Abstract: This invention relates to an abradable coating system for use in axial turbine engines. When coated onto a turbine ring seal segment the coating system may allow formation of an individualized seal between turbine blade disks and the surrounding ring seal without causing excessive wear to the blade tips. The abradable coating system includes columns of an abradable material. Thus, interference between the blades and the abradable coating system causes the individual columns to break off at the base. This abrasion mechanism may reduce blade wear and spalling of the coating system when compared to conventional coatings.

Abstract: A stator vane assembly for a gas turbine engine, the stator vane including an airfoil portion formed from a single crystal material and two platforms attached to the ends of the airfoil by shear pins that fit within slots formed between the platform and airfoil. The shear pin slots extend along the contour of the airfoil where the lowest stress levels are located. The platforms include openings having the same cross sectional shape as the curved airfoil. Because of the shear pin retainers between the airfoil and the platforms, the airfoil can be formed from a single crystal material while the platforms are un-coupled from the airfoil and can be made from a different material.

Abstract: A fan casing for a gas turbine engine has a fan track radially outward of the fan blades, and the fan track has sufficient strength and stiffness that, if a blade is released, it is broken up and deflected by the fan track rather than passing through to a containment system as in known arrangements. Optionally, a weakened region in the fan track may be provided, so that the leading edge portion of the blade will penetrate the fan track and be contained within the fan casing. This is particularly suitable for fan blades in which the stiffness and compressive strength are significantly higher in the leading edge region than in the remainder of the blade; for example, hollow metal fan blades or composite fan blades having a metal leading edge cap.

Abstract: A fan casing for a gas turbine engine has a fan track liner extending only over an upstream part of the fan blades, and the local stiffness and internal shape of the casing are arranged to promote the break-up of a released fan blade while permitting the leading edge region of the blade to pass through the fan track liner and be contained by the fan casing. This arrangement is particularly suitable for fan blades in which the stiffness and compressive strength are significantly higher in the leading edge region than in the remainder of the blade; for example, hollow metal fan blades or composite fan blades having a metal leading edge cap.

Abstract: A fluid-flow machine has at least one row of blades 5 having blade ends moving relative to one of a hub 3 and a casing 1, with a gap 11 positioned therebetween. At least one groove 7 extends essentially in a circumferential direction of the machine is in an area of the gap 11 along at least part of the circumference, with the extension of the groove 7 in the circumferential direction being large as compared to the extension of the groove 7 in the meridional flow direction. A cross-sectional area of the groove 7, in meridional view of the fluid-flow machine, essentially departs from a parallelogrammic shape and, due to its contour, is inclined in an upstream direction. A centroid of the groove cross-sectional area is provided upstream of the center of the groove aperture 12 on the main flow path.

Abstract: A stage of fan aerofoils (10) lies within a fan cowl (12). The fan duct (16) is defined in part by a hard casing (14) that in turn surrounds aerofoils (10). Hard casing (14) includes wedge members (26) that fill the annular gap between ring (14) and an outer ring (20). In the event of an aerofoil (10) breaking off, the hard ring (14) and wedges (26) absorb sufficient of the kinetic energy expended by the broken aerofoil (10), as to prevent it passing through outer ring (20) on to the fan cowl (12).

Abstract: A containment case comprises a composite core, with an inner surface, and at least one ceramic layer integrated with the composite core. The at least one ceramic layer is bonded to the inner surface of the composite core with a resin. In another embodiment, a containment case comprises a composite core, with an inner surface, and a hybrid material comprising at least one ceramic material and at least one non-ceramic material. The hybrid material is disposed on the inner surface of the composite core. The hybrid material is bonded to the inner surface of the composite core with a resin. A method of fabricating a containment case includes the steps of disposing one or more layers of ceramic fabric on an inner surface of a composite core, infusing the composite core and the one or more layers of ceramic fabric with resin, and curing the resin.

Abstract: A fan track liner within a rotor blade assembly 26 for a gas turbine engine (10, FIG. 1) comprises a radially inner casing liner member 28 and a radially outer casing liner member 30 adjacent to the radially inner casing liner member 28. A septum 32 is interposed between the radially inner and radially outer casing liner members 28, 30 and bonds together the radially inner and radially outer casing liner members 28, 30. The septum 32 includes predetermined weakened regions 36 which permit penetration of the septum 32 in use by a rotor blade in the event of impact of the radially inner casing liner 28 member by a detached rotor blade. Assembly 26 otherwise capable of resisting ice shed from the blade impacting the same region 36.

