Abstract: A turbofan gas turbine engine (10) comprises a compressor rotor (24) carrying a plurality of circumferentially spaced radially extending rotor blades (26) and a casing (28) surrounding the compressor rotor (24) and compressor rotor blades (26). A compressor rotor blade tip seal (48) comprises an annular member (30) secured to and arranged within the casing (28). The annular member (40) has a plurality of circumferentially spaced axially extending corrugations (32) and a plurality of radially spaced circumferentially extending corrugations (34). A lining (40) is secured to and arranged within the annular member (30). The lining (40) is spaced radially from the tips (27) of the compressor rotor blades (26) to form a clearance (29). An actuator (56) is provided to move the annular member (30) and lining (40) radially to control the size of the clearance (29).

Abstract: A turbofan gas turbine engine (10) comprises a fan rotor and a compressor rotor (62). The fan rotor (24) carries a plurality of circumferentially spaced radially outwardly extending fan blades (26). The compressor rotor carries a plurality of circumferentially spaced radially outwardly extending compressor blades (64). The fan rotor (24) is mounted on a fan shaft (38) and the compressor rotor (62) is mounted on a compressor shaft (68). The compressor shaft (68) is arranged around the fan shaft (38). The fan rotor (24) is rotatably mounted on a support structure (40) by a first bearing (50) and the compres˜sor rotor (62) is rotatably mounted on the support structure (40) by a second bearing (58) spaced axially from the first bearing (50). The fan and compressor rotors (24,62) are normally arranged coaxially with each other and with the support structure (40).

Abstract: A seal arrangement 30 presents a barrier in the form of a combined seal and anti-fret liner 36 across a shroud ring cavity 32 in order to inhibit air leakage in the platform gap between platforms 33. It will be understood that these gaps are provided for thermal expansion and vibration of vane assemblies but create a leakage path between passage 35 and cavity 32 which can increase aerodynamic losses and therefore reduce engine compressor efficiency. By provision of a combined seal and anti-fret liner 36 such air losses are reduced whilst combining the function of an air leakage seal and an anti-fret liner to prevent abrasive degradation of the mountings 37, 39 and platform section 33 ensures that there is a reduced component count as well as cost, weight and assembly time.

Abstract: Previously a number of techniques have been used in order to form single crystal or pre-determined crystallography components and articles. Each one of these techniques has its own particular problems, including susceptibility to error. By utilisation of a bi-crystal experiment to determine melt-back length LM and by consideration of the ingress distance d from potential initiation nucleation points on a perimeter of a seed crystal, it is possible to determine a maximum ingress length d. By ensuring that the maximum ingress length d is less than or equal to a seed crystal diameter R, it is possible to project locus from potential nucleation points C1, C2 in terms of potential radii for stray grain propagation. As the seed crystal will have a known crystalline orientation, it will be possible to consider two divergent growth curves of the crystal in terms of the stray grains propagating from the point C1, C2.

Abstract: A combustion chamber suitable for a gas turbine engine is provided with at least one Helmholtz resonator having a resonator cavity and a resonator neck in flow communication with the chamber interior. The resonator neck is provided with at least cooling holes extending through its wall for improved damping and cooling, at least one of the holes is directed towards the resonator cavity.

Abstract: An apparatus for controlling the temperature of a component, which is situated in use in a gas stream, provides a nozzle to create a jet of air at an angle to the gas stream, the jet being directed into the region of the stagnation point of the component so as to control the temperature of the component. The invention is particularly suited to preventing or reducing the formation of ice on vanes of gas turbine engines, but may also be applied to other components, and may equally be used in situations where a component is to be cooled rather than heated.

Abstract: A turbomachine including securing means that extend between the pressure wall and the suction surface and which includes an energy absorbing portion for absorbing energy after impact to the blade by a foreign object. The energy absorbing portion has a catch that provides the blade with an improved resistance to bursting.

Abstract: A turbine fuel supply system is disclosed as including a first sub-system having: a first nozzle for injecting fuel into a combustor of a turbine engine; a first valve controllable to communicate fuel to the first nozzle; and a first fuel manifold for communicating fuel to the first valve from a fuel source; the system further including a controller assembly for raising the pressure of fuel in the first fuel manifold; and the system being characterized in that the first valve is adapted to open in response to a predetermined pressure difference between the first fuel manifold and the pressure in the combustor, thereby allowing fuel to be communicated from the first fuel manifold to the first nozzle. The system may include a first recirculating conduit in fluid communication with the first fuel manifold and the fuel source, the first recirculating conduit allowing fuel not communicated by the first valve to be returned to the fuel source.

Abstract: A shaft coupling device to couple first (10) and second (12) shafts, said device comprising a fluid coupling (14) having a first portion (14a) attached to the first shaft (10) and a second portion (14b) attached to the second shaft (12) wherein the device further comprises engagement means (44) operable to switch the device between an engaged position and a disengaged position, the engaged position having each of the first and second shafts (10, 12) engaged with its respective portion (14a, 14b) of the fluid coupling and the disengaged position having at least one of the first and second shafts (10, 12) disengaged from its respective portion (14a, 14b) of the fluid coupling.

