Abstract: A blade clearance control apparatus (46) for a gas turbine engine (10) comprises a series of segments (42) which define a casing in the rotor path of high pressure shroudless turbine blades (30).The segments (42) are attached to plates (78) having surfaces co-operating with chordal lengths (74) of a number of pressure tubes (68) which are located in a frame (46). The application and release of pressure to the pressure tubes (68) causes the segments (42) to move between inner and outer positions in order to control clearance X between the tips of blades (30) and the segments (42).

Abstract: A bladed rotor (19) suitable for a gas turbine engine (10) comprises a plurality of aerofoil blades (21) mounted on the periphery of a disc (20). Each of the aerofoil blades (21) has a root part (25) which locates in the disc (20) and is constituted by two radially extending circumferentially spaced apart root portions (26,27). The circumferentially outward flanks of the root portions (26,27) are configured to define a fir-tree type blade fixing which locates in a correspondingly shaped recess (34) in the disc (20). The root part (25) configuration provides weight reduction in the rim region of the disc (20), thereby reducing stresses within the disc (20).

Abstract: A radially inwardly extending annular flange (47) is integral with an annular portion (26) of an outer shaft (23). The annular flange (47) forms an oil weir in operation. A radially outwardly extending annular sealing member (56) is mounted on an annular portion (48,50) of an inner shaft (39). The shafts (23,39) are arranged to rotate relative to each other. The outer diameter of the sealing member (56) is greater than the inner diameter of the annular flange (47). In order to introduce or remove the sealing member (56) from the hydraulic seal, past the annular flange (47), the sealing member (56) is axially split by a circumferentially extending spiral to define axially overlapping portions (56A,56B). The sealing member (56) is parted axially to form a clearance and a portion of the annular flange (47) is located in the clearance. The sealing member (56) is rotated about its axis when the slope of the spiral is parallel to the annular flange (47 ).

Abstract: A method of manufacturing an axial flow compressor stator assembly in which a plurality of aerofoil cross-section stator vanes having divergent inner and outer extents are encapsulated in a low melting point alloy so that their extents protrude from the alloy. The vanes are encapsulated in the same relationship with respect to each other which they would assume in a compressor. A moulding material is then moulded around the radially inner and outer extents to define arcuate bridging member which interconnect adjacent stator vanes. A stiffening member is bonded to the bridging piece interconnecting adjacent radially outer stator vane extents and the whole assembly heated to melt the encapsulation alloy.

Abstract: A power turbine (12) comprises alternate annular arrays of contra-rotatable turbine blades. Two arrays of turbine blades (30) and (47) have annular arrays of aerofoil propulsion blades (17,18) respectively attached thereto. Alternate arrays of turbine blades are attached to a primary drum member (55) in such a manner that below normal operating temperatures, limited relative radial movement is permitted between them. However when the normal operating temperature is reached, the relative radial movement is inhibited in such a manner that the rigidity of the power turbine (12) enhanced.

Abstract: A power turbine suitable for a gas turbine engine which a blade of one array of turbine blades has a fan propulsion blade attached to its radially outer extent. A drum member is further provided which carries at least one further annular array of turbine blades and is coaxially interposed between the arrays of turbine blades and fan propulsion blades to rotate with the same in operation. Aperatures are defined through the drum member to accommodate the attachment structure provided between the turbine blade and the fan propulsion blades such that relative radial movement is permitted between the turbine blade and the drum.

Abstract: A load transference structure between a core engine and power turbine comprises a plurality of circumferentially spaced apart radially extending spoke members, each of which is attached at its radially inner extent to first and second axially spaced apart support panels. Each of the spoke members is provided with a circumferentially extending portion positioned adjacent the panels which are adapted to cooperate with a circumferentially adjacent spoke member in such a manner that radial but not circumferential load transfer is permitted to take place between them. The load transfer structure permits load transfer to take place between the core engine casing and the power turbine.