Abstract: Shrouds (40) surrounding a stage of turbine blades (44) are cooled by a compressor airflow which is led to the downstream end of the shrouds before contacting them. The airflow passes through apertures (56) in plates (50) then over the shrouds (40) in an upstream direction, to exit from apertures (62) in the shrouds (40) in parallel with and in the same direction as the gasflow. Airflow needed is reduced relative to prior art needs, resulting in improved engine efficiency, and ejection of the air does not disturb the gasflow.

Abstract: A variable camber vane for the inlet of a gas turbine engine has a pivoting trailing section behind a fixed leading section. In order to accommodate a wide range of flow conditions the trailing section is movable through a wide angle about its upstream edge. The surrounding casing profile and the radially outer edges of the vane trailing sections are matching part-spherical surfaces. As a result of the large angle of movement the trailing section is subjected to substantial turning forces. To contain the forces the curved radially outer edges of each trailing section is formed with a broad flange which tapers in the direction of the vane trailing edge, and for aerodynamic efficiency the flange is smoothly blended into the vane aerodynamic surfaces.

Abstract: A containment casing for a gas turbine engine comprising a substantially rigid casing shell arranged in operation coaxially with an axis of rotation of the gas turbine engine. The casing extending circumferentially around an array of fan rotor blades which are arranged to rotate about the engine axis. In the region of the casing where it is predicted blade impact in the event of a fan blade failure, there is at least one reinforcing rib extending substantially radially from the casing and circumscribing the outer periphery of the casing shell. The ribs providing improved strengthening of the casing. Preferably a first rib is positioned about the casing between the operational position of the trailing edge and the mid chord point of the fan blade. A second rib is axially positioned between the operational position of the mid chord point and the leading edge of the fan blade. A third rib may also be positioned axially forward of the operational position of the leading edge of the fan blade.

Abstract: A method of applying a coating to a metallic article (10) comprises placing the metallic article within a hollow cathode (38) in a vacuum chamber (30), evacuating the vacuum chamber (30), applying a negative voltage to the hollow cathode (38) to produce a plasma and such that the material of the hollow cathode (38) is sputtered onto the metallic article (10) to produce a coating (22). A positive voltage (V1) is applied to the metallic article (10) to attract electrons from the plasma to heat the coating (22) and so inter-diffuse the elements of the metallic article (10) and the protective coating (22) and a negative voltage (V2) is applied to the metallic article (10) to attract ions from the plasma to bombard the coating (22) to minimize defects in the coating (22).

Abstract: A brush seal includes a mounting ring that carries radially inwardly directed bristles. A backing member attached to the mounting ring is positioned alongside the bristles. A chamber defined by the bristles and the backing member is pressurized with fluid to counteract the lateral forces imposed upon the bristles by a region of high fluid pressure that they are operationally adjacent. As a result of this balancing of forces, the bristles are free to move radially inward and outward so that their free ends remain in sealing engagement with an adjacent surface.

Abstract: A gas turbine engine is provided with a translatable cowl and a cascade structure which is nested within the translatable portion of the cowl. The cascade structure comprises a pair of C ducts which are exposed upon translation of a portion of the cowl. The C ducts are capable of rotation about an axis parallel to the longitudinal axis of the engine thus allowing access to the core of the engine.

Abstract: A double wall structure for a gas turbine engine has an inner wall comprising a number of tiles. The outer wall is provided with a number of apertures through which air is directed into the space between the two walls. Inclined apertures are provided in the tiles so that cooling air can pass into the combustion chamber and form a cooling film underneath the tile. Each tile is provided with a number of pedestals. The orientation of the inclined apertures is such that the axis of each aperture lies upon an unobstructed channel between the pedestals.

Abstract: A hydraulic seal (15) for providing a seal between two concentric shafts (10,12) of a gas turbine engine comprises a radially outer element (16) that defines an annular oil reservoir (21) and a radially inner element (22) having an annular fin (24) that projects into the oil reservoir (21). The radially inner seal element (22) is attached to its associated shaft (10) by frangible shear pins (26). In the event of relative axial movement between the shafts (10,12) the shear pins (26) break, thereby ensuring that such shaft relative axial movement is not unduly inhibited.

