Each door 24, the intermediate space 19 and the corresponding downstream discharge duct 30 thus together form a system for discharging air, water or debris, generally designated by the term "discharge valve". from the primary flow space 20 to the secondary flow space 27. The hub 11 thus comprises a plurality of such systems distributed around its axis T.
    [0005] When a door 24 is in the open position, a stream of air ecoped by it through the primary conduit 25, opens into the intermediate space 19 through its outlet port 26, enters the corresponding discharge conduit 30. to reach the secondary flow space 27. As can be seen in FIG. 1, the discharge ducts 30 are fixed on the one hand to the outer shell 14, in particular by screws, and on the other hand 16. They are located in an interveil zone ZC of the turbojet engine 12 which presents a risk of fire, as indicated above. Such an inter-vein zone is commonly referred to as a "core zone" using an English terminology. It is thus necessary to provide a solution to avoid fresh air supply of the inter-vein zone ZC of the turbojet engine 12, from the primary flow or the secondary flow. SUMMARY OF THE INVENTION Accordingly, there is a need to provide a solution for improving the reliability and efficiency of using a discharge valve discharge duct of a turbojet intermediate casing hub. aircraft, made of one or more composite materials, in an area of the turbojet engine with a risk of fire. In particular, there is a need to design one or more interfaces of the composite discharge duct with one or more parts of its environment, for example an outer shell, which are fire-resistant, in particular fire-tight 3036136 6 according to the ISO standard 2685. In addition, there is a need to provide such an airtightness of a discharge line interface with its environment, even in the absence of fire, in order to improve the performance of the turbojet engine. The object of the invention is therefore to at least partially remedy the needs mentioned above and the drawbacks relating to the embodiments of the prior art. The invention thus has, according to one of its aspects, an intermediate case hub for an aircraft turbojet, comprising: an internal annular shroud for delimiting externally a primary flow space of a primary gas flow in the turbojet, and secondly internally upstream of at least one inter-vein zone, the inner annular shroud being provided with at least one primary air passage orifice, - a ferrule outer ring for delimiting externally a secondary flow space of a secondary gas flow in the turbojet, and secondly internally said at least one inter-vein zone, the outer annular shell being provided with at least one secondary air passage orifice, at least one discharge valve, comprising at least one movable door capable of taking, from said at least one primary orifice, air flowing in the primary flow space e t to return to said at least one inter-vein zone the air thus withdrawn in the direction of at least one corresponding discharge inlet orifice of at least one discharge duct made of composite material and shaped to ensure a passage of air from said at least one discharge inlet port to said at least one secondary port for returning air withdrawn via said at least one discharge valve into the secondary flow space, characterized in that said at least one discharge duct is attached to the outer annular shell at said at least one secondary orifice, at least one air and fire seal being disposed between said at least one discharge duct and the outer annular shell, and in that said at least one composite material discharge conduit has a draped composite wall made of a plurality of resin impregnated plies.
    [0006] Thanks to the invention, it is possible to ensure the required fire and air tightness of the interface between a discharge duct of an aircraft turbojet engine intermediate casing hub and the outer shroud. of the hub through the good fire behavior of the composite material discharge pipe and the use of the fire-resistant seal. In particular, the invention proposes a solution involving sufficient mechanical strength in case of risk of fire and fulfilling the requirements of the ISO 2685 standard. Moreover, the principle according to the invention makes it possible to avoid recourse to the use of a protection of the metal type. Indeed, the functions of airtightness and fire are provided by the presence of the seal and the embodiment in stack of plies impregnated discharge conduit of composite material. Thus, the invention provides a significant gain in terms of mass, allowing the reduction of the number of parts needed to perform the functions of sealing and fire resistance. Furthermore, the invention may make it possible to envisage providing an interface between the discharge duct and the outer shell that is ready to be mounted. The intermediate casing hub according to the invention may further comprise one or more of the following features taken separately or in any possible technical combinations. The intermediate casing hub may in particular comprise a downstream transverse flange, connecting the inner and outer annular ferrules, delimiting upstream at least one intermediate space and downstream said at least one inter-vein zone, the downstream transverse flange comprising said at least one discharge inlet port. The air circulating in the primary flow space, taken from said at least one primary orifice, by said at least one movable door can be adapted to be returned to said at least one intermediate space in the direction of said at least one orifice discharge inlet, said at least one discharge valve comprising said at least one discharge conduit, located in said at least one inter-vein zone and connecting said at least one discharge inlet port and said at least one secondary port; , the collected air being able to circulate in said at least one intermediate space and be returned to the secondary flow space.
