Abstract: A turbine aerofoil for a gas turbine has a pressure wall and suction wall and at least one internal cavity. An aerofoil platform adjacent and generally perpendicular to the aerofoil has at least one internal cavity with a pressure wall and suction wall on respective sides of the cavity. The cavity is divided into at least two cooling chambers including a first chamber for receiving cooling air for the platform pressure wall and a second chamber for receiving cooling air for cooling the platform suction wall. The first chamber is in flow communication with the aerofoil cavity for discharge of at least part of the cooling air entering the first chamber to the aerofoil cavity.

Abstract: A gas turbine engine exhaust nozzle arrangement for the flow of exhaust gases therethrough between an upstream end and a downstream end thereof comprising a nozzle and a plurality of tabs which extend in a generally axial direction from a downstream portion of the nozzle wherein the nozzle further comprises an actuation mechanism capable of moving the tabs between a first deployed position, where the tabs interact with a gas stream to reduce exhaust noise thereof, and a second non-deployed position, where the tabs are substantially aerodynamically unobtrusive.

Abstract: A bearing hub for a gas turbine engine comprises an inner annular member defining a bearing chamber and an outer annular member extending around the inner annular member. Each annular member defines an aperture to receive a conduit which can extend to the bearing chamber. The outer annular member is configured to slidingly engage the conduit at the aperture in the outer annular member.

Abstract: A wall element for use as part of an inner wall of a gas turbine engine combustor wall structure is of cast construction and includes a plurality of cooling apertures provided therethrough and formed during the casting process. The cooling apertures may be located in positions where they could not be conventionally formed by laser drilling.

Abstract: A method of measuring temperature of a component capable of emitting thermal radiation and reflecting background radiation having the steps of: providing a pyrometer for measuring the radiation from the component, characterised by coating a part of the component with a first emissivity coating and a part of the component with a different and second emissivity coating, each with known emissivities EH and EL respectively, recording a first radiation measurement from the first emissivity coating RH and a second radiation measurement from the second emissivity coating RL, then calculating the true radiation RBlade from the component from the equation R Blade = ( R H E H - R L E H ? ( 1 - E H 1 - E L ) ) ( 1 - E L E H ? ( 1 - E H 1 - E L ) ) and relate the RBlade value to the true component temperature by calibration of the pyrometer.

Abstract: A clamp for securing an element to a structure comprises, a first part and a second part, engageable into a assembled state about the element to be secured. A flexible element is attached to the first part and the second part, which flexible element allows independent movement of the first part and second part when the clamp is not assembled. The clamp further comprises a tensioning mechanism adapted to hold the flexible element taught when the clamp is in the assembled state.

Abstract: A temperature measuring system, comprising an imaging device for focusing radiation towards a radiation detector, wherein the imaging device is a diffractive optical device.

Abstract: The ducted fan (12) of a ducted fan gas turbine engine (10) is surrounded by a compartment (26), which is packed with bristles (38). Should a blade of the fan stage (12) disintegrate during operation of engine (10), the resulting fragments, if they pass through an abradeable honeycomb structure (32), will collide with the bristles (38), which will absorb the energy exerted by the fragments, and thus prevent puncturing of the outer wall (18) of the fan duct structure.

Abstract: A boundary layer control arrangement comprises a pulse generator communicating with a surface having a fluid boundary layer thereacross. The boundary layer control arrangement further includes a fluid supply means for supplying a fluid to the surface via the pulse generator. The pulse generator is constructed such that fluid acts on the pulse generator to cause the fluid to pulse. Pulsing fluid passes from the pulse generator to the surface.

Abstract: Four actuators are instructed to move to a condition indicated by a demand line 81. Their actual conditions are indicated by four arrows. The method determines if the range 84 of actual conditions exceeds a threshold value, representing an unacceptable mismatch of actuator positions. This comparison is used to establish further demand instructions. For example, if the range is too great, and the average 82 is below the demand line 81, the devices are instructed to move to the upper extreme position at 88. If the average position 82 is above the demand line 81, the devices are instructed to move to the lower end of the range, at 90. This results in the actuator positions moving together, as they also move toward the demand line 81, reducing risks associated with actuator mismatches.

Abstract: A turbine engine arrangement incorporates an electrical machine combination which can act as a generator or motor. Thus, the combination during initial turbine engine arrangement start-up acts through this motor to drive shaft rotation. This shaft rotation creates core compressed air which is bled through a valve to an air turbine such that the electrical machine when acting as a motor or generator at low engine loads is appropriately cooled. The arrangement also includes a compressor fan which generates an airflow utilised by the air turbine in order to drive the electrical machine as a generator of electrical power when engine conditions dictate. A duct is provided to direct airflow exhausted from the air turbine to a heat exchanger arranged to exchange heat with a bypass flow from the compressor fan.

