Rolls-royce Corporation

Crispin David Bolgar

Guy David Snowsill

Michael J Sheath

Jonathan M King

Steven Aleksy Radomski

Geoffrey M Dailey

Peter Kevin Beardsley

Johnathan H Wilshaw

Michael Richard Hancox

Benjamin A Leishman

Peter Ellison Heaton

Jonathan Michael King

Geoffrey Matthew Dailey

Michael John Sheath

The present disclosure relates to a rotor assembly for a gas turbine engine, the rotor assembly comprising a first rotor stage having a first disc portion with a peripheral first rim portion and a second rotor stage, the second rotor stage having a second disc portion with a peripheral second rim portion. The second rotor stage is axially adjacent and downstream of the first rotor stage and the second rim portion has an axial extension extending towards the first rim portion such that the axial extension of the second rim portion defines a rotor drum cavity between the first and second disc portions. The second rotor stage further comprises a drive arm extending within the drum cavity to the first disc portion, the drive arm being connected to the first disc portion by at least one connector. The drive arm divides the drum cavity into radially outer rim cavity portion and a radially inner main cavity portion. The rotor assembly further comprises a rim seal located between the axial extension of the second rim portion and the first rim portion, and a pressure equalisation path extending from the rim cavity portion to the main cavity portion.

Rotor assembly

A vane for a gas turbine engine, the vane including a platform with an airfoil extending radially from the upper surface of the platform. The platform includes a joint portion which includes a circumferentially extending flange and a recessed surface both formed on either the upper or lower surface of the platform. The flange and the recessed surface extend from opposing circumferential edges of the joint portion and each include a substantially radially-extending through hole.

Vane joint

A casing assembly for a gas turbine engine includes a plurality of vane casing segments. Each vane casing segment includes an arcuate member and a pair of split-line flanges. Each split-line flange includes a first axial flange end and a second axial flange end. Each split-line flange is fixedly coupled to an adjacent split-line flange of an adjacent vane casing segment. Each vane casing segment includes at least one row of stator vanes welded to the arcuate member. Each split-line flange is at least partially and circumferentially inclined relative to a rotational axis of the gas turbine engine, such that the first axial flange end is circumferentially offset from the second axial-flange end. Further, each split-line flange includes an intersecting portion, such that at least the intersecting portion of each split-line flange is circumferentially inclined relative to the rotational axis.

Casing assembly for gas turbine engine

A compressor section for a gas turbine engine, the compressor section including a circumferential row of stator vanes, a circumferential row of rotor blades, and a casing arrangement. A row of bleed ports are provided in the casing arrangement. Each of the bleed ports extends radially and is curved such that an outlet of the port is rearward of and is circumferentially offset in a direction of rotation of the rotor blades from an inlet of the port.

Compressor

A gas turbine engine sealing air system including a bearing chamber seal to prevent lubricant fluid loss from a fluid chamber, the sealing effected by ingress of sealing air. The system includes an air vent duct coupled to the bearing chamber seal. An ejector is located within the air vent duct to pump sealing air from the bearing chamber seal and through the air vent duct. A first duct is coupled between a starter air system of the gas turbine engine and the ejector to provide motive fluid to the ejector during gas turbine engine starting.

Gas turbine engine bearing chamber seals

A pressure controlled chamber arrangement comprising a chamber, one or more seals arranged in use to limit fluid flow between the inside and outside of the chamber and a vent system arranged in use to remove a vent flow of fluid from the inside of the chamber. The vent system at least contributes to the formation of a pressure drop across the one or more seals from the outside to the inside of the chamber. The vent system comprises a flow controller arranged in use to control the quantity of fluid removal via the vent system in accordance with the sealing performance provided by the one or more seals, the vent system giving increased quantity of fluid removal for decreased sealing performance.

