Abstract: A flow reverser comprising a sleeve having a first part and a second part to define a flow path and axially separable from each other about a conjunction formed by a respective profile edges for the first part and the second part, the profile edges in a stowed position overlapping, the reverser characterised in that in a deployed position a nozzle part of each profile edge define together a nozzle jet and a reverser part of one edge profile is adjacent to a core to provide an effective flow deflector.

Abstract: An engine mounting arrangement is provided in which a mounting structure is created through fins (21, 42) extended between a core (22, 46) and a mounting ring (23, 48). The fins (21, 42) have a tangential lean and substantive planar aspect in order to create an appropriate load transfer path for torque and thrust loads in an engine when secured upon a hanger mounting (20) and through utilizing further struts (25, 26) extending from spaced locations upon the mounting structure to rearward anchor positions on a pylon (27) which also has the hanger mounting (20) attached.

Abstract: A gas turbine engine fan casing duct wall comprises an intake section and a containment casing, provided respectively with flanges. An acoustic flutter damper is secured between the flanges. The acoustic flutter damper has a skin accommodating an internal structure that dampens fan blade flutter. The skin is secured to the flanges at separate locations. In normal operation of the engine, the internal structure is sufficiently robust to support loads transmitted through the acoustic flutter damper between the intake section and the containment casing. If a blade or blade fragment detaches, the resulting deflection wave in the containment case ruptures the internal structure, and the load path between the intake section and the containment casing passes along the skin, which consequently maintains the connection between the intake duct and the containment casing, while permitting substantial radial deflection of the containment casing relative to the intake section.

Abstract: A duct wall of a fan casing of a gas turbine engine comprises an intake section and a containment casing, which are interconnected by bolts at flanges. An acoustic flutter damper is provided between the flanges to reduce or eliminate flutter arising in blades of the fan at certain important operating conditions. The damper provides flexibility at the connection between the intake section and the containment casing so that, in the event of detachment of a blade or a bladed fragment, the resulting deflection wave in the containment casing can be accommodated by displacement and/or deformation of the acoustic flutter damper, reducing the risk that the bolts will shear to allow the intake section and the containment casing to become detached from each other. The acoustic flutter damper may comprise a circumferential array of separate segments.

Abstract: A control mechanism is provided for moving at least two components of a gas turbine engine. The control mechanism comprises a moveable actuation rod. The control mechanism further comprises a first linkage arrangement which includes a first bell crank and which operatively connects the actuation rod to a first component of the gas turbine engine. Movement of the actuation rod produces an output motion of the first bell crank which in turn drives movement of the first component. The control mechanism further comprises a second linkage arrangement which includes a second bell crank and which operatively connects the actuation rod to a second component of the gas turbine engine. Movement of the actuation rod produces an output motion of the second bell crank which in turn drives movement of the second component.

Abstract: A containment system for a gas turbine engine comprises a force absorbing arrangement for absorbing the force exerted thereon by an ejected portion of a failed component of the engine. The force absorbing arrangement circumferentially surrounds a major proportion of the axial length of the engine.

Abstract: A containment system for a gas turbine engine comprises a force absorbing arrangement for absorbing the force exerted thereon by an ejected portion of a failed component of the engine. The force absorbing arrangement circumferentially surrounds a major proportion of the axial length of the engine.

Abstract: An engine mounting assembly for an engine, e.g. a gas turbine engine comprises a plurality of carrying arrangements. Each carrying arrangement comprises a rear load absorbing assembly. Each load absorbing assembly provides a rear mounting point for an engine carrier. Each carrying arrangement further comprises first and second stabilising members extending forwardly from the rear mounting point of the rear load absorbing assembly to first and second spaced forward mounting points for an engine carrier.

Abstract: A load transfer arrangement is provided to act as a stabiliser for utilisation in centring rotational instabilities which may occur in rotor devices such as those within a gas turbine engine. By providing a fixed screw thread path which comprises a groove having channels and having constrictions and which define clearance crowns an acceptable eccentricity range it is possible to slow screw thread driving motion and therefore load transfer only to the positions where the rotational eccentricity enters tapers leading to the constrictions. In such circumstances an orbit thread is then progressively brought into confinement and continuous engagement with a fixed screw thread defined by crowns of channels of the fixed screw thread path defined by the continuous groove through the channels. In such circumstances wide eccentricity is allowed initially but as rotational speed reduces greater and more continuous driving motions are provided whilst avoiding excessive loading during early stage operation.

Abstract: In order to accommodate loading in a number of situations such as with regard to thrusts in a gas turbine engine it is known to provide thrust bearings. These thrust bearings can be large and may be difficult to accommodate. By providing a duplex or multiplex thrust bearing which may be of increased axial length but reduced radial width easier accommodation may be achieved. By providing a thrust element 27 which compromises a balance beam from which extend elastic hinges to a mounting to a fixed bearing frame and ends of bearing races 21, 22 an appropriate thrust response can be achieved dependent upon bearing loads presented through bearing elements 23, 24. By rotation of the balance beam and flexing of the elastic hinges 30, 31, 32 an appropriate bearing load share is achieved.

Abstract: A static structure is provided such as that used as an engine structure 30 within a gas turbine engine. A structure 30 comprises a hub or core 33 with a concentric ring or casing 32. The hub 33 and casing 32 are secured together through spoke or vane elements 31 which are sloped axially and tangentially relative to a principal axis X-X. In such circumstances any roll torque reaction caused by differentially rotating hub and ring motions is retained by the tangential inclination of the spokes or vanes whilst axial sloping of these spokes or vanes 31 and axially sloping of the struts 34 creates bracing in the direction of the principal axis X-X. In such circumstances essentially robust triangles are formed radially and longitudinally respectively each comprising spokes or vanes 31 and struts 34 secured at axially spaced locations on the hub or core 33.