Abstract: A casing for a gas turbine includes a construction providing improved structural efficiency. Improved load paths and means for transmitting loads in the engine case are disclosed.

Abstract: A clip used to attach an acoustic panel to a fan casing of a gas turbine engine is “U” shaped. One arm of the clip is provided with a spacing portion cut into the material of the arm and which may be bent to define the spacing between the outer face of the arm and a support. An engagement portion, which may also be cut into the material of the clip, may be bent to engage a recess in a tenon inserted into the receiving portion.

Abstract: A device that deploys during a blade portion failure that encompasses the engine casing is provided to prevent damage caused by the failure. A blade containment device such as a flexible interwoven structure or a net encompasses the engine casing and the blade failure path, or is embedded in the engine casing, and deploys in the event of a blade failure, such as the breaking of an engine blade. The net deploys when the blade portion has failed and has started expanding along the blade failure path, and may include a number of links attached to one another. Upon impact with a broken blade portion, thus transferring the kinetic energy of the blade portion to the net, some or all of the links may break off from each other, thus allowing the net to deploy and dissipating the kinetic energy generated by the expanding broken blade. Consequently, the blade remains contained inside the net, and blade containment is improved.

Abstract: Composite containment casings including a body having an interior and a fragment catcher integrally joined to the interior of the body of the containment casing.

Abstract: The present invention relates to a device for, in a gas turbine engine, braking a turbine comprising a rotor driving a shaft capable of rotating with respect to a stator in the event of said shaft breaking. The device is one which comprises a first braking member (110) provided with at least one abrasive element (110A) and a second braking member (120) comprising a ring-shaped element (120A) made of a material capable of being eroded by the abrasive element (110A), the two braking members being secured one of them to the rotor and the other to the stator and coming into contact with one another through axial displacement of the rotor once the shaft has broken, the abrasive element (110A) of the first braking member (110) eroding the ring-shaped element (120A) of the second braking member (120).

Abstract: A method for assembling a gas turbine engine includes coupling a low-pressure turbine to a core turbine engine, coupling a gearbox including an input, a first output, and a second output, to the low-pressure turbine, coupling a first fan assembly to the gearbox first output, and coupling a second fan assembly to the gearbox second output.

Abstract: One embodiment of a transition-to-turbine seal (300) comprises a transition exit seal (302) adapted for fixed attachment to an exit rail (350) of a transition (109), and a stage-1 seal (340). The stage-1 seal (340) comprises an axial seal slot (348) comprising parallel walls (346, 347) adapted to slidingly engage a first male member (310) of the transition exit seal (302). The stage-1 seal (340) also comprises a radial seal slot (354) comprising parallel walls (351, 353) adapted to slidingly engage a plate (338) of a row 1 vane segment (330). The transition-to-turbine seal (300) thus provides both axial and radial freedom of movement. Materials and coatings for the transition exit seal (302) and the stage-1 seal (340) are selected to promote preferential wear of these components relative to wear of the surface of the row 1 vane segment (330). Wear of the exit rail (350) is reduced or eliminated due to the noted fixed attachment.

Abstract: A blade containment of a gas turbine engine comprises a casing defining an annular wall radially spaced from and immediately surrounding a plurality of blades radially extending from a rotor, and at least one layer of metal wires wrapped around and directly attached to the annular wall.

Abstract: A gas turbine engine rotor blade containment assembly (38) comprises a generally cylindrical, or frustoconical, metal casing (40). The metal casing (40) has at least one arched portion (62,64,66) and each arched portion (62,64,66) extends circumferentially around the metal casing (40). Each arched portion (62,64,66) comprises a radially inwardly extending arch (68) in which the trough (74) of the arch (68) is radially inwardly of a straight line (76) interconnecting the axial ends (70,72) of the arch (68). The trough (74) of the arch (68) is radially inwardly of the at least one of the axial ends (70,72) of the at least one arch (68). The arched portions (62,64,66) introduce compressive stresses into the metal casing (40) during a fan blade 34 impact and enable the metal casing (40) to withstand the impact. The fan blade containment assembly (38) is lighter for large diameter turbofan gas turbine engine (10).