Abstract: Within an intake for a gas turbine engine provision is provided through stator vanes 25, 45 in the stator whereby back pressure from a necessary obstruction 34, 46 can be utilised to balance forward pressure variations caused by intake droop or crosswinds in order to reduce those forward pressure detriments for more efficient engine operation according to a desired objective regime. Typically the flow through the intake 20 is analysed and then an appropriate positioning of the stator vanes 25, 45 determined in order to provide approximate balance between the forward pressures and back pressures. Normally, a combination of camber variation of stator vanes 45a and stagger variation of stator vanes 45b are utilised in order to achieve a desired momentum balance around the circumference of the intake 20. It will be understood that both the forward pressures and the back pressures are differentially variable about the circumference such that one opposes the other.

Abstract: The fan cowl (12) of a ducted fan gas turbine engine (10) has a downstream portion including an inner casing (30) that surrounds a honeycomb structure (36) in abutting engagement. Honeycomb structure (36) is made up of a number of closely spaced honeycomb blocks (36a) fixed about a liner (40), one or more of which blocks, if struck by a broken fan blade portion, will move relative to those not struck, in a radially outward direction, thus reducing the crushing force of the blow, and only flexing casing (30) instead of permanently deforming or puncturing it.

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 fault current limiter comprising: an input node; an output node; a variable impedance element coupled between the input node and the output node; a multi-cycle refrigeration system, comprising a first coolant cycle for cooling the variable impedance element and a second coolant cycle, thermally coupled to the first coolant cycle, for cooling the first coolant, wherein the variable impedance element comprises Magnesium Diboride.

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 seal arrangement 32 between a turbine shroud seal segments 38,40 in a gas turbine engine. The arrangement 32 comprises facing slots 34, 36 provided respectively in the segments 38,40 with offset spaces 42, 44 provided between the segments 38,40. A strip 46 is cemented into the slot 36 such that the flow path 50 passes through the space 42, around the free end of the strip 46, and out through the offset space 44, thereby providing a longer flow path.

Abstract: A circumferential row of aerofoil members span an annular duct for carrying a flow of compressible fluid. The duct is centred on the axis of a turbomachine. Pressure and suction surfaces of neighbouring aerofoil members bound respective sectoral passages which receive the flow of compressible fluid. The row comprises at least one radial endwall to which the tangent to the trailing edge of each aerofoil member at midspan is substantially orthogonal. Each aerofoil member has reverse compound lean. Each aerofoil member further has a leading edge which has a position at the endwall which is upstream of its position at midspan.

Abstract: A compressor or fan is controlled to avoid stall, surge or flutter. The control system includes a look-up table from which is retrieved a margin-adjusted limiting value of a control parameter of the compressor, using inputs representing inlet guide vane angle (IGV), non-dimensional speed of the compressor (N/R?T), vehicle Mach number (Mn) and inlet geometry (IG). A variable stall, surge or flutter safety margin about a nominal limiting value of a control parameter is pre-established to reflect threats to the working, stall, flutter and surge lines, for example as a result of measurement inaccuracies, and from this the margin-adjusted limiting value is derived. An actual value of the control parameter is determined in real time, and the actual value is compared with the margin-adjusted limiting value. An output signal indicating stall, surge or flutter risk is generated if the actual value falls within the stall, surge or flutter safety margin.

Abstract: A marker paint comprises 42 wt % to 52 wt % pigment, 31 wt % to 37 wt % acrylic resin and 17 wt % to 21 wt % silicone resin excluding solvent. The pigment comprises cobalt titanium oxide, titanium nickel antimony oxide, cobalt aluminium oxide or cadmium sulphide selenium to give a green, yellow, blue or red colour respectively. The marker paint is used in conjunction with an irreversible temperature indicating paint to determine if there is any debris and/or dirt on the irreversible temperature indicating paint to affect the accuracy of the temperature measurement of the irreversible temperature indicating paint.

Abstract: An internal fluid cooling system for a gas turbine aerofoil comprises a plurality of multi-pass cooling arrangements each of which consists of a serpentine passage in the interior of the aerofoil. The cooling fluid—in particular air—is supplied to an inlet end of each passage and exhausted through a multiplicity of discharge holes to provide tip, leading edge, trailing edge and surface film cooling. The inlet end of a first serpentine passage is positioned close to the leading edge and flows rearwards while the inlet end of the second serpentine passage is positioned close to the trailing edge and flows forwards. These serpentine passages are disposed side-by-side, one adjacent the pressure surface and the other adjacent the suction surface on opposite sides of a main load carrying member which comprises a major part of the internal structure of the aerofoil.

Abstract: A blade assembly for a rotary component comprises an aerofoil member and a displacement apparatus on the aerofoil member for displacing a detached first portion of the aerofoil member in a rearward direction relative to a second portion of the aerofoil member. On failure of the aerofoil member, the displacement apparatus displaces the first portion from the second portion in the rearward direction.