Abstract: A gas turbine engine (10) comprising a rotating member (16), a support structure (22) and a centering device (30), the centering device (30) centres the rotation of the rotating member (16) which is supported within the support structure (22), characterised in that the device includes a first member (34) and a second member (36) that contact one another, one of the first and second members (34, 36) extending from the rotating member and the other of the first and second members (34, 36) extending from the support structure (22), the first member (34) comprises a biasing surface (46) and the second member (36) comprises a seating surface (40), the device requiring a predetermined force to be applied to move between a centered condition and a first out of centre condition and in the centered condition the first member (34) is seated against the seating surface (40) of the second member (36), and in the first out of centre condition the first member contacts the biasing surface (46) which is arranged to provide

Abstract: A rotor assembly for a gas turbine engine compressor includes a rotor disc for mounting a plurality of radially extending rotor blades for rotation thereon, an annulus filler for securing to the disc to provide a radially inner airflow surface for air passing through the rotor assembly and a securing arrangement for securing the annulus filler to the rotor disc, the annulus filler being configured such that it is biased into a condition in which the securing arrangement mechanically retains the annulus filler on the disc, but is movable against its bias to allow the annulus filler to be removed from the disc.

Abstract: A gas turbine engine comprising a core engine surrounded by a core casing, at least one C-duct having radially inner and outer walls and a thrust reverser unit; the engine is mounted to an aircraft pylon via at least one thrust member attached to the core engine; the thrust reverser unit is mounted within the outer wall of the C-duct; the C-duct is secured to the core casing via a vee-groove attachment comprising vee-groove and vee-blade parts, the attachment is arranged to transfer loads between the thrust reverser unit and the pylon; wherein one cooperating part of the attachment is mounted to a torsion box mounted on the inner wall of the C-duct.

Abstract: A gas turbine engine comprising a core engine surrounded by a core casing, at least one C-duct having radially inner and outer walls and a thrust reverser unit; the engine is mounted to an aircraft pylon via at least one thrust member attached to the core engine; the thrust reverser unit is mounted within the outer wall of the C-duct; the C-duct is secured to the core casing via a vee-groove attachment comprising vee-groove and vee-blade parts, the attachment is arranged to transfer loads between the thrust reverser unit and the pylon; wherein one cooperating part of the attachment is mounted to a torsion box mounted on the inner wall of the C-duct.

Abstract: A hollow aerofoil (30) comprises an aerofoil portion (40) having a leading edge (42), a trailing edge (44), a concave pressure surface wall (46) extending from the leading edge (42) to the trailing edge (44) and a convex suction surface (48) extending from the leading edge (42) to the trailing edge (44). The concave pressure surface wall (46) and the convex suction surface wall (48) are integral and define a cavity (50). A plurality of webs (52) extend across the cavity (50) between the concave pressure surface wall (46) and the convex suction surface wall (48). At least one of the webs (52A) extends substantially perpendicularly to the concave pressure surface wall (46) and the convex suction surface wall (48) and at least one of the webs (52B) extends substantially diagonally to the concave pressure surface wall (46) and the convex suction surface wall (48).

Abstract: Some aircraft configurations have an engine arrangement comprising engines as part of an aft fuselage. In order to accommodate such engine arrangement positions, wings are rearwardly displaced compared to other aircraft configurations for balance across the fuselage. By creating empennage functions utilising the nacelle of engines as well as flaps to create rudder and elevator functions, it is possible to accommodate larger engine sizes more suitable for noise control with a reduced necessity for designed rearward movement of wings.

Abstract: An engine mounting arrangement is provided in which a mounting structure is created through fins (21, 42) extended between a core (22, 46) and a mounting ring (23, 48). The fins (21, 42) have a tangential lean and substantive planar aspect in order to create an appropriate load transfer path for torque and thrust loads in an engine when secured upon a hanger mounting (20) and through utilizing further struts (25, 26) extending from spaced locations upon the mounting structure to rearward anchor positions on a pylon (27) which also has the hanger mounting (20) attached.

Abstract: A mounting arrangement of a gas turbine engine comprising a core engine, surrounded by a nacelle having a bypass duct, the mounting arrangement comprising a front mount, a rear mount and characterised by the mounting arrangement further comprising a fail-safe mount axially spaced from the front and rear mounts. The fail-safe mount comprises a number of A-frames connecting the core engine and the bypass duct. In normal operation the fail-safe mount does not transmit loads between the engine and a pylon, except in the event of damage to either the front or rear mounts.

Abstract: A rotor stage (13, 14, 15, 17, 18, 19) of a gas turbine engine (10) comprises an annular retaining plate (38) capable of preventing axial movement of blades (30). The retaining plate (38) is secured to the disc (32) via a bayonet arrangement (49). A locking assembly (48) is provided to prevent relative rotation between the disc (32) and the retaining plate (38). The locking assembly (48) comprises a locking plug (50) configured in a generally Y-shaped cross section having a channel portion defined by arms (56), which engage upstream and downstream of the retaining plate (38) and a leg part (58) configured to span between bayonet parts (40 and 42), thereby preventing relative rotation between the disc (32) and the retaining plate (38). The assembly (48) further comprises a securing plate (52) configured to span between circumferentially adjacent castellations (40) thereby preventing the locking plug (50) disengaging the disc (32) and retaining plate (38).