Abstract: A fan rotor assembly (30) for a gas turbine engine (10) comprises a fan rotor (32) and a plurality of radially extending fan blades (34). Each fan blade (34) has a root (64) which is retained in a corresponding groove (54) in the periphery of the hub (54) of the fan rotor (32). Each fan blade root (64) comprises four axially spaced root portions (66, 68, 70 and 72) which locate in a corresponding number of axially spaced groove portions (56, 58, 60 and 62). The flanks (94, 98) of two of the root portions (66, 70) are arranged at an angle to the axis of the fan rotor (32) and the flanks (96, 100) of the other two root portions (68, 72) are arranged at an angle in the opposite sense. The flanks (74, 76, 78 and 80) of the groove portions (56, 58, 60 and 62) are arranged at the same angle as the flanks (94, 96, 98 and 100) of the corresponding root portions (66, 68, 70 and 72). The arrangement provides improved retention of the fan blades (34) and reduces the weight of the fan blades (34) and fan rotor (32).

Abstract: An apparatus (42) for inspecting an as cast turbine blade (40), with impingement cooling passages (36) (see FIG. 2 ), comprises a vacuum pump (52) arranged to evacuate the interior of the turbine blade (40) through pipe (48) and valve (54) and a supply of steam (46) arranged to supply steam to the interior of the turbine blade (40) through the pipe (48) and a valve (50). An infrared camera (56) is arranged to view the outer surface (32) of the turbine blade (40) to detect hot spots produced on the outer surface (32) of the turbine blade (40) by jets of steam impinging on the inner surface (35) of the turbine blade (40) from the cooling passages (36). A processor (60) records the images produced by the camera (56) and analyses the images to determine if the cooling passages (36) have been formed and if they have been formed in the correct position.

Abstract: A fuel injector (32) for a gas turbine engine combustion chamber (22) comprises a fuel injector nozzle (34) and a fuel feeding member (36). A wire mesh (76) is located in a space (74) defined between a hollow outer member (72) arranged around and spaced from a hollow inner member (70) of the fuel feeding member (36) and a fuel nozzle body (54) of the fuel injector nozzle (34). The wire mesh (76) rubs against the hollow outer member (72), the hollow inner member (70) and the fuel nozzle body (54) to frictionally damp vibrations of the fuel injector (32).

Abstract: A combustion chamber assembly comprises a plurality of three stage lean burn combustion chambers (28) each of which comprises a primary combustion zone (36), a secondary combustion zone (40) and a tertiary combustion zone (44). Each of the combustion zones (36,40,44) is supplied with premixed fuel and air by respective fuel and air mixing ducts (76,78,80,92). A plurality of transition ducts (118) are provided at the downstream ends of the combustion chambers (28) to receive the exhaust gases. A damping ring (130) is connected to all of the transition ducts (118) by bolts (138) which pass through apertures (128) in flanges (126) on the transition ducts (118). The bolts (138) are biased by springs (142) such that frictional contact between the damping ring (130) and the flanges (126) damps harmful vibrations in the transition ducts (118).

Abstract: A hybrid control system executes tasks within a transaction which is executed in a given order. The order in which the tasks are executed is inversely proportional to their deadlines. The deadlines are assigned an initial deadline D, an initial maximum end-to-end delay of the transaction is calculated using the initial deadlines D and the given order of the tasks.The deadline of the task with the largest deadline is reduced by an increment to give a new deadline and if the task following it has an increment to give a new deadline. A new end-to-end delay is calculated using the new deadlines and the given order of the tasks. These steps are repeated until the new end-to-end delay does not exceed the desired maximum end-to-end delay.

