A blade track for a gas turbine engine includes a plurality of blade track segments. The blade track segments are arranged circumferentially around a central axis to form the blade track.

A gas (or steam) turbine is disclosed including: a rotor with at least one array of rotor blades, a stator with a casing and a shroud ring; the shroud ring extends around the array of blades rotor and the casing extends around the shroud ring. The shroud ring has radial size independent from temperature due to its material and is movably coupled with the casing so to allow the casing of the stator to thermally expand and contract during operation of the turbine while maintaining the shroud ring radial size. Also, the rotor thermally expands and contracts during operation of the turbine, and, at working temperature, tip regions of the rotor blades are in close proximity to an inner region of the shroud ring so that clearance is small or zero at working condition.

A shroud attaching structure of a gas turbine includes: a support provided around an axis of the gas turbine and inside a casing of the gas turbine in a radial direction; and a shroud assembly attached to the support so as to cover an inner peripheral surface of the support, the shroud assembly being formed by laminating a large number of plate-shaped shroud elements containing a ceramic matrix composite. The shroud elements are lined up in a circumferential direction of the support. The adjacent shroud elements are arranged so as to be slidable on each other.

A turbine shroud segment has a body extending axially between a leading edge and a trailing edge and circumferentially between a first and a second lateral edge. Upstream and downstream plenums are defined in the body. The upstream plenum has a plurality of cooling inlets. The downstream plenum has a plurality of cooling outlets. A flow constricting slot extends across the body between the first and second lateral edges. The flow constricting slot fluidly connects the downstream plenum to the upstream plenum.

The present technique presents a blade 1 for a gas turbine 10. The blade 1 includes an airfoil 100 having an airfoil tip part 100a and a pressure side 102 and a suction side 104 meeting at a leading edge 106 and a trailing edge 108 and defining an internal space 100s of the airfoil 100. A squealer tip 80, 90 is arranged at the airfoil tip part 100a. The squealer tip 80, 90 comprises a suction side rail 90. The suction side rail 90 comprises a chamfer part 90x and at least one squealer tip cooling hole 99. The chamfer part 90x comprises a chamfer surface 9. An outlet 99a of the at least one squealer tip cooling hole 99 is disposed at the chamfer surface 9.

The present technique presents a blade 1 for a gas turbine 10. The blade 1 includes an airfoil 100 having an airfoil tip part 100a and a pressure side 102 and a suction side 104 meeting at a leading edge 106 and a trailing edge 108 and defining an internal space 100s of the airfoil 100. A squealer tip 80, 90 is arranged at the airfoil tip part 100a. The squealer tip 80, 90 comprises a suction side rail 90. The suction side rail 90 comprises a chamfer part 90x and at least one squealer tip cooling hole 99. The chamfer part 90x comprises a chamfer surface 9. An outlet 99a of the at least one squealer tip cooling hole 99 is disposed at the chamfer surface 9.

A case clearance control system comprising a blade outer air seal support structure having a plurality of protrusions extending radially from the blade outer air seal support structure opposite a blade outer air seal proximate the blade outer air seal support structure; a thermal control ring coupled to the blade outer air seal support structure, the thermal control ring including a plurality of receivers configured to couple with the plurality of protrusions; a thermal break formed between the plurality of protrusions and the plurality of receivers, the thermal break configured to control heat transfer between the blade outer air seal support structure and the thermal control ring; and a plurality of flow passages formed between the blade outer air seal support structure, the thermal control ring and the plurality of protrusions, the plurality of flow passages configured to allow cooling air flow to condition the thermal control ring and maintain thermal control ring dimensions.

A case clearance control system comprising a blade outer air seal support structure having a plurality of protrusions extending radially from the blade outer air seal support structure opposite a blade outer air seal proximate the blade outer air seal support structure; a thermal control ring coupled to the blade outer air seal support structure, the thermal control ring including a plurality of receivers configured to couple with the plurality of protrusions; a thermal break formed between the plurality of protrusions and the plurality of receivers, the thermal break configured to control heat transfer between the blade outer air seal support structure and the thermal control ring; and a plurality of flow passages formed between the blade outer air seal support structure, the thermal control ring and the plurality of protrusions, the plurality of flow passages configured to allow cooling air flow to condition the thermal control ring and maintain thermal control ring dimensions.

A blade outer air seal (BOAS) for a gas turbine engine according to one disclosed non-limiting embodiment of the present disclosure includes a seal body having a radially inner face and a radially outer face that axially extend between a leading edge portion and a trailing edge portion; and a multitude of cooling fins disposed around the radially outer face of the seal body, at least one of the multitude of cooling fins having at least one aperture.

Example apparatus, systems, and articles of manufacture to control deflection mismatch are disclosed herein. Further examples and combinations thereof include: A deflection limiter comprising an inner shroud segment to support a stator structure, the inner shroud segment including a first end face and a first outer upper portion, the first end face positioned radially inward and aft relative to the first outer upper portion, and an outer shroud segment to support the inner shroud segment, the outer shroud segment including a second end face and a second outer upper portion, the second end face positioned aft relative to the first end face and the second outer upper portion positioned aft relative to the first outer upper portion of the inner shroud segment, the second end face coupled to the first end face of the inner shroud segment and the second outer upper portion coupled to the first outer upper portion.