Abstract: A fan turbine engine arrangement 10 is provided in which a core engine 1 is aligned such that an input 11 of a nacelle 2 is aligned with the upwash of a wing 4 to which the arrangement 10 is secured. The engine 1 drives a fan 14 such that bypass flows 16 produced by the fan 14 are guided by a duct 8 to present a downward component for uplift. The core engine 1 presents its output flow through a nozzle 7 which is aligned with the fundamental axis Y-Y of the engine 1 in such circumstances, the bypass flows cross the core engine 1 flow from the nozzle 7 unless that engine is adjustable during operation in order to maintain alignment dependent upon lift requirements.

Abstract: A compressor blade (30) comprises a vibration damping coating (54) on a first surface of at least one portion of an erosion resistant material (56). The vibration damping coating (54) comprises a plurality of segments (58). The portion of erosion resistant material (56) and the vibration damping coating (54) are adhesively bonded to the compressor blade (30) such that the vibration damping coating (54) is arranged between the surface (50) of the compressor blade (30) and the portion of erosion resistant material (56).

Abstract: A permanent magnet electrical machine (10) has a rotor (12) and an armature (18). Armature coils (34) are received in pairs of slots (30). The walls of the slots (30) are parallel to themselves and to the walls of the other slot (30) of the same slot pair. Thus, parallel sided locator teeth (32) exist between the slots (30) of a slot pair, and tapering spacer teeth (44) exist between adjacent slot pairs. The spacer teeth (44) are narrower than the locator teeth (32) to provide a greater magnetic flux linkage with the locator teeth, and thus increasing the emf within the corresponding coil (34).

Abstract: A leakage drain (30) is used to provide an indication that leakage of a fluid has occurred within an engine. Previously such leakage was detected by visual inspection of a retained drop at the outlet of a drain. With the present invention a sensor arrangement comprising electrodes (33, 34; 133, 134) is located within or about the leakage drain (30) whereby leakage fluid creates a film across those electrodes. Relative presence or absence or thickness of film causes variation in a detected electrical response (capacitance, reactance, resistance or impedance) and this variation is utilised in order to determine whether fluid leakage has occurred. Additionally by time of flight analysis using electrostatic sensors (35, 36; 135, 136) it is possible to determine fluid flow rate. By analysis at different incident frequencies a characteristic response for a particular type of leakage fluid is obtained. Thus, it is possible to deduce the type of leakage fluid as well as the rate of leakage.

Abstract: In order to more effectively utilise seed crystals 35, 55 to achieve a single crystal grain orientation for a component without the problems of utilising a helix constriction previously necessary to avoid epitaxial grain competition and growth. The present invention creates a wax component pattern 30. This pattern 30 comprises integral sections of wax for a mould component section and for a spacer section with the seed crystal 35 or holder for that crystal therebetween. This pattern 30 is then utilised in order to form a final refractory mould within which the component is formed. By appropriate choice of the spacer section 32, an appropriate spacing between an upper surface of the seed 55 which will be the initial interface with molten castable material to form the component and a chiller surface through which heat is transferred can be determined in order to achieve successful transfer of the seed 55 orientation to the formed component.

Abstract: An aircraft (10) comprises a wing (12) having an upper surface (16) and a lower surface (18) and at least one turbofan gas turbine engine (20) mounted on the wing (12). The axis (34) of the at least one turbofan gas turbine engine (20) is arranged substantially in the plane (14) of the wing (12) of the aircraft (10). The at least one turbofan gas turbine engine (20) has an intake, the intake comprising a hollow member (44) rotatably mounted coaxially with the at least one turbofan gas turbine engine (20) such that in use the hollow member (44) is rotatable between a first position, a low-speed condition, in which air flowing over and/or above the upper surface (16) of the wing (12) flows into the intake of the at least one turbofan gas turbine engine (20) and a second position, a high-speed condition, in which air flowing along and/or below the lower surface (18) of the wing flows into the intake of the at least one turbofan gas turbine engine.