Pressure controlled chamber

A gas turbine engine has a compressor section with rotational compressor components rotatable with respect to static compressor components. A compressed air bleed arrangement is provided to cool one or more other rotational components of the gas turbine engine. The compressed air bleed arrangement takes a flow of compressed air from the compressor section along an off-take passage. The off-take passage opens in the compressor section at an off-take port. The off-take passage is rotatable, in use, with the rotational compressor components. The compressed air bleed arrangement is operable to provide the air in the off-take passage with higher static pressure than the air in the compressor section at the off-take port, by diffusing the air in the off-take passage. The off-take passage further includes off-take vanes, operable to increase the tangential velocity of the air in the off-take passage compared with the air at the off-take port.

Gas turbine engine havinga rotatable off-take passage in a compressor section

A gas turbine engine comprising a turbine section cooling system and a method of cooling a turbine section of a gas turbine engine is provided. The gas turbine engine comprises in flow series a compressor section, a combustor, and a turbine section, the engine further comprising a turbine section cooling system. The turbine section cooling system including a first compressed air bleed arrangement and a second compressed air bleed arrangement. The first compressed air bleed arrangement bleeds a first flow of compressed air from a high pressure stage of the compressor section. The first flow of compressed air bypasses the combustor and arrives at the turbine section to form a sealing and/or cooling flow at a row of stator vanes upstream of an adjacent rotor disc. The second compressed air bleed arrangement bleeds a second flow of compressed air from one or more lower pressure stages of the compressor section. The second flow of compressed air bypasses the combustor and arrives at the turbine section to form a cooling flow. A first portion of the cooling flow is routed to a front face of the rotor disc and a second portion of the cooling flow is routed to a rear face of the rotor disc.

Gas turbine engine with turbine cooling arrangement

Apparatus is provided comprising first and second coaxial shafts, the first and second shafts extending along, and being rotatable about, a central axis. The first and second shafts are coupled together by a splined major coupling at a coupling region of the apparatus, which major coupling enables the shafts to co-rotate about the central axis. A torque element is connected to the first and second shafts, the torque element being operable to force the first and second shafts to relatively rotate about the central axis in order to apply a preload to the major coupling. The apparatus may be comprised in a gas turbine engine, whereupon the first shaft may be a fan shaft and the second shaft may be a LP turbine shaft.

Splined couplings

A casing component () of a turbomachine, the casing component () comprising: a plurality of casing elements () which define a diameter of the casing component; and an actuation means operable to change the diameter of the casing component (), wherein the actuation means changes the diameter of the casing component () as a function of a rotational speed of a rotatable component () disposed within the casing component ().

Casing component

A ducted fan gas turbine assembly has a propulsive fan driven by a core engine. The turbine assembly has an annular bypass duct which receives an airflow from the fan and is bounded on the radially inward side by the core engine and on the radially outward side by an engine nacelle. The turbine assembly has a thrust reverser assembly having a vane arrangement including a first set of turning vanes for channelling airflow from the duct in a rearward direction, a second set of turning vanes for channelling airflow from the bypass duct in a forward direction. The thrust reverser assembly includes a plurality of blocker doors which are deployable to block the bypass duct. The nacelle has stationary and movable cowl portions that cooperate with the thrust reverser assembly to provide several operational configurations.

Ducted fan gas turbine assembly

A control mechanism for moving at least two components of a gas turbine engine, the control mechanism comprises first and second crankshafts and a torsion bar connected to both crankshafts, each crankshaft is connected to a different component, the crankshafts engage one another via one-way abutment means and the torsion bar is pre-twisted so that up to a predetermined rotational extent the two crankshafts rotate together and beyond that extent only the first crankshaft rotates, thereby one of the two components moves more than the other component.

Multi-rotational crankshaft arrangement

A seal arrangementpresents a barrier in the form of a combined seal and anti-fret lineracross a shroud ring cavityin order to inhibit air leakage in the platform gap between platforms. It will be understood that these gaps are provided for thermal expansion and vibration of vane assemblies but create a leakage path between passageand cavitywhich can increase aerodynamic losses and therefore reduce engine compressor efficiency. By provision of a combined seal and anti-fret linersuch air losses are reduced whilst combining the function of an air leakage seal and an anti-fret liner to prevent abrasive degradation of the mountingsand platform sectionensures that there is a reduced component count as well as cost, weight and assembly time.

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Nottingham, United Kingdom

Crispin David Bolgar