Abstract: A component for the turbine of a gas turbine engine is provided. The component two facing walls interconnected by one or more generally elongate divider members to partially define side-by-side, generally elongate, cooling fluid passage portions which form a multi-pass cooling passage within the component. The passage portions are connected in series fluid flow relationship by respective bends formed by joined ends of neighbouring of the passage portions. The component further includes one or more core tie linking passages formed in the divider members. One or more differential pressure reducing arrangements are formed in the multi-pass cooling passage adjacent respective of the core tie linking passages.

Abstract: A rotor blade 40 for a gas turbine engine has an aerofoil portion 42 from a root 48 to a tip 54. In use, combustion gas may leak over the tip 54 from the pressure face 52 to the suction face 50. A gutter 62 extends across the tip 54 to entrain any over tip leakage gap. The floor of the gutter defines an increased depth portion 72 at the exit end of the gutter 62.

Abstract: A turbine blade for a gas turbine engine has an aerofoil portion extending from a root to a tip. The tip carries winglets. A gutter extends across the tip to entrain gas leaking around the tip (over tip leakage). The aerofoil portion has a mean camber line and the gutter has a center line. In the examples described, the conditions that (a) the mean camber line and the centre line coincide at the exit when viewed from the tip towards the root, and (b) the mean camber line and the center line are parallel at the exit when viewed as aforesaid, are not both fulfilled.

Abstract: A cooling arrangement is provided for a turbine disc of a gas turbine engine. The turbine disc has a plurality of circumferentially spaced disc posts forming fixtures therebetween for a row of turbine blades. Each turbine blade has an attachment formation which engages at a respective fixture, a platform radially outwardly of the attachment formation such that the adjacent platforms of the row form an inner endwall for the working gas annulus of the engine, and an aerofoil which extends radially outwardly from the platform. A respective cavity is formed between an exposed radially outer surface of each disc post and the inner endwall. The cooling arrangement has at each disc post, a cooling plate located in the respective cavity and spaced radially outwardly from the exposed outer surface of the disc post to form a cooling channel between the cooling plate and the exposed outer surface.

Abstract: A component for the turbine of a gas turbine engine is provided. The component two facing walls interconnected by one or more generally elongate divider members to partially define side-by-side, generally elongate, cooling fluid passage portions which form a multi-pass cooling passage within the component. The passage portions are connected in series fluid flow relationship by respective bends formed by joined ends of neighbouring of the passage portions. The component further includes one or more core tie linking passages formed in the divider members. One or more differential pressure reducing arrangements are formed in the multi-pass cooling passage adjacent respective of the core tie linking passages.

Abstract: A cooling arrangement is provided for a turbine disc of a gas turbine engine. The turbine disc has a plurality of circumferentially spaced disc posts forming fixtures therebetween for a row of turbine blades. Each turbine blade has an attachment formation which engages at a respective fixture, a platform radially outwardly of the attachment formation such that the adjacent platforms of the row form an inner endwall for the working gas annulus of the engine, and an aerofoil which extends radially outwardly from the platform. A respective cavity is formed between an exposed radially outer surface of each disc post and the inner endwall. The cooling arrangement has at each disc post, a cooling plate located in the respective cavity and spaced radially outwardly from the exposed outer surface of the disc post to form a cooling channel between the cooling plate and the exposed outer surface.

Abstract: A rotor blade 40 for a gas turbine engine has an aerofoil portion 42 from a root 48 to a tip 54. In use, combustion gas may leak over the tip 54 from the pressure face 52 to the suction face 50. A gutter 62 extends across the tip 54 to entrain any over tip leakage gap. The floor of the gutter defines an increased depth portion 72 at the exit end of the gutter 62.

Abstract: A turbine blade (40) for a gas turbine engine has an aerofoil portion (42) extending from a root (48) to a tip (54). The tip (54) carries winglets (56, 58). A gutter (62) extends across the tip (54) to entrain gas leaking around the tip (54) (over tip leakage). The aerofoil portion (42) has a mean camber line and the gutter (62) has a centre line. In the examples described, the conditions that (a) the mean camber line and the centre line coincide at the exit when viewed from the tip towards the root, and (b) the mean camber line and the centre line are parallel at the exit when viewed as aforesaid, are not both fulfilled.