Abstract: A mounting arrangement 31 incorporates a blade pocket cavity 35, 135, 235 between a blade platform 34, 134, 234 and a rotor disc post 53. Coolant air is presented within the cavity 35, 135, 235 in order to provide cooling of the platform 34, 134, 234. In order to inhibit ingress of hot air from adjacent wheel spaces 36, an air curtain is created across an open end 45, 145 whereby coolant air is contained within the cavity 35, 135, 235. Air flow from the air curtain itself may pass along a surface 48, 148 to provide more direct cooling for the platform 34, 134, 234. Apertures 49, 50 may be provided in order to bleed coolant air from the cavity 35 and also draw an air flow 51 across the surface 48 for greater cooling efficiency.

Abstract: A gas turbine engine nose cone (44) comprises a spinner (46) and a fairing (48). The spinner (46) has a conical upstream portion (50) and a cylindrical base portion (52). The fairing (48) is frustoconical and surrounds the base portion (52) of the spinner (46). The outer surface (68) of the fairing (48) forms a continuation of the outer surface (56) of the conical upstream portion (50) of the spinner (46). A fibrous material (86) extends around the upstream end (82) of the fairing (48). The outer surface (68) of the fairing (48) has a skin (80) to protect the fairing (48) from erosion. The skin (80) extends around the upstream end (82) of the fairing (48) and is infiltrated into the fibrous material (86) to form a composite material and to bond the skin (80) to the upstream end (82) of the fairing (48).

Abstract: A sealing arrangement (36) for sealing a leakage gap between relatively moveable parts in a flow path between a region of high pressure (A) and a region of low pressure (B). The sealing arrangement comprises an annular sealing land (52) having a radially inner surface (54) and a radially outer surface (56) projecting from one of the said moveable parts and at least two brush seal elements (40,42). The brush seal elements (40,42) comprise a mass of bristles arranged in an annular shape, one end of the bristles being fixed (48), the other being free (46). The bristles are packed together in a layer and positioned one side of a backing member (44). At least one brush seal element is provided on the radially inner surface (54) of the sealing land and at least one brush seal element is provided on the radially outer surface (56) of the sealing land (52).

Abstract: A turbine blade arrangement 100 comprises turbine blades 101, 102 which are secured in adjacent positions to a rotor disc 106 with a cavity 107 defined between root segments and platform segments 103, 104. Within the cavity 107 a flow deflector 112 is provided normally as an insert such that through a recessed portion 113 coolant flow B from a coolant path 111 is constrained to remain adjacent to a rim surface 105. By constraining the coolant flow B to remain adjacent to the surface 105 greater cooling efficiency is achieved. Inner surfaces of the deflector 102 may also be coated with low emissivity materials to reduce radiant heat flux transfer. Typically the flow deflector 112 supports a damper member 109 in association with the platform segments 103, 104.

Abstract: A Method and Machine for Shaft Imbalance Determination Correcting unbalance with regard to rotors (11, 101) and shafts is important in order to improve performance of those rotors (11, 101) or shafts within a practical machine or engine. Previously, result orientated iterative balancing of the rotor or shaft has been achieved in which mass has been added or possibly intrusive and distorted adjustments in drag and stiffness presented through bearings or alteration of the rotor (11, 101) itself have been utilised to improve rotor balance performance. In the present invention, a dynamic bearing 13 is utilised in order to provide force components with some components in phase with rotor deflection in order to simulate negative stiffness and/or some force components in phase with rotor velocity in order to simulate negative drag upon the rotor (11, 101).

Abstract: An aircraft gas turbine engine including a first hydraulic pump connected to a starter motor and a second hydraulic pump connected to the high-pressure compressor. The engine being arranged such that during starting the starter motor drives the first pump, to pump hydraulic fluid to operate the second pump as a motor to drive the high pressure compressor. During running the first and second pumps are connected to the aircraft hydraulic system.

Abstract: A hot forming die (10) has forming surfaces (16,18) and the forming surfaces have a nickel oxide layer (20). A stop off material coating (22) builds up on the nickel oxide layer (20) in operation. A method of cleaning the forming surfaces (16,18) of the hot forming die (10) includes washing the forming surfaces (16,18) of the hot forming die (10) in water to remove the stop off material (22) without removing the nickel oxide layer (20) from the forming surfaces (16,18). The forming surfaces (16,18) may be soaked in water (26) and/or a jet (30) of pressurized water may be directed onto the forming surfaces (16,18). The stop off material is quickly removed without damaging the forming surfaces (16,18). The retention of the nickel oxide layer (20) improves the quality of the hot-formed articles and the interval between cleaning of the hot forming die (10) is increased.