    [0007] Said at least one airtight and fire-proof seal may especially be made, at least in part, of silicone. Said at least one air and fire seal may be composed of a superposition of different folds of fabrics, in particular glass and / or ceramic.
    [0008] The outer annular shell may or may not have an annular boss. Similarly, said at least one discharge duct may or may not comprise an annular skew. When the outer annular shell comprises an annular boss and the said at least one discharge duct comprises an annular cavity, the fixing of the at least one discharge duct to the outer annular collar can be achieved by means of the annular boss and the annular grinding, in particular by means of a screwing through the annular boss and the annular grinding. The annular boss may extend totally or partially around the at least one seal to the air and fire. Likewise, the annular furring may extend totally or partially around said at least one air and fire seal. Preferably, the assembly formed by the annular boss and the annular recess may extend all around said at least one air and fire seal, forming a separation between said at least one seal and said at least one inter-vein area.
    [0009] As a variant, said at least one discharge duct may comprise a partial annular recess, extending over at least two opposite edges of said at least one discharge duct, in particular upstream and downstream edges, fixing said at least one duct discharge to the outer annular shroud being carried out by means of partial annular grinding, in particular by means of a screwing through the partial annular recess 25 for raising said composite wall, said at least one seal being housed between said composite wall and the outer annular shell. Then, the outer annular shell may be devoid of annular boss. Furthermore, a control gate may be disposed at said at least one secondary orifice. Said at least one air and fire seal may be arranged all around the control gate, the control gate being fixed to said at least one discharge duct, in particular by means of a screwing . In addition, said at least one air and fire seal may be located between the control grid and the assembly formed by the annular boss and the annular sieve. Moreover, the thickness of said at least one discharge duct and / or the number of plies impregnated with said at least one discharge duct may be chosen in particular as a function of the composite material (s) and / or architecture of said at least one discharge duct. This thickness and / or this number of impregnated folds are preferably sufficient to ensure the fire barrier function. Thus, for example, said at least one discharge duct may comprise a composite wall with a thickness of at least 1.5 mm, in particular 2 mm. In addition, the composite wall of said at least one discharge duct may comprise a number of impregnated folds at least equal to 3, in particular at least 4.
    [0010] In a preferred manner, said at least one discharge duct may in particular comprise a draped composite wall with a thickness of at least 2 mm and a number of impregnated folds at least equal to 4. In addition, said at least one discharge duct The composite material can in particular be made from a thermosetting resin of the bismaleimide type, for example such as marketed by CYTEC under the reference CYCOM®5250-4, and a laminate of carbon braids, in particular biaxial braids. and / or triaxial. The stacking strategy of the impregnated folds constituting the composite wall of said at least one discharge duct advantageously directly influences the integrity of said at least one discharge duct subjected to a fire hazard. In addition, the invention also relates, in another of its aspects, to an intermediate casing for an aircraft turbojet, characterized in that it comprises a hub as defined above.
    [0011] In addition, the invention further relates, in another of its aspects, to an aircraft turbojet, characterized in that it comprises an intermediate casing as defined above. The intermediate casing hub, the intermediate casing and the aircraft turbojet according to the invention may comprise any of the features set out in the description, taken individually or in any technically possible combination with other characteristics. BRIEF DESCRIPTION OF THE DRAWINGS The invention will be better understood on reading the detailed description which follows, of an example of non-limiting implementation thereof, as well as on the examination of the figures, diagrams and partial, of the accompanying drawing, in which: - Figure 1 shows, in axial section, an example of a hub of an intermediate casing for an aircraft turbojet, - Figure 2 illustrates, in schematic and partial axial section, the principle attaching a discharge duct to the outer shell of an aircraft turbojet intermediate casing hub according to the invention, that is to say the realization of the interface between the discharge duct and the outer shell of the hub, - Figures 3 and 4 respectively show, in schematic and partial axial section, a principle of embodiment of the upstream wall and the downstream wall of the discharge duct illustrated in Figure 2, and - the figure 5 illustrates, according to a view in partial perspective, an exemplary alternative embodiment of the discharge duct of the aircraft turbojet intermediate casing hub according to the invention of Figures 2, 3 and 4. In all of these figures, identical references may designate Identical or similar elements. In addition, the different parts shown in the figures are not necessarily in a uniform scale, to make the figures more readable.