Abstract: A casing section (30) for rotary apparatus (13, 14) of a gas turbine engine (10) is described. The casing section (30) comprises a partially circumferential casing member (32) and a plurality of radially inwardly extending vanes (34) fixed on the casing member (32).

Abstract: A turbine disk (20) has a sealing plate (28) fixed to it by bayonet fixing features (36) and (40) formed in sealing plate flange (30) and disk flange (38). Abutments (42) and (44) are formed on the bayonet fixings, between which a multi part anti relative rotation locking member (46) is positioned in abutting relationship.

Abstract: A method of selecting a component of optimal design comprising automated steps of a) measuring the periphery of at least part of said component; b) producing a 2D co-ordinate dataset (20) from said measurements; c) processing the 2D co-ordinate dataset (20) so as to produce a representation of characteristics of the periphery of the component; d) assigning a value to the difference between the characteristics of the periphery of the component and characteristics of a component of optimal design; and e) conducting a parametric analysis of the differences between the characteristics of the periphery of the component and the characteristics of a component of optimal design. The method further comprises the step of using the results of the parametric analysis to determine whether the component meets a predetermined criteria for acceptability.

Abstract: A securing arrangement for securing a manifold to a casing (surrounding a rotary assembly) is disclosed the securing arrangement comprises a bracket defining a recess co-operable with a radially extending part of the casing. Securing means, provide to secure the bracket to the aforesaid radially outwardly extending part.

Abstract: A gas turbine engine (2) comprises a fan unit (4, 18, 36, 42, 44, 46) in flow relationship with an engine core and a bypass duct, of which said engine core and bypass duct (16) are in parallel flow relationship with each other. The engine core comprises a compressor (6), a combustor (8) and a turbine (10), with an inner casing (12) provided around said engine core which defines the engine core intake (32). The bypass duct (16) is defined by an outer casing (14) radially spaced apart from the fan unit (4, 18, 36, 42, 44, 46) and the inner casing (12) along at least part of the length of the gas turbine engine (2). Bypass air compression means (28) are provided such that, under substantially all engine power conditions, air at exit from the bypass duct (16) is at a greater pressure than air delivered to the engine core intake (32).

Abstract: The present invention relates to laser deposition and more particularly to use of deposition wire (5, 21) in order to create structures or components. The original deposition wire (5, 21) is reconfigured in order to increase its lateral dimension and so provide a greater overlap with a presented beam (3) from a laser (2). The presented beam (3) therefore more consistently melts the presented reconfigured deposition wire (31, 41, 51, 61, 71, 81) in order to form the desired structure or component.

Abstract: A spacer arrangement 1, 40 is provided in which wedges 5, 6; 45, 46 are arranged to have an interface 7, 47 between them such that a thermally variable member 8, 48 can cause relative motion between the wedges 5, 6; 45, 46 to adjust an arrangement lateral dimension dependent upon temperature. In such circumstances with the arrangement 1, 40 positioned within a gap 3 between a casing 14 and unison ring 13, retention of position of that unison ring 13 can be maintained. The unison ring is arranged to adjust position of variable vanes.

Abstract: An engine mount for a gas turbine comprises a first compliant structure having a range of compliant movement governed by a first stiffness value. The mount further comprises stiffening means which act upon the compliant structure through part of the range of compliant movement such that the mount has at least two modes of operation, a first mode governed only by the first stiffness value of the compliant structure, and a second mode governed by both the first stiffness value and the stiffening means.

Abstract: The present invention provides a method of operating an aircraft whereby actual acceleration of an aircraft is used as the direct measure of the capability of the aircraft to takeoff. Thus, a minimum takeoff acceleration is determined for an aircraft and then through appropriate sensing means for acceleration, actual acceleration of the aircraft is determined and compared with that predetermined minimum takeoff acceleration. If the actual aircraft acceleration exceeds the minimum takeoff acceleration then an indication such as a buzzer or signal is provided. The present method may also be used in cooperation with existing takeoff determining systems utilising engine thrust as the means of determining aircraft capability to takeoff. The authority of the various aircraft takeoff criteria can be adjusted as required. Furthermore, the present aircraft acceleration regime may be utilised with regard to adjusting existing engine thrust means for determining takeoff or cruise capability.