Abstract: A component for a gas turbine engine, comprising: first and second walls; a coolant channel defined by the space between the first and second walls; and a first rib extending between the first and second walls to the end of the coolant channel in a coolant flow direction, such that the coolant channel is bifurcated in the coolant flow direction.

Abstract: A vane assembly includes an aerofoil having a leading edge, a trailing edge, and a pressure surface and a suction surface defined between the leading edge and the trailing edge. The aerofoil includes a blade member forming the trailing edge, at least a portion of the pressure surface and at least a portion of the suction surface. The blade member is formed of a first material. The aerofoil further includes a spar at least partly enclosed by the blade member and forming at least a portion of the leading edge. The spar further forms at least one cooling channel and supports at least a portion of an interior surface of the blade member. The spar is formed of a second material different from the first material. The second material has a greater impact resistance than the first material.

Abstract: The present disclosure concerns removal of entrained contaminant particles in a coolant airflow for a gas turbine engine. Example embodiments include a coolant airflow assembly for a gas turbine engine, comprising: a coolant feed passage connected between a supply of coolant air and an inlet of a component to be cooled, the coolant feed passage defining a coolant airflow path and comprising first and second opposing internal faces (305, 306), the inlet of the component connected to the coolant airflow path through one of the first and second internal faces (305, 306) of the coolant feed passage; and a particulate filter for removing entrained particles from the coolant airflow path, comprising: a first filter panel extending from the first face into the coolant airflow path upstream of the inlet of the component; and a second filter panel extending from the second face into the coolant airflow path upstream of the first filter panel.

Abstract: A vane assembly includes an aerofoil having a leading edge, a trailing edge, and a pressure surface and a suction surface defined between the leading edge and the trailing edge. The aerofoil includes a blade member forming the trailing edge, at least a portion of the pressure surface and at least a portion of the suction surface. The blade member is formed of a first material. The aerofoil further includes a spar at least partly enclosed by the blade member and forming at least a portion of the leading edge. The spar further forms at least one cooling channel and supports at least a portion of an interior surface of the blade member. The spar is formed of a second material different from the first material. The second material has a greater impact resistance than the first material.

Abstract: Aerofoil including: a body comprising first body portion and second body portion bonded to first body portion, first and second body portions having outer faces forming an outer surface of the aerofoil, and inner faces opposing the outer faces; a channel formed within the aerofoil by the inner faces of the first body portion and the second body portion; and a dividing sheet provided within the channel, such that a cooling passage is formed between the dividing sheet and the inner face of at least one of the first body portion and second body portion, cooling passage arranged to conduct cooling fluid adjacent the body, for cooling the body; wherein the body has a rear face and an opposing front face; and the dividing sheet extends parallel to at least a portion of a rear face of the body, such that the cooling passage formed parallel to the rear face.

Abstract: A method of manufacturing the aerofoil, and an aerofoil. The method includes: casting a first body portion having a passage passing there-through; casting a second body portion; and joining the first body portion and the second body portion to form the aerofoil. The passage extends from an opening at a first end at or near a leading edge to an opening at a second end at or near a trailing edge. The passage is formed during the casting process by a core; and the core is supported at the first and second ends through the respective openings of the passage.

Abstract: A component for a gas turbine engine, comprising: first and second walls; a coolant channel defined by the space between the first and second walls; and a first rib extending between the first and second walls to the end of the coolant channel in a coolant flow direction, such that the coolant channel is bifurcated in the coolant flow direction.

Abstract: A method for manufacturing a component having a defined geometry includes: a) defining a pre-component geometry including interim shape elements and additional, sacrificial elements for supporting interim elements; b) on a base plate, depositing multiple layers of a powder including a material from which the pre-component will be manufactured; c) sintering the powder to form the pre-component to the defined geometry; d) removing at least some of the sacrificial elements from the pre-component; e) subjecting the remaining pre-component from step d) to a HIP step; and f) removing remaining sacrificial elements from the pre-component product of step e) to provide a component to the defined component geometry. In the definition of the pre-component geometry, the interim elements differ from the corresponding final shape elements in the defined component geometry such that during the HIP step, the interim shape elements adjust to form final shape elements in the defined component geometry.

Abstract: The present disclosure concerns removal of entrained contaminant particles in a coolant airflow for a gas turbine engine. Example embodiments include a coolant airflow assembly for a gas turbine engine, comprising: a coolant feed passage connected between a supply of coolant air and an inlet of a component to be cooled, the coolant feed passage defining a coolant airflow path and comprising first and second opposing internal faces (305, 306), the inlet of the component connected to the coolant airflow path through one of the first and second internal faces (305, 306) of the coolant feed passage; and a particulate filter for removing entrained particles from the coolant airflow path, comprising: a first filter panel extending from the first face into the coolant airflow path upstream of the inlet of the component; and a second filter panel extending from the second face into the coolant airflow path upstream of the first filter panel.

Abstract: A cooled gas turbine engine component includes a wall having a plurality of effusion cooling apertures extending there-through from a first surface to a second surface. Each aperture has an inlet in the first surface and an outlet in the second surface. Each aperture includes an inlet portion, a collection chamber, a metering portion, a U-shaped bend portion and a diffusing portion arranged in flow series from the inlet to the outlet. The inlet portion of each aperture is arranged substantially perpendicularly to a surface of the collection chamber. The metering portion of each aperture is arranged to extend longitudinally from a first lateral side of the collection chamber and the diffusing portion of each aperture is arranged at an angle to the second surface. Each outlet has a quadrilateral shape in the second surface of the wall and each outlet is displaced laterally from the metering portion.