    [0012] DETAILED DESCRIPTION OF A PARTICULAR EMBODIMENT Throughout the description, it is noted that the terms upstream and downstream are to be considered with respect to a main direction F of normal gas flow (from upstream to downstream ) for a turbojet engine 12. Furthermore, the T-axis of the turbojet engine 12 is called the radial axis of symmetry of the turbojet engine 12. The axial direction of the turbojet engine 12 corresponds to the axis of rotation of the turbojet engine 12, which is the direction of rotation. the axis T of the turbojet engine 12. A radial direction of the turbojet engine 12 is a direction perpendicular to the axis T of the turbojet engine 12. In addition, unless otherwise stated, the adjectives and adverbs axial, radial, axially and radially are used as reference 10 in the aforementioned axial and radial directions. In addition, unless otherwise stated, the terms inner (or inner) and outer (or outer) are used with reference to a radial direction so that the inner part of an element is closer to the T axis of the turbojet engine 12 than the outer part of the same element. FIG. 1 has already been described previously in the section relating to the prior art and the technical context of the invention. Referring to Figure 2, there is shown, in schematic and partial axial section, a principle of attachment of a discharge duct 30 to the outer shell 14 of a hub 10 of intermediate casing 11 of an aircraft turbojet engine 12 according to FIG. an exemplary embodiment of the invention, that is to say the embodiment of the interface between the discharge duct 30 and the outer shell 14 of the hub 10. Furthermore, FIGS. 3 and 4 respectively illustrate, in schematic and partial axial section, , a principle of embodiment of the upstream composite wall 30b and the downstream composite wall 30a of the discharge duct 30. In addition, FIG. 5 illustrates, in a partial perspective view, an alternative embodiment of the discharge duct 30 of the Figures 2, 3 and 4. The hub 10 of the intermediate casing 11 according to the invention, associated with Figures 2, 3 and 4 described below, may in particular be of the same type as that described previously with reference to Figure 1. Auss i, for the parts not shown in Figures 2, 3 and 4, reference should be made to the previous description of Figure 1.
    [0013] As described above, the discharge duct 30 is located in the inter-vein zone ZC and connects the discharge inlet port 31 and the secondary orifice 29. The discharge duct 30 is then capable of sampling from the discharge inlet 31, air flowing in the intermediate space 19 and return to the secondary flow space 27 the air thus withdrawn. According to the invention, the discharge duct 30 is fixed to the outer annular shell 14 at the secondary orifice 29. In addition, a gasket 33 to air and fire is disposed between the duct discharge 30 and the outer ring ferrule 14.
    [0014] More specifically, as can be seen in FIG. 2, the outer ring 14 comprises an annular boss 37 and the discharge duct 30 comprises an annular ring 36. The attachment of the discharge duct 30 to the outer shell 14 is then realized. by means of a screwing 34 through the annular boss 37 and the annular recess 36.