Abstract: A turbine engine fan casing for containing a separated blade fragment comprising a support case having an axially extending impact zone, a substantially structurally continuous penetrable covering circumscribing the impact zone; and a containment covering circumscribing the penetrable covering. The containment covering comprises a textile fabric containing threads in a warp direction and a fill direction, wherein the threads in the warp direction are of a different denier than the threads in the fill direction.

Abstract: A gas turbine engine blade containment system that includes a laminate material including first and second layers. The first layer is an inner layer within the second layer. The second layer includes a plurality of resiliently compressive springs extending transversely relative to the first layer, and each spring is encased in a crushable support material. The first layer includes a hard material to blunt a broken blade. The springs in the second layer are compressed and the crushable support material of the second layer is crushed to absorb energy of a broken blade.

Abstract: A method for assembling a gas turbine engine includes coupling a low-pressure turbine to a core turbine engine, coupling a counter-rotating fan assembly including a forward fan assembly and an axially aft fan assembly to the low-pressure turbine such that the forward fan assembly rotates in a first direction and the aft fan assembly rotates in an opposite second direction, and coupling a booster compressor to the low-pressure turbine such that the booster compressor rotates in the first direction.

Abstract: A protective ring (7) for a fan casing (2) of a gas-turbine engine equipped with a fan having especially metallic fan blades (3) includes a one-piece fiber-composite ring (8) with a ceramic cover (9) arranged on the ring inner surface facing the fan blades, and preferably formed by individual panels embedded in a ductile material, this cover (9)—upon the impingement of broken-off blade fragments—forms a crack network corresponding to an arrangement of the sintered starting ceramic particles and subsequently disintegrates into small particles, thereby absorbing a great part of the kinetic energy caused by the impinging blade fragment. Consequential damage to the protective ring is thereby minimized.

Abstract: A gas turbine fan blade containment assembly includes a fan case having an inner surface surrounding a jet engine fan and an outer surface. Mounted on the inner surface and across a blade containing region of the fan case is a load spreader layer for initially receiving a point load from a fan blade release (a “blade-out event”). A band layer is mounted to an outer surface of the fan case for carrying at least a portion of a hoop tensile load on the fan case resulting from the blade-out event, and separator film layer is mounted between the outer surface of the fan case and the band layer to retard the formation of stress concentrations in the band layer. In one embodiment, the load spreader layer includes a plurality of circumferentially-arrayed load spreader layer segments.

Abstract: A pressure relief arrangement for a pump, which includes a pump housing assembly with a pumping chamber therein, is provided in the pump housing assembly of the pump and includes a section mounted for movement between a normal operating position and a venting position, further includes a shearing element adapted to retain the section in the normal operating position, the section being mounted such that upon pressure within the pumping chamber increasing to a specified level, the pressure acts on the section and the shearing element will fail thereby permitting movement of the section from the normal operating to the venting position.

Abstract: A method for assembling a gas turbine engine including providing a core gas turbine engine at least partially defined by a frame and having a drive shaft rotatable about a longitudinal axis of the core gas turbine engine. A low-pressure turbine is coupled to a core turbine engine. A counter-rotating fan assembly is coupled to the low-pressure turbine. The counter-rotating fan assembly includes a first fan assembly and a second fan assembly. A booster compressor is coupled to the second fan assembly. A gearbox is securely coupled to the frame so that the gearbox is circumferentially positioned about the drive shaft. The first fan assembly is rotatably coupled to an input of the gearbox such that the first fan assembly rotates in a first direction. The second fan assembly is rotatably coupled to an output of the gearbox such that the second fan assembly rotates in a second direction opposite the first direction.

Abstract: A method facilitates assembling a gas turbine engine. The method comprises coupling a low-pressure turbine rotatable about a drive shaft to a counter-rotating fan assembly including a first fan assembly and a second fan assembly wherein the first fan assembly rotates in a first direction and the second fan assembly rotates in an opposite second direction. The method also comprises coupling a planetary gearbox substantially circumferentially about the drive shaft such that an input of the gearbox is coupled to the low-pressure turbine and an output of the gearbox is coupled to the counter-rotating fan assembly, and such that the gearbox is positioned within a lubrication fluid sump that substantially circumscribes the gearbox.

Abstract: A gearbox includes a support structure, at least one sun gear coupled within the support structure, and a plurality of planetary gears coupled within the support structure. The support structure includes a first portion, an axially aft second portion, and a thrust spring coupled between the first and second portions.

