Abstract: A method of testing flow in an airfoil includes isolating a first cooling passage network of the airfoil by applying plugs to internal inlet orifices of the other ones of the cooling passage networks in the airfoil. The plugs block flow from an internal core cavity of the airfoil into the other ones of the cooling passage networks. The inlet orifices of the first cooling passage network are left open. A test flow is then applied to the internal core cavity and flows through the inlet orifices of the first cooling passage network.

Abstract: Methods for manufacturing airfoils of gas turbine engines are provided. The methods include forming a main body core, the main body core including a feed cavity core portion, the main body core configured to form at least a part of an airfoil including an airfoil body, a platform, and an attachment element, forming a platform circuit core having a platform core extension, wherein the platform circuit core is configured to form a cooling circuit in the platform, wherein at least one of the feed cavity core portion and the platform core extension includes a notch, assembling the platform circuit core to the main body core such that the platform core extension engages with the feed cavity core portion at the notch, casting an airfoil using the assembled platform circuit core and main body core, and removing a platform core extension element that is formed at the location of the notch.

Abstract: A core assembly includes a core that includes an exterior surface that has a recessed area that extends along the exterior surface. An insert includes a contact surface that corresponds to the recessed area.

Abstract: A component for a gas turbine engine includes a first component end, a second component end and a cavity extending from the first component end to the second component end. A baffle assembly is located in the cavity including a first baffle portion inserted into the cavity and a second baffle portion inserted into the cavity. The first baffle portion overlaps the second baffle portion. A vane for a gas turbine engine includes an outer platform, an inner platform and an airfoil portion extending. A core extends through the airfoil portion. A baffle is positioned in the core and is offset from the core wall. The baffle includes a first baffle portion inserted into the core and a second baffle portion inserted into the core. One of the first baffle portion or the second baffle portion overlaps the other of the first baffle portion or the second baffle portion.

Abstract: Airfoils bodies having a first core cavity and a second core cavity located within the airfoil body that is adjacent the first core cavity. The second core cavity is defined by a first cavity wall, a second cavity wall, a first exterior wall, and a second exterior wall, wherein the first cavity wall is located between the second core cavity and the first core cavity and the first and second exterior walls are exterior walls of the airfoil body. The first cavity wall includes a first surface angled toward the first exterior wall and a second surface angled toward the second exterior wall. At least one first cavity impingement hole is formed within the first surface and a central ridge extends into the second core cavity from at least one of the first cavity wall and the second wall and divides the second core cavity into a two-vortex chamber.

Abstract: A component for a gas turbine engine includes, among other things, an airfoil that includes a pressure sidewall and a suction sidewall that meet together at both a leading edge and a trailing edge, the airfoil extending radially from a platform to a tip, a tip pocket formed in the tip and terminating prior to the trailing edge, and one or more heat transfer augmentation devices formed in the tip pocket.

Abstract: A method for forming a hole within a ceramic matrix composite component includes forming a first core portion for a ceramic matrix composite component; embedding a hollow member into the first core portion at a desired location; wrapping the first core portion with a first ceramic matrix composite material; inserting a rod through the hollow member and into the first core portion; removing the hollow member; assembling a second core portion to the first core portion such that the rod extends into the second core portion; and wrapping the first core portion and the second core portion with a second ceramic matrix composite material.

Abstract: An airfoil includes a core structure, a passage, and a panel attached with the core structure adjacent the passage. The panel has an exterior gas path side and an opposed interior side bordering the passage. The interior side includes at least one flow guide in the passage.

Abstract: Airfoils for gas turbine engines are provided. The airfoils include an airfoil body extending between leading and trailing edges in an axial direction, between pressure and suction sides in a circumferential direction, and between a root and tip in a radial direction, a first shielding sidewall cavity located adjacent one of the pressure and suction sides proximate the root of the airfoil body and extending radially toward the tip, a second shielding sidewall cavity located adjacent the other of the pressure and suction sides proximate the root of the airfoil body and extending radially toward the tip, and a shielded sidewall cavity located between the first shielding sidewall cavity and the second shielding sidewall cavity, wherein the shielded sidewall cavity is not adjacent either of the pressure or suction sides proximate the root and transitions to be proximate at least one of the pressure and suction sides proximate the tip.