    [0015] In addition, advantageously, the assembly formed by the annular boss 37 and the annular sander 36 extends all around the seal 33, forming a separation between the seal 33 and the inter-vein zone ZC . Furthermore, a control gate 32 is disposed at the secondary orifice 29. The seal 33 is then arranged all around the control gate 32, which is fixed to the discharge duct 30 via Thus, the seal 33 is located between the control grid 32 and the assembly formed by the annular boss 37 and the annular sander 36. The seal 33 can for example be made, at the less in part, silicone. In particular, it may comprise a superposition of different folds of fabrics, in particular glass and / or ceramic. In addition, the discharge duct 30 is made of composite material. It comprises an upstream composite wall 30b and a downstream composite wall 30a. According to the invention, the upstream 30b and downstream 30a walls are draped, that is to say they are obtained by draping a plurality of plies impregnated with resin, these plies 3036136 13 including including carbon braids, biaxial or triaxial, and the resin being in particular of the bismaleimide type. The thickness of the discharge duct 30 and the number of impregnated folds of the discharge duct 30 may be chosen as a function of the composite material (s) 5 and the architecture of the discharge duct 30. As can be seen in FIG. , the downstream composite wall 30a of the discharge duct 30 comprises in particular three successive portions al, a2 and a3. The first portion al comprises, for example, two large diameter carbon T2 biaxial braids with a thickness of approximately 0.55 mm, and for example at least six or even eight flat carbon triaxial T3 braids of equal thickness. about 0.25 mm. In this way, the thickness E1 of the first portion a1 is at least about 2.6 mm or even at least about 3.1 mm. In addition, the second portion a2 comprises, for example, two large diameter carbon T2 biaxial braids with a thickness of about 0.55 mm, and for example two small, diameter-diameter, carbon T1 biaxial braids. equal to about 0.55 mm. In this way, the thickness E2 of the second portion a2 is about 2.7 mm. In addition, the third portion a3 comprises for example two biaxial braids T2 of carbon, large diameter, two biaxial braids T1 of carbon, small diameter, thickness equal to about 0.55 mm. In this way, the thickness E3 of the third portion a3 is about 2.2 mm. Moreover, as can be seen in FIG. 4, the upstream composite wall 30b of the discharge duct 30 also comprises in particular three successive portions b1, b2 and b3. The first portion b1 comprises, for example, two large diameter carbon T2 biaxial braids with a thickness of approximately 0.55 mm, and for example at least six or even eight flat carbon triaxial T3 braids of equal thickness. about 0.25 mm. In this way, the thickness E4 of the first portion b1 is at least about 2.6 mm, or even at least about 3.1 mm, or even at least about 3.8 mm. In addition, the second portion b2 comprises, for example, two large diameter carbon T2 biaxial braids with a thickness of about 0.55 mm, and for example two small diameter carbon, small diameter biaxial braids T1, thickness equal to about 0.55 mm. In this way, the thickness E6 of the second portion b2 is about 2.7 mm. In addition, the third portion b3 comprises, for example, two large diameter carbon T2 biaxial braids, two small diameter carbon biaxial braids T1, with a thickness of about 0.55 mm. In this way, the thickness E7 of the third portion b3 is about 2.2 mm. Finally, the thickness E5 of the upstream composite wall 30b at the bend is for example about 1.6 mm. Furthermore, FIG. 5 shows a partial embodiment of the discharge duct 30 of the intermediate casing hub 11 according to the invention described previously with reference to FIGS. 2, 3 and 4. For example, the outer annular ferrule 14 (not visible in FIG. 5) is devoid of annular boss. Then, the discharge duct 30 is coupled to the outer annular shell 14 without surface offset.
    [0016] To do this, more precisely, the discharge duct 30 comprises an annular ring 36 which is only partial. This extends only on the two opposite upstream edges 38b and downstream 38a of the discharge duct 30. Then, the attachment of the discharge duct 30 to the outer annular shell 14 can for example be achieved by means of the ring cavity 36 partially thanks to screwing through the partial ring recess 36 to raise the composite walls 30a, 30b. The seal 33, necessary to prevent flames from passing through the sides of the duct 30, is thus housed between the composite walls 30a, 30b and the outer annular shell 14. Of course, the invention is not limited to the embodiment which has just been described. Various modifications may be made by the skilled person.
    权利要求:
    Claims (13)
    [0001]
    REVENDICATIONS1. Hub (10) for intermediate casing (11) for an aircraft turbojet engine (12), comprising: an inner annular ring (13) intended to delimit externally a primary flow space (20) of a flow primary gas in the turbojet engine (12), and secondly internally upstream of at least one inter-vein zone (ZC), the inner annular shell (13) being provided with at least one primary orifice for the passage of (21), an outer annular shroud (14) for delimiting externally a secondary flow space (27) of a secondary gas flow (F2) in the turbojet engine (12), and on the other hand internally said at least one inter-vein zone (ZC), the outer annular ferrule (14) being provided with at least one secondary air passage orifice (29), - at least one discharge valve (23) , comprising at least one movable door (24) adapted to take, from said at least one primary orifice (21), air flowing in the space this primary flow (20) and return to said at least one inter-vein area (ZC) the air thus withdrawn towards at least one corresponding discharge inlet port (31) of at least one duct discharge device (30) made of composite material and shaped to provide an air passage from said at least one discharge inlet (31) to said at least one secondary orifice (29) for returning air taken via said at least one minus a discharge valve (23) in the secondary flow space (27), characterized in that said at least one discharge duct (30) is attached to the outer annular shell (14) at said at least one secondary orifice (29), at least one air and fire seal (33) being disposed between said at least one discharge duct (30) and the outer annular ferrule (14), and in that said at least one discharge duct (30) of composite material comprises a composite wall (30a, 30b) draped, constituted a plurality of plies impregnated with resin. 3036136 16
    [0002]
    2. Crankcase hub according to claim 1, characterized in that it comprises: - a downstream transverse flange (16), connecting the annular inner ring (13) and outer 5 (14), delimiting upstream at least one space intermediate (19) and downstream said at least one inter-vein area (ZC), the downstream transverse flange (16) comprising said at least one discharge inlet (31), the air flowing in the space of primary flow (20), taken from said at least one primary orifice (21), by said at least one movable gate (24) being adapted to be returned to said at least one intermediate space (19) towards said at least one discharge inlet orifice (31), said at least one discharge valve (23) comprising said at least one discharge conduit (30), located in said at least one inter-vein zone (ZC) and connecting said at least one discharge inlet (31) and said at least one secondary orifice (29), the air can circulate in said at least one intermediate space (19) and being returned to 15 the secondary flow space (27).
