Abstract: An internally cooled airfoil for a gas turbine engine has a hollow airfoil body defining a core cavity. An insert is mounted in the core cavity. A cooling gap is provided between the insert and the hollow airfoil body. A plurality of standoffs project across the cooling gap. Trip-strips projecting laterally between adjacent standoffs. The trip-strips and the standoffs may be integrated into a unitary heat transfer feature.

Abstract: A combustor heat shield comprises a heat shield body adapted to be mounted to a combustor wall with a back side of the heat shield body in spaced-apart facing relationship with the combustor wall to define an air gap between the heat shield body and the combustor wall. At least one nozzle opening is defined in the heat shield bod. The opening is bordered by a nozzle opening boss. The boss extends from the back side of the heat shield body across the air gap for sealing engagement with an adjacent part of the combustor. An annular array of effusion holes is provided adjacent the nozzle opening boss. The effusion holes extend through the heat shield body for passing cooling air from the back side to a front side of the heat shield body. Fins are interspersed between the effusion holes on the back side of the heat shield.

Abstract: An internally cooled airfoil, such as a turbine blade, has an airfoil section extending between a tip and a root. The interior of the airfoil includes a distribution of lands at the trailing edge in the span direction. A width of each of the lands is a widest dimension in the span direction of the land in the interior of the airfoil. A pitch is a distance in the span direction between centerlines of two adjacent lands. The pitch is constant throughout the distribution of the lands. The distribution of the lands includes at least two different widths.

Abstract: A heat shield for a combustor of a gas turbine engine has a heat shield adapted to be mounted to a combustor wall with a back face of the heat shield in spaced-apart facing relationship with an inner surface of the combustor wall to define an air gap. Rails extend from the back face of the heat shield across the air gap. Grooves are defined in at least one of the rails. The rail grooves are in fluid flow communication with the air gap when the heat shield is mounted to the combustor wall.

Abstract: A combustor for a gas turbine engine comprises a single skin liner defining a combustion chamber. The single skin liner has an inner surface facing the combustion chamber and an outer surface exposed to a coolant flow discharged in a plenum extending from the outer surface of the single skin liner to the engine casing. Cooling holes extend through the single skin liner. Open flow guiding channels are provided on the outer surface of the single skin liner, the open flow guiding channels being uncovered and aligned with the flow of air over the outer surface of the single skin liner.

Abstract: A combustor for a gas turbine engine comprises a single skin liner defining a combustion chamber. The single skin liner has an inner surface facing the combustion chamber and an outer surface exposed to a coolant flow discharged in a plenum extending from the outer surface of the single skin liner to the engine casing. Cooling holes extend through the single skin liner. Cooling protuberances, such as fins or pin fins, project integrally from the outer surface of the single skin liner into the plenum, the cooling fins being interspersed between the cooling holes.

Abstract: A cooling arrangement provides cooling around a dilution hole defined in a liner circumscribing a combustion chamber of a gas turbine engine. The cooling arrangement comprises a hollow boss projecting from an outer surface of the liner about the dilution hole. The hollow boss defines an internal cavity extending circumferentially around the dilution hole. The internal cavity has an inlet in fluid flow communication with an air plenum surrounding the liner and an outlet in fluid flow communication with the combustion chamber.

Abstract: A turbine blade for a gas turbine engine with an airfoil portion defined by a perimeter wall surrounding at least one enclosure, the perimeter wall having a plurality of cooling holes defined therethrough and providing fluid communication between the at least one enclosure and a gaspath of the gas turbine engine. The plurality of cooling holes includes at least one set of holes selected from the group consisting of a first set, a second set, a third set, a fourth set, a fifth set and a sixth set, wherein the first, second, third, fourth, fifth and sixth sets of holes respectively include the holes numbered A1 to A8, B1 to B10, C1 to C9, D1 to D6, E1 to E7 and F1 to F6 each located such that a central axis thereof extends through the respective point 1 and point 2 having a nominal location in accordance with the X, Y, Z Cartesian coordinate values set forth in Table 3.

Abstract: An internally cooled turbine vane for a gas turbine engine has coolant flow channels between the interior walls of the vane and an insert, where the channels serve to convey a portion of the cooling air flow from a pressure side chamber to a suction side chamber. The turbine vane defines a radially extending passage with a dividing wall defining a front section and a rear section; the rear section having interior walls spaced apart from an insert to define the pressure side chamber and the suction side chamber. The insert may receive cooling air and conveys the cooling air into the pressure side chamber and the suction side chamber. A front surface of the insert or a rear surface of the dividing wall may have a clearance gap and an air flow channel communicating between the pressure side chamber and the suction side chamber.

Abstract: A heat shield for a combustor of a gas turbine engine has a heat shield adapted to be mounted to a combustor wall with a back face of the heat shield in spaced-apart facing relationship with an inner surface of the combustor wall to define an air space. Concentric rails extend from the back face of the heat shield across the air space surrounding a nozzle opening in the heat shield. Effusion holes are provided between the concentric rails and extend between the back and front faces. Fins may be placed between the effusion holes.

Abstract: There is provided a combustor comprising a dome and a shell extending from the dome defining a combustion chamber. A dome heat shield is mounted to the dome inside the combustion chamber. A front heat shield is mounted to the shell and spaced therefrom. The dome heat shield has a lip extending generally away from the dome heat shield and generally parallel to the shell and spaced inwardly of the front heat shield to define a gap between the lip and the front heat shield. The front heat shield has a leading edge opposite the lip. The combustor has impingement holes extending through the shell and disposed to direct impingement cooling jets to the upstream portion of the front heat shield. The leading edge, of the front heat shield has at least one scallop defining an opening and disposed to allow the impingement cooling jets to impinge directly on a portion of the peripheral lip adjacent the scallop.