Abstract: A method of manufacturing the aerofoil, and an aerofoil. The method includes: casting a first body portion having a passage passing there-through; casting a second body portion; and joining the first body portion and the second body portion to form the aerofoil. The passage extends from an opening at a first end at or near a leading edge to an opening at a second end at or near a trailing edge. The passage is formed during the casting process by a core; and the core is supported at the first and second ends through the respective openings of the passage.

Abstract: Aerofoil including: a body comprising first body portion and second body portion bonded to first body portion, first and second body portions having outer faces forming an outer surface of the aerofoil, and inner faces opposing the outer faces; a channel formed within the aerofoil by the inner faces of the first body portion and the second body portion; and a dividing sheet provided within the channel, such that a cooling passage is formed between the dividing sheet and the inner face of at least one of the first body portion and second body portion, cooling passage arranged to conduct cooling fluid adjacent the body, for cooling the body; wherein the body has a rear face and an opposing front face; and the dividing sheet extends parallel to at least a portion of a rear face of the body, such that the cooling passage formed parallel to the rear face.

Abstract: An aerofoil configured to be connected to a second component, the aerofoil comprising a first face including a bi-cast groove for joining the aerofoil to the second component by bi-casting; a second face connected to a main body of the aerofoil; an angled sealing face between the first and second face, the angled sealing face being angled with respect to the second face and configured to provide a sealing interface between the first face and the second face; and a concave face forming a recessed surface at the intersection of the sealing face and the second face.

Abstract: A method of bi-casting a first component (10) to a second component (11), wherein the first component comprises a channel (13) having an opening at a surface of the component, the method comprising inserting a pin (12) into the channel (13) to a closure position to block a cross-section of said channel, positioning the first and second components to be adjacent such that at least one space is defined between the first and second components, pouring liquid bi-cast metal (15) into the at least one space between the first component and the second component such that the first component is joined to the second component by the bi-cast metal and the pin is secured in the closure position.

Abstract: A gas turbine engine comprising a rotor blade, an inner casing and an outer casing, a guide vane, located upstream of the rotor blade and between the inner and outer casing, the guide vane comprising a vane contact member, wherein the guide vane is configured to pivot upon axial displacement of the inner casing relative to either or both of the outer casing and the rotor blade; and, a seal segment located radially outwardly of the rotor blade and mounted for radial displacement relative to the rotor blade; wherein at least a portion of the seal segment axially overlaps, and slidably engages a downstream end of the guide vane so that a pivoting of the guide vane causes the vane contact member to undergo a radially outward displacement which causes a radially outward displacement of at least a portion of the seal segment.

Abstract: A cooled gas turbine engine component includes a wall having a plurality of effusion cooling apertures extending there-through from a first surface to a second surface. Each aperture has an inlet in the first surface and an outlet in the second surface. Each aperture includes an inlet portion, a collection chamber, a metering portion, a U-shaped bend portion and a diffusing portion arranged in flow series from the inlet to the outlet. The inlet portion of each aperture is arranged substantially perpendicularly to a surface of the collection chamber. The metering portion of each aperture is arranged to extend longitudinally from a first lateral side of the collection chamber and the diffusing portion of each aperture is arranged at an angle to the second surface. Each outlet has a quadrilateral shape in the second surface of the wall and each outlet is displaced laterally from the metering portion.

Abstract: A method for manufacturing a component having a defined geometry includes: a) defining a pre-component geometry including interim shape elements and additional, sacrificial elements for supporting interim elements; b) on a base plate, depositing multiple layers of a powder including a material from which the pre-component will be manufactured; c) sintering the powder to form the pre-component to the defined geometry; d) removing at least some of the sacrificial elements from the pre-component; e) subjecting the remaining pre-component from step d) to a HIP step; and f) removing remaining sacrificial elements from the pre-component product of step e) to provide a component to the defined component geometry. In the definition of the pre-component geometry, the interim elements differ from the corresponding final shape elements in the defined component geometry such that during the HIP step, the interim shape elements adjust to form final shape elements in the defined component geometry.

Abstract: Methods, apparatus, computer programs, and non-transitory computer readable storage mediums for repairing an aerofoil of a gas turbine engine A method of repairing an aerofoil of a gas turbine engine, the method comprising: controlling machining of at least one of: a first aerofoil and a second aerofoil; the machining causing an increase in a throat area defined by the first aerofoil, the second aerofoil, a first platform and a second platform, the first aerofoil and the second aerofoil being arranged to couple to the first platform and to the second platform; and controlling provision of a coating to at least one of the first platform and the second platform subsequent to controlling machining to reduce the throat area.

Abstract: A vane is provided for directing hot gases in a gas turbine engine. The vane includes a hollow aerofoil portion, which in use spans the working gas annulus of the engine. The vane further includes an impingement tube which forms a covering over the interior surface of the aerofoil portion and which has jet-forming apertures formed therein for the production of impingement cooling jets. The impingement tube includes two tube portions which are separately insertable into position into the aerofoil portion to form the covering. The impingement tube further includes an expansion member which, when the tube portions are in position in the aerofoil portion, is locatable in the aerofoil portion to urge each tube portion outwardly and thereby holds the tube portions in position against the aerofoil portion.

Abstract: A strip seal for providing a seal between adjacent components, which include a first groove and a second groove angled with respect to and intersecting the first groove. The grooves of respective components are opposed to one another when the components are assembled. The strip seal includes a first strip section to span between and be received in the opposed first grooves, and a second strip section to span between and be received in the opposed second grooves. One end of the first strip section is arranged to butt against the second strip section at the intersection of the grooves. The other end of the first strip section has at least one deformable projection that can engage the end of the first groove. The projection deforms as it engages the end of the groove and urges the end of the first strip section against the second strip section.