    [0003]
    3. Intermediate case hub according to claim 1 or 2, characterized in that said at least one gasket (33) to the air and fire is made, at least in part, of silicone. 20
    [0004]
    4. Intermediate casing hub according to one of the preceding claims, characterized in that the outer annular shell (14) comprises an annular projection (37) and in that the at least one discharge duct (30) comprises a ring cavity (36), the fixing of said at least one discharge duct (30) to the outer annular shell (14) being effected by means of the annular boss (37) and the annular sieve (36), in particular by means of a screwing (34) through the annular boss (37) and the annular recess (36).
    [0005]
    Intermediate casing hub according to Claim 4, characterized in that the assembly formed by the annular boss (37) and the annular recess (36) extends around said at least one seal ( 33) to air and fire, forming a separation between said at least one seal (33) and said at least one inter-vein area (ZC). 5
    [0006]
    Intermediate casing hub according to one of claims 1 to 3, characterized in that said at least one discharge duct (30) comprises a partial annular recess (36) extending over at least two opposite edges (38a). , 38b) of said at least one discharge duct (30), the fixing of said at least one discharge duct (30) to the outer annular shroud (14) being effected by means of the partial ring recess (36), in particular by screwing through the partial ring cavity (36) to raise said composite wall (30a, 30b), said at least one seal (33) being housed between said composite wall (30a, 30b) and the outer annular shell (14). 15
    [0007]
    Intermediate casing hub according to claim 6, characterized in that the outer annular shell (14) has no annular boss.
    [0008]
    An intermediate crankcase hub according to any one of the preceding claims, characterized in that a control grid (32) is disposed at said at least one secondary orifice (29), said at least one seal (33) to air and fire being arranged all around the control gate (32), the control gate (32) being fixed to said at least one discharge duct (30), in particular by means of a screwing (35). 25
    [0009]
    An intermediate crankcase hub according to claim 4 or 5 and according to claim 8, characterized in that said at least one air and fire seal (33) is located between the control grid (32). and the assembly formed by the annular boss (37) and the annular squelch (36). 3036136 18
    [0010]
    10. Crankcase hub according to any one of the preceding claims, characterized in that said at least one discharge duct (30) comprises a composite wall (30a, 30b) of thickness (E1-E7) at least equal to 1.5 mm, in particular 2 mm.
    [0011]
    11. Intermediate casing hub according to any one of the preceding claims, characterized in that the composite wall (30a, 30b) of said at least one discharge duct (30) comprises a number of impregnated plies at least equal to 3, in particular at least 4.