Abstract: A method for assembling a gas turbine engine including providing a core gas turbine engine at least partially defined by a frame. The gas turbine engine has a drive shaft that is rotatable about a longitudinal axis of the core gas turbine engine. A counter-rotating fan assembly is coupled to the low-pressure turbine. The counter-rotating fan includes a first fan assembly that rotates in a first direction and a second fan assembly that rotates in an opposite second direction. A booster compressor is coupled to the second fan assembly. A first bearing assembly is mounted between the drive shaft and the frame and operatively coupled to the first fan assembly to transmit thrust force from the first fan assembly to the frame. A second bearing assembly is mounted with respect to the second fan assembly and operatively coupled to the second fan assembly to transfer thrust force from at least one of the second fan assembly and the booster compressor to the second bearing assembly.

Abstract: Device for limiting turbine overspeed in a turbomachine, the turbine comprising a rotor formed of disks mounted on a turbine shaft and carrying moving blades, stages of fixed blades disposed between the disks of the rotor, and means for shearing the moving blades of the upstream disk of the rotor, these means being carried by a circumferential rim of a stage of fixed blades of the turbine.

Abstract: A containment system comprises at least one passage obstructer extending from a housing inlet cover through a diffuser passage and into a housing back cover. The passage obstructer includes a fastener portion, such as a bolt, and an obstructing portion extending from and integral to a shaft of the fastener portion. The diameter of the obstructing portion is less than the diameter of the shaft, allowing the obstructing portion to bend upon impact with a burst impeller fragment.

Abstract: The present invention relates to a device for, in a gas turbine engine, braking a turbine comprising a rotor, having at least one disk (63?) with a rim (63?B), driving a shaft and capable of rotating with respect to a stator, this device being for the event of said shaft breaking and comprising a first braking member (110), secured to said rim and provided with at least one cutting element (110A), and a second braking member (120) secured to the stator downstream of the rim (63?B), comprising a ring-shaped element (120A) made of a material that can be cut by the cutting element (110A), the two braking members coming into contact with one another through axial displacement of the rotor once the shaft has broken, the cutting element (110A) of the first braking member (110) cutting the ring-shaped element (120A) of the second braking member (120).

Abstract: An arrangement of bearing supports for a rotating shaft mounted on bearings connected to an engine stator structure by bearing supports of which at least one is made up of N>1 elements effectively acting in parallel to simultaneously connect the bearing to the structure, of which N?1 of the N elements form this connection by fusible links, is made up of a first bearing supported by a first bearing support and a second bearing supported by a second bearing support, that of the two bearing supports which makes up the N>1 elements being the second bearing support.

Abstract: A turbofan gas turbine engine (10) comprises a fan rotor and a compressor rotor (62). The fan rotor (24) carries a plurality of circumferentially spaced radially outwardly extending fan blades (26). The compressor rotor carries a plurality of circumferentially spaced radially outwardly extending compressor blades (64). The fan rotor (24) is mounted on a fan shaft (38) and the compressor rotor (62) is mounted on a compressor shaft (68). The compressor shaft (68) is arranged around the fan shaft (38). The fan rotor (24) is rotatably mounted on a support structure (40) by a first bearing (50) and the compres˜sor rotor (62) is rotatably mounted on the support structure (40) by a second bearing (58) spaced axially from the first bearing (50). The fan and compressor rotors (24,62) are normally arranged coaxially with each other and with the support structure (40).

Abstract: An integrally bladed rotor disk, commonly referred to as an IBR or a blisk, includes partial cuts in the rotor disk extending from the rotor disk surface and along side the blades to form a partial root portion below each blade. A damaged blade can be removed from the rotor disk by breaking off the partial root portion from the disk or additionally cutting the disk to join the two cuts forming the partial root portion of the damaged blade. A replacement blade having a root already formed on the blade can then be inserted into the remaining slot of the rotor disk and secured to the rotor disk by any known means. The partial formed root of the blade can be in the shape of a dove tail root or a fir tree root.

Abstract: The fan cowl (12) of a ducted fan gas turbine engine (10) has a downstream portion including an inner casing (30) that surrounds a honeycomb structure (36) in abutting engagement. Honeycomb structure (36) is made up of a number of closely spaced honeycomb blocks (36a) fixed about a liner (40), one or more of which blocks, if struck by a broken fan blade portion, will move relative to those not struck, in a radially outward direction, thus reducing the crushing force of the blow, and only flexing casing (30) instead of permanently deforming or puncturing it.

