Abstract: An annulus filler system bridges the gap between two adjacent blades attached to a rim of the rotor disc of a gas turbine engine. The system includes an annulus filler having a lid which extends between the adjacent blades and defines an airflow surface for air being drawn through the engine. The filler also has a support body extending beneath the lid and terminating in an elongate foot which, in use, extends along a groove provided in the rim of the disc. The groove has a neck which prevents withdrawal of the foot through the neck in a radially outward direction of the disc. The system further includes a sleeve which, after installation of the filler, is slidably locatable into a gap between the foot and sides of the groove.

Abstract: A turbomachine casing assembly including a first casing element locatable radially outward of one or more rotating aerofoil elements of a turbomachine, a second casing element located radially distal to the first casing element is provided. A void is provided between a radially proximal face of the first casing element and a radially proximal face of the second casing element, at the first end of the first casing element. The first end of the first casing element is located against a corresponding first end of the second casing element by an energy absorbing element, with the first casing element; including a cantilever. The cantilever is arranged in a region extending along the first casing element from the first end; and having a pivot axis at a mid portion thereof.

Abstract: A fan outlet guide vane for a gas turbine engine comprises a generally radially-extending internal structural member and at least one surface member securable in use to the internal structural member so as to define an aerofoil. The gas turbine engine includes an annular first structure and an annular second structure, the second structure being radially outward of the first structure so as to define a duct between them, and in use the internal structural member of the vane is secured between the first structure and the second structure. In a preferred embodiment, the surface member or surface members are secured only to the internal structural member, and not to the first or second structures.

Abstract: An injection-molded thermoplastic mini-platform is adhesively bonded, in use, to the aerofoil surfaces of a fan blade in a gas turbine engine. The mini-platform is divided into a number of leaves, to increase its flexibility, and ribs beneath the leaves provide additional stiffness in the radial direction. In the event that a fan blade is released in operation, the mini-platform is designed to detach from the blade and/or to break apart, so that damage to other components of the engine is minimized.

Abstract: A fan assembly for a gas turbine, the assembly comprising: a fan surrounded by a fan casing; a resilient, radially flexible containment material wrapped around the outside of the fan casing for containing a fan blade which has breached the casing during a fan blade-off event, and a retaining cable arrangement for resisting separation of a first portion of the casing from a second portion of the casing as a result of said breach occurring between the first and second portion; wherein the cable arrangement comprises: a retaining cable on the outside of the containment material, secured between the first and second portion of the casing, the retaining cable being sufficiently radially compliant for radial deflection to accommodate radial flexing of the containment material caused by impact of said fan blade in between the first and second portion; and a locking arrangement operable to subsequently lock the retaining cable between the front and rear portions of the casing whereby separation of the front and rear

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

Abstract: An annulus filler assembly for a rotor of a turbomachine, the assembly comprising: an annulus lid (56) having a radially outwardly facing surface (58) for forming an inner wall of a flow annulus of the rotor and a radially inwardly facing surface; and a frame for supporting the annulus lid (40), the frame being mountable to a disc of the rotor such that the annulus lid (56) is spaced away from the disc, wherein the frame (40) comprises a connection portion (46) which, in use, passes through an aperture (60) in the annulus lid (56) from the radially inwardly facing surface towards the radially outwardly facing surface (58) such that at least a portion of the connection portion (46) is visible from the radially outwardly facing surface (58); the assembly further comprising a locking element (68) which locks the connection portion (46) to the annulus lid (56) via the visible portion of the connection portion (46).

Abstract: A securing arrangement 32 for securing an actuating lever 20 to a spindle 14, for example a spindle 14 of a variable inlet guide vane 10 in a gas turbine engine, comprises a fastening formation 38 rotatably co-operable, in use, with a corresponding fastening formation 30 on the spindle 14. The securing arrangement 32 is rotatable relative to the spindle 14 to a securing position in which the respective fastening formations 30, 38 co-operate so that the securing arrangement 32 secures the actuating lever 20 to the spindle 14. The securing arrangement 32 includes anti-rotation means 42, preferably in the form of anti-rotation tabs 44, for preventing rotation of the securing arrangement 32 from the securing position.

Abstract: High bypass ratio turbofan propulsion engines are designed with a two-step approach to surviving foreign object damage to the fan assembly. First, a containment system encircling the fan assembly to capture any part of a released fan blade in the event of damage occurring, and two in order to cope with the vibrations resulting from a rotating unbalanced fan assembly the support structure for the bearing of the rotating assembly is constructed with frangible parts to introduce a snubber gap to permit the rotating, unbalanced fan assembly and bearing to orbit thereby to reduce the resulting vibrations. However, in the case of very wide chord fan assemblies, the largest acceptable snubber gap is insufficient to reduce the induced vibrations.

Abstract: A fan assembly 1 for a gas turbine, the assembly 1 comprising: a fan surrounded by a fan casing 2; a resilient, flexible containment material 4 around the outside of the fan casing 2 for containing a detached fan blade which breaches the casing 2 during a fan blade off event, and a retaining cable arrangement for resisting separation of a first and second portion of the casing 10, 12 as a result of said breach, wherein the cable arrangement comprises: a continuous length of retaining cable 14 extending around the fan casing 2 on the outside of the containment material 4, the retaining cable 14 being slidably or rollably secured to each of the first and second portions 10, 12 at a plurality of positions around the fan casing 2 to form respective cable spans 14a, 14b between the casing portions 10, 12, there being sufficient overall slack in the cable 14 for accommodating radial flexing of the underlying containment material 4 upon impact of said detached fan blade.