    [0012]
    12. Intermediate casing (11) for an aircraft turbojet engine (12), characterized in that it comprises a hub (10) according to any one of the preceding claims. 15
    [0013]
    13. Aircraft turbojet (12), characterized in that it comprises an intermediate casing (11) according to claim 12. 5 10
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    FR3041380B1|2019-08-23|ASSEMBLY FOR AIR CIRCULATION DEVICE FOR TURBOMACHINE
    FR3072128B1|2019-11-08|INTERMEDIATE CASTER HUB DISCHARGE DUCT FOR AN AIRCRAFT TURBOJET ENGINE COMPRISING AN INTERNAL RIB
    CA2980688A1|2016-10-06|Discharge flow duct of a turbine engine comprising a vbv grating with variable setting
    FR2933127A1|2010-01-01|DEVICE FOR COLLECTING AIR IN A TURBOMACHINE
    FR3011584A1|2015-04-10|EXTENSION OF INTERMEDIATE CASING
    CA3048800A1|2018-07-05|Intermediate housing hub comprising discharge flow guiding channels formed by the discharge fins
    FR3009039A1|2015-01-30|INTERMEDIATE CASTER HUB FOR AIRCRAFT TURBOJET ENGINE COMPRISING AIR GUIDE DEFLECTORS
    FR3072127B1|2019-11-01|INTERMEDIATE CASTER HUB DISCHARGE DUCT FOR AN AIRCRAFT AIRCRAFT COMPRISING COOLING CHANNELS
    EP3545177B1|2020-12-30|Bypass turbomachine fitted with bleed system
    EP3778383A1|2021-02-17|Forward section of nacelle of an aircraft propulsion assembly comprising a thermal transition zone
    FR3023260A1|2016-01-08|PROPELLANT AIRCRAFT ASSEMBLY
    WO2015001258A1|2015-01-08|Method for repairing a panel by applying a doubler
    FR3032943A1|2016-08-26|NACELLE FOR A DOUBLE FLOW AIRCRAFT AIRCRAFT
    EP3084177B1|2018-02-07|Housing made from an organic-matrix composite material promoting the discharge of fumes
    同族专利:
    公开号 | 公开日
    US20180291841A1|2018-10-11|
    CN107624142A|2018-01-23|
    CA2985826A1|2016-11-24|
    FR3036136B1|2019-07-12|
    WO2016185119A1|2016-11-24|
    CN107624142B|2019-05-17|
    RU2017143793A3|2019-09-20|
    US10337456B2|2019-07-02|
    BR112017023022A2|2018-07-03|
    EP3295010A1|2018-03-21|
    BR112017023022B1|2022-02-01|
    RU2714388C2|2020-02-14|
    RU2017143793A|2019-06-17|
    EP3295010B1|2019-07-03|
    引用文献:
    公开号 | 申请日 | 公开日 | 申请人 | 专利标题
    EP1854989A2|2006-05-12|2007-11-14|Rohr, Inc.|Cover of a bleed air relief system in a turbomachine|
    EP2383453A2|2010-04-30|2011-11-02|General Electric Company|Flow mixing vent system|
    FR2961251A1|2010-06-15|2011-12-16|Snecma|Hub for use in intermediate casing of e.g. ducted-fan turbine engine of aircraft, has flange including air outlet orifices formed downstream from bleed valves, and deflectors permitting guiding of air between inlet and outlet orifices|
    FR3012846A1|2013-11-07|2015-05-08|Snecma|INTERMEDIATE CASTER HUB FOR AIRCRAFT TURBOJET AIRBORNE COMPRISING A DEFORMABLE CONDUIT OF AIR AND DEBRIS CANALIZATION|FR3064029A1|2017-03-15|2018-09-21|Safran Aircraft Engines|JOINT FOR SEALING LIGHT AND ASSEMBLY COMPRISING SUCH A JOINT|
    WO2019069011A1|2017-10-05|2019-04-11|Safran Aircraft Engines|Discharge duct of an intermediate housing hub for an aircraft turbojet engine comprising cooling channels|
    FR3072128A1|2017-10-10|2019-04-12|Safran Aircraft Engines|INTERMEDIATE CASTER HUB DISCHARGE DUCT FOR AN AIRCRAFT TURBOJET ENGINE COMPRISING AN INTERNAL RIB|
    FR3083832A1|2018-07-10|2020-01-17|Safran Aircraft Engines|DISCHARGE VALVE FOR TAKING FLUID CIRCULATING IN A TURBOMACHINE VEIN|
    FR3098550A1|2019-07-12|2021-01-15|Safran Aircraft Engines|GASKET FOR TURBOMACHINE INTERCASE|