Abstract: An unshrouded turbine rotor blade 24 for use particularly in gas turbine engines comprising an aerofoil 30 having a leading edge 32 and a trailing edge 34. The radially outer extremity of the aerofoil 30 having a passage 42 defined by a peripheral wall 44. An aperture is formed within the wall 44 in the proximity of the trailing edge 32 of the aerofoil portion 30. The walled passage 42 is provided to capture and retain air or gas flowing over the tip of the aerofoil 30 and redirect the flow through the aperture 46 at the trailing edge of the aerofoil 30.

Abstract: A new nickel base superalloy suitable for compressor or turbine discs of gas turbine engines with fatigue crack propagation resistance equal to Waspaloy, tensile strength higher than Waspaloy and higher operating temperature than Waspaloy or UDIMET 720 family of alloys. The nickel base superalloy has a preferred composition by weight % of 14.0-19.0% cobalt, 14.35-15.15 Chromium, 4.25-5.25 Molybdenum, 1.35-2.15 tantalum, 3.45-4.15 titanium, 2.85-3.15 aluminium, 0.01-0.025 boron, 0.012-0.033 carbon, 0.05-0.07 zirconium, 0.5-1.0 hafnium, up to 1.0 rhenium, up to 2.0 tungsten, less than 0.5 niobium, up to 0.1 yttrium, up to 0.1 vanadium, up to 1.0 iron, up to 0.2 silicon, up to 0.15 manganese and balance nickel plus incidental impurities.

Abstract: A non-obstructive fluid diagnostic technique is described which enables the measurement of different parameters in a fluid. The diagnostic technique includes the steps of focusing a laser (10) in a fluid volume (14) to generate a laser induced breakdown spark (15). The characteristics of the initial laser induced breakdown spark (15) are measured and compared to the characteristics of a delayed image of the spark, formed by light emission due to recombination and excited molecules and ions generated by the laser induced spark (15). The differences between the initial and delayed images being used to diagnose characteristics of the fluid, such as velocity, temperature and pressure.

Abstract: A fluid seal comprises a plurality of seal segments (24) which form a peripheral ring at the tip of turbine rotor blades (20). The upstream end (25) of each seal segment (24) is mounted on the static structure (27). The downstream end (26) of each seal segment (24) is supported by an inner flange (34) on a vane (32). The vane (32) is also provided with a radially outer flange (36) which locates in a radially inclined slot (31) in the casing (30). During engine transients, the vane (32) expands axially forward relative to the casing (30). The outer flange (36) rides up the inclined slot (31) which causes the seal segment (24) to move radially outward and increases the seal clearance. However, during stabilized engine running, the relative axial movement between the casing (30) and the vane (32) is reduced. The outer flange (36) contracts down the inclined slot (31) to move the seal segment (24) radially inwards; the combination reduces the tip clearance and prevents excessive gas leakage.

Abstract: A wax pattern assembly (10) is removed from a ceramic shell mould (30) by removing a support member (20) from the wax pattern assembly (10). The ceramic shell mould (30) is then heated to melt the wax out of the ceramic shell mould (30). The removal of the support member (20) provides a cavity (22) into which the wax may expand without cracking the ceramic shell mould (30). The support member (20) is provided with an external thread to allow it to be easily removed from the wax pattern assembly (10).

Abstract: An improved aluminide coating especially for superalloy substrates. Slurry coating compositions of eutectic metal alloy powders and of non-eutectic metal powders, including an elemental powder of a silicide-former in a heat-curable liquid binder. A process for coating the substrates and the coated metal parts. The coatings have improved resistance to developing cracks and to hot corrosion and oxidation.

Abstract: The material removal rate of a creep-feed grinding operation may be increased significantly by use of a type of porous grinding wheel in combination with jet of coolant liquid at high pressure directed at the periphery of the wheel in advance of the cutting point. The apparatus for performing the method may comprise a multi-axis machining centre adapted to enable the coolant nozzle(s) to be retracted to provide working clearance for automatic tool changer operation and to re-position the aiming point of the nozzle(s) over an angular range relative to a cutting point.