Rolls-royce Corporation

Other

Rolls-royce Ple

Rolls-ryce Plc

Alec George Dodd

Derek A Roberts

Mark A Halliwell

Martin G Rose

Jagnandan K Bhangu

Anthony G Gale

Terence R Pellow

Martin Hamblett

Edwin Pateman

Stephen N Dancer

The present invention relates to a cooling arrangement in which a fractal passage network is provided incorporating several passage stages. Each passage stagehas substantially the same passage cross-section for consistency with other passage stages to facilitate flow pressure control across the fractal passage network. Normally, the length of passages in each passage stage are also specifically determined in order to achieve a consistent flow pressure control. The arrangement is particularly utilized with regard to cooling within turbine engine components such as blade aerofoils.

Cooling arrangement

A turbine blade () has cooling air passageways () and (.) through the leading edge wall portion () which are positionally arranged so as to intersect each other within the wall thickness so as to transmit mechanical stresses into the thicker, non-perforated material of the blade aerofoil (). Further passageways near the blade root portion () do not intersect, the reduced cooling in that area causes expansion and stress absorption.

Turbine blade

A machine for forming and fitting core support pins into wax ( ) surrounding a core ( ) prior to covering the wax ( ) with a frit ( ), includes one or more punch devices ( ), ( ) within its structure, so as to enable the forming of one or more indentations in each pin, intermediate its ends, thereby providing a designed weakness in the pin to cause it to more reliably follow slight relative movement which occurs between the pins and the part being cast, thus reducing stresses in the surface of the part.

Apparatus for performing foundry work

Tip clearance apparatus for a gas turbine engine comprises a shroud ring having curved portions so as to allow eccentric offset and hence asymmetric movement of the shroud. The shroud ring is mounted within a guide also having corresponding curved portions and movement of the shroud ring is controlled by the use of sensors.

Gas turbine engine system

A mold support ( ) is disclosed for supporting a plurality of molds ( ) in a casting apparatus having a first zone for introducing molten material into the molds and a second zone in which the material in the molds can cool. The mold support ( ) is capable of relative movement from the first zone to the second zone thereby forming a solidification front in the material in each mold, whereby the solidification front moves through the material as the material solidifies. Each mold ( ) defines a main axis (A) along which the solidification front can move and a centrally extending plane (P) perpendicular to the main axis. The mold support ( ) comprises carrying means upon which the molds ( ) can be disposed. The geometry of the carrying means is such that the centrally extending plane of a first mold ( ) arranged on the carrying means at or towards a central region thereof is not co-planar with the centrally extending plane of a second mold spaced from the first mold ( ) away from the central region, and such that the main axis (A) of each mold ( ) extends substantially perpendicular to the solidification front to be formed therein.

Mould support arrangement

Two stages of relatively rotatable guide vanes ( ) are arranged in tandem. Shroud portions ( ) on the vanes overlap in loose engagement, until an associated gas turbine engine starts operating. Gas loads then act on the downstream vane stage ( ) and move it in a downstream direction until the opposing flanks of radially opposing lands ( ) on respective shrouds ( ) engage, whereupon, gas leakage via the overlap onto the turbine casing ( ) is at least substantially reduced.

Tandem guide vanes

A hollow liner ( ) for fitting in the interior of a aerofoil vane ( ) is crimped prior to fitting within the vane ( ). Following fitting, the liner ( ) is expanded so as to fit snugly within the vane ( ). The technique permits the insertion of liners ( ) into vanes ( ) that are so configured as to preclude the insertion of conventional liners.

Method of producing a lining artefact

The shroud ( ) of a shrouded turbine blade ( ) has a convex leading edge ( ) the apex of which extends towards the trailing edge of the shroud ( ) of an adjacent, preceding guide vane. The extension enables the containment of a reverse flow of high pressure gas from the juncture of the blade aerofoil and shroud, until the pressure has dropped sufficiently to avoid nullifying gas flow from the corresponding portion of the guide vane ( ) and thereby reduces excessive hot gas leakage from between the opposing shrouds edges.

Turbine blades

A gas turbine blade ( ) which has a rounded trailing edge ( ), is provided with a row of side by side arranged ceramic fibres ( ) along the trailing edge ( ). During operation of the turbine blade ( ), the rounded shape of trailing edge ( ) causes gasflows to break from the rounded edge before reaching the edge extremity. The presence of the fibres ( ) prevent the formation of vortices in the gasflow, and thereby improve turbine efficiency.

Turbine blade and manufacture thereof

Shrouds (,) surrounding a stage of turbine blades (,) are cooled by a compressor airflow which is led to the downstream end of the shrouds before contacting them. The airflow passes through apertures (,) in plates (,) then over the shrouds (,) in an upstream direction, to exit from apertures (,) in the shrouds (,) 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.

Turbine cooling

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.

Fluid seal

A turbine (10) for a gas turbine engine includes an annular array of turbine aerofoil blades (12) which are mounted on a disc (19). Each of the aerofoil blades (12) is provided with a radially inner platform (21). Each platform (21) includes a passage (37) into which leaked cooling air flows. The passages (37) are disposed in a direction having a circumferential component so that cooling air is exhausted from them in a direction that is generally opposite to that in which the disc (19) operationally rotates.

Turbine

A variable diameter shroud ring (21) for the turbine of a gas turbine engine (10) comprises a support structure (25) which carries an annular array of circumferentially spaced apart ceramic segments (26). The radially outer surfaces of the segments (26) are overlaid by a plurality of circumferentially extending metallic sheets (38). The sheets (38) serve to inhibit gas leakage through gaps between the segments (26) as the diameter of the shroud ring (21) is varied.

Turbine shroud ring

A gas turbine engine casing includes a plurality of coaxially interconnected ring members each provided with a flange at each of their axial extents to facilitate interconnection of the ring members by bolts; each ring member is provided with an integral array of inwardly directed stator vanes and carries an abradable sealing material to cooperate with the tips of rotor blades of the engine.

Gas turbine engine casing

A cast and hot isostatically pressed gamma titanium aluminide turbine blade (26) has an aerofoil portion (38) including a concave pressure surface (44) and a convex suction surface (46). The concave pressure surface (44) has surface irregularities (48), produced by the action of the hot isostatic pressing process on voids (56) located within the cast turbine blade (26), located in that region (50) of the concave pressure surface (44) where in operation aerodynamic separation (G) occurs to minimise the aerodynamic effects of the surface irregularities (48) upon the operation of the turbine blade (26). The remaining portion of the concave pressure surface (44) and the whole of the convex suction surface (46) are substantially free of surface irregularities (48).

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Derby, United Kingdom

Alec George Dodd