Abstract: Described is reinforcement of a fan case in a gas turbine jet engine. In one embodiment, a containment ring and a hear resistance ring are shrink interference fit on the inside diameter of the fan case, the containment ring where the large fan blades turn, and the heat resistance ring where heated air from backfiring heats up the fan case. In one example, the containment ring is made of a sue alloy to provide added strength to the fan case should a fan blade break, containing the fan blade within the fan case. Also, the containment ring may extent forward of at least the leading edge of each fan blade and aft of at least the trailing edge of each fan blade. The heat resistance ring is made of titanium or other suitable material. Additionally, one or more stiffener rings may be shrink interference fit on the outside diameter of the fan case.

Abstract: A pressure relief arrangement for a pump (10) which includes a pump housing assembly with a pumping chamber (25) therein, the pump housing assembly including a section (41) mounted for movement between a normal operating position and a venting position, a shearing element (45) being adapted to retain the section in the normal operating position, the section (41) being mounted so that upon the pressure within the pumping chamber (25) can act on the section (41) and upon the pressure within the pumping chamber (25) reaching a specified pressure, the shearing element (45) will fail thereby permitting movement of the section (41) from the normal operating to the venting position.

Abstract: Described is reinforcement of a fan case in a gas turbine jet engine. In one embodiment, a containment ring and a heat resistance ring are shrink interference fit on the inside diameter of the fan case, the containment ring where the large fan blades turn, and the heat resistance ring where heated air from backfiring heats up the fan case. In one example, the containment ring is made of a super alloy to provide added strength to the fan case should a fan blade break, containing the fan blade within the fan case. The heat resistance ring is made of titanium or other suitable material. Additionally, one or more stiffener rings may be shrink interference fit on the outside diameter of the fan case. The containment ring and stiffener rings can reduce the flight weight of the fan case and lower the material costs, while increasing the containment strength of the fan case. Other embodiments are described and claimed.

Abstract: A diffuser includes multiple vanes rotatable about pivot pins between multiple positions. A mounting plate supports a backing plate. The vanes are arranged between the backing plate and the shroud. The pivot pins are in a slip fit relationship with the backing plate, vanes and shroud and are threadingly received by bosses in the mounting plate. The mounting plate and pivot pin arrangement better enables the backing plate and shroud to remain parallel with one another during deflection of the backing plate and/or shroud. Structure, such as integral protrusions and/or bolts, extend from between the backing plate and shroud through apertures in the vanes to better contain the vanes in the event of a catastrophic failure.

Abstract: In a turbofan gas turbine engine, a fan shaft is rotatably mounted and radially supported by a bearing in a bearing support structure supported from a fixed engine structure by radially frangible bolts and radially extending spokes. The radially inner ends of the radially extending spokes are mounted on a common member that engages the bearing support structure. The radially outer ends of the radially extending spokes are mounted on a fixed engine structure located radially outwardly of the bearing support structure. The radially extending spokes are held in tension and include a super elastic material to exert a radially inward restoring force on the bearing support structure, subsequent to any radial excursion of at least part of the fan shaft relative to an engine rotational axis (X) following any fracture of the frangible bolts.

Abstract: A gas turbine engine rotor blade containment assembly (38) comprising a generally cylindrical, or frustoconical, casing (40). The casing (40) having at least one member (54) and the member (54) has a first portion (53) extending radially from the casing (40) and a second portion (55) extending axially from the first portion (53) of the member (54). At least one cassette (80) being provided and one or more liner panels (74) being arranged in each cassette (80). A first end (84) of each cassette (80) being locatable on the second portion (55) of the member (54) and the first end (84) of each cassette (80) being shaped to correspond to the shape of the first and second portions (53,55) of the at least one member (54) to protect the first and second portions (53,55) of the at least one member (54). The second end (86) of each cassette (80) having at least one axially extending member (96) removably securable to the casing (40).

Abstract: A composite fan casing for a gas turbine engine includes, in an exemplary embodiment, a core having a plurality of core layers of reinforcing fiber bonded together with a thermosetting polymeric resin. Each core layer includes a plurality of braided reinforcing fiber tows. The braids of reinforcing fiber tows are aligned in a circumferential direction with each fiber tow including a plurality of reinforcing fibers. The core also includes at least one additional fiber tow braided into at least one predetermined axial location to form at least one concentrated stiffening ring.