    US11131406B2|2019-06-24|2021-09-28|Rolls-Royce Deutschland Ltd & Co Kg|Seal for engine firewall|US6086326A|1998-02-27|2000-07-11|United Technologies Corporation|Stator structure for a track opening of a rotary machine|
    US7624581B2|2005-12-21|2009-12-01|General Electric Company|Compact booster bleed turbofan|
    US7918081B2|2006-12-19|2011-04-05|United Technologies Corporation|Flame prevention device|
    FR2923270B1|2007-11-06|2014-01-31|Airbus France|TURBOMOTEUR WITH ADAPTED COLD FLUX TUBE|
    FR2976022B1|2011-05-31|2015-05-22|Snecma|TURBOMACHINE WITH DISCHARGE VALVES LOCATED AT THE INTERMEDIATE COVER|
    US9399951B2|2012-04-17|2016-07-26|General Electric Company|Modular louver system|
    US9322337B2|2012-06-20|2016-04-26|United Technologies Corporation|Aerodynamic intercompressor bleed ports|
    US9982598B2|2012-10-22|2018-05-29|General Electric Company|Gas turbine engine variable bleed valve for ice extraction|
    FR3027073B1|2014-10-10|2017-05-05|Snecma|TWO-PIECE ASSEMBLY COMPRISING A REMOVABLE CENTERING CUTTER FOR AIRCRAFT TURBOMACHINE|
    FR3064029B1|2017-03-15|2021-04-30|Safran Aircraft Engines|AIR-FIRE SEAL AND ASSEMBLY INCLUDING SUCH A SEAL|
    US20190055889A1|2017-08-17|2019-02-21|United Technologies Corporation|Ducted engine compressor bleed valve architecture|ES2694157T3|2015-09-22|2018-12-18|Airbus Defence And Space Sa|Aircraft bleed conduit in composite material|
    BE1024605B1|2016-09-27|2018-04-26|Safran Aero Boosters S.A.|CARTRIDGE WITH SUCTION ARMS FOR AXIAL TURBOMACHINE|
    法律状态:
    2016-05-20| PLFP| Fee payment|Year of fee payment: 2 |
    2016-11-18| PLSC| Publication of the preliminary search report|Effective date: 20161118 |
    2017-04-13| PLFP| Fee payment|Year of fee payment: 3 |
    2018-02-09| CD| Change of name or company name|Owner name: SAFRAN AIRCRAFT ENGINES, FR Effective date: 20170717 Owner name: SAFRAN, FR Effective date: 20170717 |
    2018-04-23| PLFP| Fee payment|Year of fee payment: 4 |
    2019-04-19| PLFP| Fee payment|Year of fee payment: 5 |
    2020-04-22| PLFP| Fee payment|Year of fee payment: 6 |
    2021-04-21| PLFP| Fee payment|Year of fee payment: 7 |
    优先权:
    申请号 | 申请日 | 专利标题
    FR1554378|2015-05-15|
    FR1554378A|FR3036136B1|2015-05-15|2015-05-15|INTERMEDIATE CASTER HUB FOR AIRCRAFT TURBOREACTOR COMPRISING A COMPOSITE DISCHARGE DUCT|FR1554378A| FR3036136B1|2015-05-15|2015-05-15|INTERMEDIATE CASTER HUB FOR AIRCRAFT TURBOREACTOR COMPRISING A COMPOSITE DISCHARGE DUCT|
    CN201680028039.1A| CN107624142B|2015-05-15|2016-05-13|The middle casing hub including compound escape pipe for aircraft turbine jet engine|
    RU2017143793A| RU2714388C2|2015-05-15|2016-05-13|Inner housing of intermediate housing for aircraft turbojet engine, intermediate housing, including such inner housing, and turbojet engine, including such intermediate housing|
    US15/573,970| US10337456B2|2015-05-15|2016-05-13|Intermediate casing hub for an aircraft turbojet engine including a composite outlet pipe|
    EP16731208.1A| EP3295010B1|2015-05-15|2016-05-13|Intermediate casing hub for an aircraft turbojet engine including a composite outlet pipe|
    CA2985826A| CA2985826A1|2015-05-15|2016-05-13|Intermediate casing hub for an aircraft turbojet engine including a composite outlet pipe|
    PCT/FR2016/051134| WO2016185119A1|2015-05-15|2016-05-13|Intermediate casing hub for an aircraft turbojet engine including a composite outlet pipe|
    BR112017023022-4A| BR112017023022B1|2015-05-15|2016-05-13|Intermediate crankcase hub and intermediate crankcase for aircraft turbojet as well as aircraft turbojet|
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