Abstract: A guide vane arrangement 20 for a gas turbine engine (10, FIG. 1) includes a vane member 21 extending between inner and outer platforms 22, 24 which are respectively mounted on inner and outer mounting members 42, 34. One of the inner and outer platforms 22, 24 includes a resilient that abuts the respective inner or outer platform 22, 24 to permit relative movement between the inner or outer platform 22, 24 and the respective inner or outer mounting member 42, 34.

Abstract: A fan assembly for a gas turbine, the assembly comprising: a fan surrounded by a fan casing; a resilient, radially flexible containment material wrapped around the outside of the fan casing for containing a fan blade which has breached the casing during a fan blade-off event, and a retaining cable arrangement for resisting separation of a first portion of the casing from a second portion of the casing as a result of said breach occurring between the first and second portion; wherein the cable arrangement comprises: a retaining cable on the outside of the containment material, secured between the first and second portion of the casing, the retaining cable being sufficiently radially compliant for radial deflection to accommodate radial flexing of the containment material caused by impact of said fan blade inbetween the first and second portion; and a locking arrangement operable to subsequently lock the retaining cable between the front and rear portions of the casing whereby separation of the front and rear p

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 fan outlet guide vane for a gas turbine engine comprises a generally radially-extending internal structural member and at least one surface member securable in use to the internal structural member so as to define an aerofoil. The gas turbine engine includes an annular first structure and an annular second structure, the second structure being radially outward of the first structure so as to define a duct between them, and in use the internal structural member of the vane is secured between the first structure and the second structure. In a preferred embodiment, the surface member or surface members are secured only to the internal structural member, and not to the first or second structures.

Abstract: A securing arrangement 32 for securing an actuating lever 20 to a spindle 14, for example a spindle 14 of a variable inlet guide vane 10 in a gas turbine engine, comprises a fastening formation 38 rotatably co-operable, in use, with a corresponding fastening formation 30 on the spindle 14. The securing arrangement 32 is rotatable relative to the spindle 14 to a securing position in which the respective fastening formations 30, 38 co-operate so that the securing arrangement 32 secures the actuating lever 20 to the spindle 14. The securing arrangement 32 includes anti-rotation means 42, preferably in the form of anti-rotation tabs 44, for preventing rotation of the securing arrangement 32 from the securing position.

Abstract: An injection-moulded thermoplastic mini-platform (42) is adhesively bonded, in use, to the aerofoil surfaces of a fan blade (312) in a gas turbine engine. The mini-platform (42) is divided into a number of leaves (46), to increase its flexibility, and ribs (48) beneath the leaves (46) provide additional stiffness in the radial direction. In the event that a fan blade is released in operation, the mini-platform (42) is designed to detach from the blade (312) and/or to break apart, so that damage to other components of the engine is minimised.

Abstract: A seal for an annulus filler for a gas turbine engine comprises a seal body releasably attached in use to the fan annulus filler, the seal body supporting two seal members so that in use one seal member is brought into contact with, and forms a seal against, an aerofoil surface of one of the fan blades, and the other seal member is brought into contact with, and forms a seal against, an aerofoil surface of the other of the fan blades.

Abstract: An axial flow compressor (16) comprises a plurality of variable stator vanes (32) circumferentially spaced apart and extending radially. Each variable stator vane (32) is rotatably mounted on a compressor casing (28). A control ring (38) surrounds the compressor casing (28). Each variable stator vane (32) is connected to the control ring (38) by a respective one of a plurality of operating levers (40) and the control ring (38) is spaced from the compressor casing (28) by a clearance. A plurality of bimetallic strips (50) are arranged circumferentially and are positioned radially between the control ring (38) and the compressor casing (28) and the bimetallic strips (50) control the clearance between the control ring (38) and the compressor casing (28) whereby any error of the variable stator vane (32) angular position is reduced. Each bimetallic strip (50) extends radially outwardly from the compressor casing (28) towards the control ring (38).

Abstract: A vane arrangement for a compressor of a gas turbine engine comprises a plurality of circumferentially arranged vanes, a plurality of levers and a control ring. Each vane comprises a fixed upstream portion secured to a casing and a movable downstream portion pivotally mounted to the casing. The movable portion of each vane is pivotally mounted at the upstream end of the movable portion adjacent the fixed portion of the respective vane. The movable portion of each vane has a spindle extending through an aperture in the casing. Each aperture in the casing has an elongate shape in cross-section and each aperture has a bush arranged within the aperture and around the spindle to fill the elongate shaped aperture. The downstream end of the fixed portion of each vane has substantially the same shape as the upstream end of the movable portion of the respective vane.

Abstract: A variable vane arrangement (36) for a compressor (26) of a gas turbine engine (10) comprises a plurality of circumferentially arranged vanes (38), a plurality of operating levers (64) and a control ring (66). Each vane (38) comprises a fixed upstream portion (40) secured to a casing (30) of the compressor (26) and a movable downstream portion (42) pivotally mounted to the casing (30). The movable portion (42) of each vane (38) is pivotally mounted about an axis (Y) at the upstream end (52) of the movable portion (42) adjacent the fixed portion (40) of the respective vane (38). The movable portion (42) of each vane (38) has a spindle (60) arranged to extend through a respective aperture (46) in the casing (30). Each aperture (46) in the casing (30) has an elongate shape in cross-section and each aperture (46) has a bush (120) arranged within the aperture (46) and around the spindle (60) to fill the elongate shaped aperture (46).