Abstract: A turbomachine comprising at least a first rotary assembly with a first shaft, a stator, and bearings secured to the stator and suitable for supporting the shaft, one of the bearings being capable of giving way when an unbalance occurs in the rotary assembly, at least one segment of the shaft being covered in a covering which comprises at least one piece of fabric impregnated with resin and which is for coming into contact with a surrounding portion of the turbomachine when the unbalance occurs in order to protect the first shaft and ensure that there is no risk of the first shaft rupturing. The turbomachine is intended to be used mainly in the field of aviation, and more particularly as an airplane turbojet.

Abstract: A device (40) for limiting overspeed in a turbomachine in the event of breakage of the turbine shaft (24), comprising means (38, 44) of propelling a pin (42) into the path of the blades (22), these propelling means being controlled by means (36) that detect overspeed or that the turbine shaft has broken.

Abstract: A turbine module for a gas turbine engine includes at least one turbine disk and a disk-shaped component mounted on the turbine disk upstream relative to the gas flow. The module can be assembled to the compressor of the engine, wherein the component is preassembled to the turbine disk, before the module is assembled to the compressor, by bolting to an attachment flange fixedly attached to the turbine disk.

Abstract: A turbomachine including at least a first rotary assembly with a first shaft, a stator, and bearings secured to the stator and suitable for supporting the first shaft, one of the bearings being capable of giving way when an unbalance occurs in the first rotary assembly. The turbomachine further comprising brake arrangement for braking the first shaft when the unbalance occurs, and bond arrangement acting at the end of braking to bond a portion of the first shaft to a surrounding portion of the turbomachine. The turbomachine is intended to be used mainly in the field of aviation, and more particularly as an airplane turbojet.

Abstract: A gas turbine engine component and a method for making a composite gas turbine engine component, the method includes providing a tool having a first and second surface, the first and second surface arranged at an angle to each other. A fibrous segment is applied to the first surface. The fibrous segment has an end terminating adjacent to the angle between the first surface and the second surface. A fibrous preform is positioned on the fibrous segment and on the second surface. A matrix material is then provided on one or more of the fibrous preform or the fibrous segment. The fibrous segment and fibrous preform are heated to a temperature sufficient to allow expansion of the end of the fibrous segment. The matrix material is then cured to form a composite article having an angular portion having substantially uniform fiber distribution.

Abstract: A fan case and method of manufacturing a fan case, for a gas turbine engine, where the fan case has a hollow tubular metal shell with: a central axis of symmetry; an inlet; an outlet; and peripheral wall about the axis to encompass the tips of a plurality of rotary fan blades, and the shell wall includes an upstream portion defining an annular abradable material recess that extends axially having an upstream end at the shell inlet and a downstream end located upstream from the shell outlet.

Abstract: A flange of a molecular pump is connected to a vacuum vessel by first bolts that extend through elongate holes in the flange and are bolted to the vacuum vessel. One or more buffering members are attached to the flange by the first bolts that extend through first openings in the buffering members and by second bolts that extend through second openings in the buffering members and are bolted to the flange. In the event of malfunction of the molecular pump that generates an abnormally high torque, the buffering members plastically deform and absorb the shock energy caused by the abnormal torque thereby preventing damage to the vacuum vessel.

Abstract: The turbomachine of the invention extends longitudinally along an axis, and includes a rotor attached to drive shaft, arranged to rotate around an axis, supported by at least a first bearing, mounted on the fixed structure of the turbomachine by a bearing support element. The turbomachine is characterised by the fact that it includes a stop ring, mounted on the fixed structure of the turbomachine, to cooperate with the support element of the first bearing and, in the event of displacement of the rotor in relation to the fixed structure, to perform a function of axial retention of the rotor in an even manner, with no angle effect between the axis of the turbomachine and the axis of the drive shaft.

Abstract: A gas turbine engine rotor blade containment assembly (38) comprises a containment casing (40) and the containment casing (40) has a plurality of axially spaced ribs (66). The ribs (66) extend circumferentially and radially inwardly from the containment casing (40) to define a plurality of circumferentially extending grooves (76) and a perforate sheet (70) is secured to the radially inner ends (68) of the ribs (66). In the arrangement of the present invention the blade containment function and the acoustic treatment function are combined.