Abstract: A method comprises inspecting a ceramic matrix composite component assembled in a gas turbine engine to determine an extent of damage to the ceramic matrix composite component, determining a repair technique to repair the damage to the ceramic matrix composite component based on the extent of damage to the ceramic matrix composite component, and repairing the ceramic matrix composite component using the repair technique.

Abstract: A gas turbine engine includes a shroud ring, a vane ring, and a secondary flow assembly. The shroud ring extends around an associated turbine wheel and includes a carrier segment and a blade track segment comprising ceramic matrix composite materials. The blade track segment and the carrier segment defining a shroud cavity radially therebetween. The vane ring includes a heat shield comprising ceramic matrix composite materials that forms an outer platform, an inner platform, and an airfoil defining a heat shield cavity radially through the heat shield. The secondary flow assembly includes a recirculating flow circuit having a discharge tube and a vane pressurizing tube; and the secondary flow assembly is configured to cool the blade track segment. The flow circuit is further configured to pressurize the heat shield cavity thereby providing a seal against the gases from the primary gas path entering the heat shield cavity.

Abstract: A turbine assembly for use with a gas turbine engine includes a bladed wheel assembly, a vane assembly, and an inner seal. The bladed wheel assembly is adapted to interact with gases flowing through a gas path of the gas turbine engine. The vane assembly is located upstream of the bladed wheel assembly and adapted to direct the gases at the bladed wheel assembly. The inner seal is configured to block gases from passing around the vane assembly.

Abstract: A turbine vane assembly adapted for use in a gas turbine engine includes a plurality of turbine vanes, an outer vane support, and an inner vane support. The plurality of turbine vanes comprise ceramic matrix composite material and are adapted to interact with hot gases flowing through a gas path of the gas turbine engine during use of the turbine vane assembly.

Abstract: A turbine vane assembly includes a plurality of turbine vanes, an outer vane support, and an inner vane support. The turbine vanes are adapted to interact with gases flowing through a gas path of the gas turbine engine. The outer vane support is arranged radially outward of the turbine vane and configured to receive aerodynamic loads from the turbine vanes. The inner vane support is arranged radially inward from the outer vane support to locate the turbine vane assembly therebetween.

Abstract: An assembly adapted for use in a gas turbine engine includes a carrier and a blade track segment. The carrier extends at least partway about an axis. The blade track segment is supported by the carrier radially relative to the axis to define a portion of a gas path of the assembly.

Abstract: A method comprises inspecting a ceramic matrix composite component assembled in a gas turbine engine to determine an extent of damage to the ceramic matrix composite component, determining a repair technique to repair the damage to the ceramic matrix composite component based on the extent of damage to the ceramic matrix composite component, and repairing the ceramic matrix composite component using the repair technique.

Abstract: An assembly adapted for use in a gas turbine engine includes a carrier and a blade track segment. The carrier extends at least partway about an axis. The blade track segment is supported by the carrier radially relative to the axis to define a portion of a gas path of the assembly.

Abstract: A gas turbine engine includes a shroud ring, a vane ring, and a secondary flow assembly. The shroud ring extends around an associated turbine wheel and includes a carrier segment and a blade track segment comprising ceramic matrix composite materials. The blade track segment and the carrier segment defining a shroud cavity radially therebetween. The vane ring includes a heat shield comprising ceramic matrix composite materials that forms an outer platform, an inner platform, and an airfoil defining a heat shield cavity radially through the heat shield. The secondary flow assembly includes a recirculating flow circuit having a discharge tube and a vane pressurizing tube; and the secondary flow assembly is configured to cool the blade track segment. The flow circuit is further configured to pressurize the heat shield cavity thereby providing a seal against the gases from the primary gas path entering the heat shield cavity.

Abstract: A turbine vane assembly includes a plurality of turbine vanes, an outer vane support, and an inner vane support. The turbine vanes are adapted to interact with gases flowing through a gas path of the gas turbine engine. The outer vane support is arranged radially outward of the turbine vane and configured to receive aerodynamic loads from the turbine vanes. The inner vane support is arranged radially inward from the outer vane support to locate the turbine vane assembly therebetween.

Abstract: A turbine assembly for use with a gas turbine engine includes a bladed wheel assembly, a vane assembly, and an inner seal. The bladed wheel assembly is adapted to interact with gases flowing through a gas path of the gas turbine engine. The vane assembly is located upstream of the bladed wheel assembly and adapted to direct the gases at the bladed wheel assembly. The inner seal is configured to block gases from passing around the vane assembly.

Abstract: A turbine vane assembly adapted for use in a gas turbine engine includes a plurality of turbine vanes, an outer vane support, and an inner vane support. The plurality of turbine vanes comprise ceramic matrix composite material and are adapted to interact with hot gases flowing through a gas path of the gas turbine engine during use of the turbine vane assembly.

Abstract: The present invention provides an airfoil for a first stage turbine blade having an external surface with first and second sides. The external surface extends spanwise between a hub and a tip and streamwise between a leading edge and a trailing edge of the airfoil. The external surface includes a contour substantially defined by Table 1 as listed in the specification.

Abstract: A cooling arrangement is provided in which an incident coolant flow is presented to a passage such that a proportion of that flow is diverted in order to create a standing relative overpressure in a chamber compared to the actual static pressure of the flow in the passage. In such circumstances through flow transfer apertures in the walls of the chamber forced coolant flows are presented for impingement upon surfaces and therefore greater heat transfer. Generally, coolant flow is presented to both ends of the chamber in order to create the relative standing overpressure in the chamber relative to the current static pressure presented in the passage through the incident flow. In such circumstances through utilizing the dynamic rotation or other motion of a component incorporating the arrangement a higher relative pressure is achieved for greater force coolant flow projection compared to the incident or actual static pressure of the incident coolant flow in the.

Abstract: The present invention provides an airfoil for a first stage turbine blade having an external surface with first and second sides. The external surface extends spanwise between a hub and a tip and streamwise between a leading edge and a trailing edge of the airfoil. The external surface includes a contour substantially defined by Table 1 as listed in the specification.

Abstract: A cooling arrangement 11 is provided in which an incident coolant flow 23 is presented to a passage 14 such that a proportion of that flow is diverted in order to create a standing relative overpressure in a chamber 24 compared to the actual static pressure of the flow 23 in the passage 14. In such circumstances through flow transfer apertures in the walls of the chamber 24 forced coolant flows are presented for impingement upon surfaces 26, 27 and therefore greater heat transfer. Generally, coolant flow is presented to both ends of the chamber 24 in order to create the relative standing overpressure in the chamber 24 relative to the current static pressure presented in the passage 14 through the incident flow 23.

Abstract: A heat transfer arrangement comprises a target member (28 or 30), and an impingement member (32). The impingement member (32) defines a fluid path (38 or 40) there through to direct fluid onto the target member. The arrangement includes a fluid directing formation (42) on the impingement member (32). The fluid path (38 or 40) extends through the fluid directory formation (42) such that the fluid path directs fluid to exit there from at an exit angle that is substantially orthogonal to the fluid directing formation.

Abstract: A component such as a turbine blade for a gas turbine engine has a cooling arrangement comprising a supply passage 18 within the blade from which extend cooling passages 28 of a first array which open at discharge openings 29, for example at a trailing edge of the blade. Cooling passages 30 of a second array extend into the blade from discharge openings 31, and intersect the passages 28. The passages 30 terminate short of the supply passage 18. As a result of this arrangement, the limiting flow area for cooling air is defined by the cooling passages 28 of the first array, and is not affected by the accuracy with which the cooling passages 30 of the second array intersect the cooling passages 28 of the first array.