Abstract: A refractory metal core (RMC) finishing method according to an exemplary aspect of the present disclosure includes, among other things, performing a plurality of finishing operations on a plurality of RMC samples, analyzing one or more properties of at least a portion of the plurality of RMC samples and selecting a combination of finishing operations for generating an RMC having desirable properties for manufacturing a part free from defects.

Abstract: An airfoil includes an airfoil structure that defines a portion of an airfoil profile and a panel that defines a different portion of the airfoil profile. The panel has perimeter edges that meet the airfoil structure in a perimeter joint, and there is a perimeter seal disposed in the perimeter joint.

Abstract: A gas turbine engine article includes a blade that has a platform, an airfoil, and a root. The platform has a gaspath side and a non-gaspath side. The airfoil extends radially from the gaspath side and defines a leading end and a trailing end. The root is configured to secure the blade. The root extends radially from the non-gaspath side of the platform. The root defines forward and aft axial faces, and a radial channel in the aft axial face. The radial channel leads to an inlet orifice in the root. A cooling passage extends in the root from the inlet orifice.

Abstract: A component according to an exemplary aspect of the present disclosure includes, among other things, a wall and a hollow vascular engineered lattice structure formed inside of the wall. The hollow vascular engineered lattice structure has an inlet hole and an outlet hole that communicate fluid into and out of the hollow vascular structure. The hollow vascular engineered lattice structure further has at least one resupply inlet hole between the inlet hole and the outlet hole with respect to a dimension of the component to communicate additional fluid into the hollow vascular engineered lattice structure.

Abstract: A refractory metal core (RMC) finishing method according to an exemplary aspect of the present disclosure includes, among other things, performing a plurality of finishing operations on a plurality of RMC samples, analyzing one or more properties of at least a portion of the plurality of RMC samples and selecting a combination of finishing operations for generating an RMC having desirable properties for manufacturing a part free from defects.

Abstract: An airfoil includes a core structure that defines a cooling passage and first and second panels that are attached with the core structure. Each of the first and second panels includes an exterior gas path side, an opposed interior side, and a side edge. The side edge has a bearing portion that defines a bearing surface. The bearing surface of the first panel abuts the bearing surface of the second panel at a bearing interface. There are channels in the bearing portion of the first panel. Each of the channels has a first end that opens to the cooling passage and a second end that opens to the exterior gas path side at one of the side edges. Each of the channels is defined on a side by the bearing surface of the second panel.

Abstract: An airfoil includes an airfoil body that has a geometrically segmented coating section. The geometrically segmented coating section includes a wall that has an outer side. The outer side includes an array of cells and there is a coating disposed in the array of cells. The cells include respective secondary mechanical bonding features that secure the coating to the wall.

Abstract: An airfoil includes an airfoil body that has a geometrically segmented coating section. The geometrically segmented coating section includes a wall having an outer side. The outer side has an array of cells, and there is a coating disposed in the array of cells.

Abstract: Airfoil bodies having a first core cavity and a second core cavity located within the airfoil body and adjacent the first core cavity, wherein the second core cavity is defined by a first cavity wall, a second cavity wall, a first exterior wall, and a second exterior wall, wherein the first cavity wall is located between the first and second core cavities. The first cavity wall includes a first surface angled toward the first exterior wall and a second surface angled toward the second exterior wall. At least one first cavity impingement hole is formed within the first surface. At least one circuit exit is located in the first exterior wall, the at least one circuit exit arranged to expel air from the second core cavity through the first exterior wall.

Abstract: A component for a gas turbine engine includes a root with a neck that extends into a fir tree with at least one tooth, the root includes a feed passage in communication with a multiple of cooling passages that extend through the neck and fir tree.

Abstract: A combustor for a gas turbine engine includes a combustor shell having a shell opening therethrough, a combustor panel having a stud attached thereto, the stud extending through the shell opening. The stud includes a standoff to define an intermediate passage between the combustor shell and the combustor panel. A retainer is attached to the stud. A washer surrounds the stud and is positioned between the retainer and the combustor shell. The washer at least partially defines a cooling flow passage configured to direct a cooling airflow through the shell opening to impinge the cooling flow on at least one of the stud or the standoff.