Abstract: An internally cooled airfoil for a gas turbine engine has a hollow airfoil body including pressure and suction sidewalls defining a cooling passage therebetween. A combination of pedestal and trip-strips are used in the cooling passage to enhance heat transfer while minimizing the coolant pressure drop across these features.

Abstract: A combustor heat shield comprises a heat shield body adapted to be mounted to a combustor wall with a back side of the heat shield body in spaced-apart facing relationship with the combustor wall to define an air gap between the heat shield body and the combustor wall. At least one nozzle opening is defined in the heat shield bod. The opening is bordered by a nozzle opening boss. The boss extends from the back side of the heat shield body across the air gap for sealing engagement with an adjacent part of the combustor. An annular array of effusion holes is provided adjacent the nozzle opening boss. The effusion holes extend through the heat shield body for passing cooling air from the back side to a front side of the heat shield body. Fins are interspersed between the effusion holes on the back side of the heat shield.

Abstract: A gas turbine engine combustor has a dome and a shell extending axially from the dome. The dome and the shell cooperates to define a combustion chamber. A dome heat shield is mounted to the dome inside the combustion chamber. A front heat shield is mounted to the shell inside the combustion chamber. The dome heat shield and the front heat shield have axially overlapping portions cooperating to define a flow guiding channel. The flow guiding channel has a length (L) and a height (h). The length (L) is at least equal to the height (h).

Abstract: A gas turbine engine airfoil has a hollow airfoil section extending chordwise between a leading edge and a trailing edge. The airfoil has a leading edge cooling passage and a separate serpentine passage for cooling a remaining portion of the airfoil. The serpentine passage has at least three segment serially interconnected in fluid flow communication. The leading edge cooling passage and the serpentine cooling passage have separate coolant inlets. The coolant inlet of the serpentine passage comprises a primary inlet branch connected in fluid flow communication with a first one of the segments of the serpentine passage and a secondary inlet branch connected in flow communication with a last one of the segments, thereby providing for a portion of the flow passing through the coolant inlet of the serpentine passage to be directly fed into the last segment of the serpentine passage.

Abstract: A gas turbine engine airfoil has a hollow airfoil section extending chordwise between a leading edge and a trailing edge. The airfoil has a leading edge cooling passage and a separate serpentine passage for cooling a remaining portion of the airfoil. The serpentine passage has at least three segment serially interconnected in fluid flow communication. The leading edge cooling passage and the serpentine cooling passage have separate coolant inlets. The coolant inlet of the serpentine passage comprises a primary inlet branch connected in fluid flow communication with a first one of the segments of the serpentine passage and a secondary inlet branch connected in flow communication with a last one of the segments, thereby providing for a portion of the flow passing through the coolant inlet of the serpentine passage to be directly fed into the last segment of the serpentine passage.

Abstract: A partially coated blade for a gas turbine engine, including a fillet surface surrounding the airfoil section and connecting it to the platform section. A radially outermost portion of the pressure side and leading edge is covered by a thermal barrier coating. This portion extends radially from a first limit to the blade tip. The first limit is located at a radial distance from the platform of at most 21% of the maximum span. The fillet surface is free or substantially free of the thermal barrier coating. In another embodiment, a second portion of the pressure side and of the leading edge is free or substantially free of the thermal barrier coating, extending radially from the platform section to a second limit located a radial distance from the platform section corresponding to at least 5% of the maximum span. A method of applying a thermal barrier coating is also discussed.

Abstract: A turbine blade for a gas turbine engine with an airfoil portion defined by a perimeter wall surrounding at least one enclosure, the perimeter wall having a plurality of cooling holes defined therethrough and providing fluid communication between the at least one enclosure and a gaspath of the gas turbine engine. The plurality of cooling holes includes at least one set of holes selected from the group consisting of a first set, a second set, a third set, a fourth set, a fifth set and a sixth set, wherein the first, second, third, fourth, fifth and sixth sets of holes respectively include the holes numbered A1 to A8, B1 to B10, C1 to C9, D1 to D6, E1 to E7 and F1 to F6 each located such that a central axis thereof extends through the respective point 1 and point 2 having a nominal location in accordance with the X, Y, Z Cartesian coordinate values set forth in Table 3.

Abstract: There is provided a combustor comprising a dome and a shell extending from the dome defining a combustion chamber. A dome heat shield is mounted to the dome inside the combustion chamber. A front heat shield is mounted to the shell and spaced therefrom. The dome heat shield has a lip extending generally away from the dome heat shield and generally parallel to the shell and spaced inwardly of the front heat shield to define a gap between the lip and the front heat shield. The front heat shield has a leading edge opposite the lip. The combustor has impingement holes extending through the shell and disposed to direct impingement cooling jets to the upstream portion of the front heat shield. The leading edge, of the front heat shield has at least one scallop defining an opening and disposed to allow the impingement cooling jets to impinge directly on a portion of the peripheral lip adjacent the scallop.

Abstract: A heat shield for a combustor of a gas turbine engine has a heat shield adapted to be mounted to a combustor wall with a back face of the heat shield in spaced-apart facing relationship with an inner surface of the combustor wall to define an air space. Concentric rails extend from the back face of the heat shield across the air space surrounding a nozzle opening in the heat shield. Effusion holes are provided between the concentric rails and extend between the back and front faces. Fins may be placed between the effusion holes.