Abstract: A mounting arrangement (46) for mounting a turbofan gas turbine engine (10) on an aircraft pylon (44). The gas turbine engine (10) comprises a core engine (24) having a core engine casing (26). The mounting (46) comprises a first mounting (48) for mounting the core engine casing (26) on the pylon (44) and a second mounting (50) for mounting the core engine casing (26) on the pylon (44). The first mounting (48) comprises a first hinge adjacent (52) the core engine casing (26) and a second hinge adjacent (54) adjacent the pylon (44). The first hinge (52) is arranged parallel to the axis (S) of the gas turbine engine (10) to form a roll hinge. The second hinge (54) is arranged in a plane perpendicular to the axis (S) of the gas turbine engine (10) to form a pitch hinge. The second mounting (50) comprises a third hinge (56) adjacent the core engine casing (26) and a fourth hinge (58) adjacent the pylon (44).

Abstract: A fan or compressor blisk for a gas turbine engine, lift fan or the like, comprises a disc having a radially outer rim, and a plurality of aerofoil blades circumferentially spaced around and extending radially outwards from the rim. The rim has a radially inward facing surface on its underside on at least one side of the disc, the radially inward facing surface having at least one indentation for reducing the cross-sectional area of the rim, in the circumferential direction of the disc, in the region of the indentation.

Abstract: A nozzle guide vane for a gas turbine comprises at least one aerofoil having a leading edge section provided with at least one internal cavity for conveying cooling fluid through the said leading edge section, and a plate like wall section extending from the leading edge section to provide a continuous surface between the aerofoil leading edge section and the aerofoil trailing edge on the suction side of the aerofoil.

Abstract: Apparatus for maintaining the temperature of a component of a gas turbine engine below a predetermined maximum working temperature comprises a reservoir for a cooling fluid having a boiling point below the working temperature and in which the component is immersed or with which it is in contact. At least two heat exchangers are associated with the reservoir and operable to effect condensation of vaporized cooling fluid and return of same to the reservoir. Preferably the apparatus is a closed system incorporating three heat exchangers respectively adapted to effect heat exchange with engine fuel, compressed air derived from a fan or low pressure compressor of the engine and ambient air. Means may advantageously be provided to ensure return of condensed cooling fluid to the reservoir when the attitude of the reservoir is altered as a result of aircraft maneuvers.

Abstract: A gas turbine engine (10) comprises a fan rotor (26) having a plurality of fan blades (24) and an apparatus (34) for detecting an impact on the fan blades (24). The apparatus (34) comprises speed sensors (48,48B) arranged to measure the speed of rotation of the fan rotor (26) and the turbine rotor (58) and to produce speed signals. A processor unit (40) analyses the speed signals to determine the angle of twist between the ends of the fan rotor (26). The processor unit (40) is arranged to analyse the angle of twist to determine if the angle of twist of the fan rotor (40) oscillates by more than a predetermined level from an average level. The processor unit (40) is arranged to produce a signal indicative of an impact to at least one of the fan blades (24) if the angle of twist oscillates by more than the predetermined level and sends a signal to an indicator device (44,46).

Abstract: A gas turbine engine flanged shaft (120) is provided with a plurality of anti-score plates (6a, 6b) each of which has two openings which are aligned with corresponding apertures on the flange. Bolts (4) pass through the openings and apertures to attach the anti-score plates (6a, 6b) to the flange (2). The anti-score plates (6a, 6b) have different masses to facilitate shaft balancing.

Abstract: A rotor blade, a rotor disc (24) or load bearing assembly having at least one bearing surface (22) for contact with a corresponding bearing surface (28). At least one selected area of the bearing surface (30,32) which, in operation, is an area of alternating stress greater than about 50 MPa (peak to peak), is configured to have a co-efficient of friction lower than the remainder of the bearing surface (34). The selected area (30,32) having a co-efficient of friction lower than the remainder of the bearing surface (34) may be provided by the application of a dry film lubricant, with the remaining area(s) (34) of said bearing surfaces (22) being substantially free of the said coating.