Helical skin cooling passages for turbine airfoils

An airfoil for a gas turbine engine may comprise an airfoil body having an outer diameter (OD) end extending between a leading edge and a trailing edge and having an ID end located opposite the airfoil body from the OD end. The airfoil body defines a helical skin cooling passage extending between the ID end of the airfoil and the OD end of the airfoil. The airfoil body may further define a main body core. The helical skin cooling passage may thermally shield the main body core.

Directional cooling arrangement for airfoils

An airfoil according to an example of the present disclosure includes, among other things, an internal wall and an external wall. The external wall defines pressure and suction sides between a leading edge and a trailing edge, and the airfoil section defines a mean camber line that extends between the leading and trailing edges to bisect a thickness of the airfoil section. A first cavity and a second cavity are separated by the internal wall. The second cavity is bounded by the external wall at the leading edge. At least one crossover passage within the internal wall connects the first cavity to the second cavity. The crossover passage defines a passage axis. The passage axis defines a passage angle with respect to the mean camber line such that the passage axis extends transversely from the mean camber line to intersect a surface of the second cavity.

FLOAT WALL COMBUSTOR PANELS HAVING AIRFLOW DISTRIBUTION FEATURES

Combustor panel bodies with first and second sides, a pin array of cooling pins extending from the first side, with each pin extending a first height from the first side and having a pin diameter, each pin separated from adjacent pins by a pin array separation distance. A structural protrusion extending from the first side with no pins located at a position within a flashing distance that is equal to a protrusion separation distance plus one half of the pin diameter. A location of the pin is measured from a center point of the pin to a closest point on the exterior surface of the structural protrusion. A pin array scallop is integrally formed with the structural protrusion forming a reduction in material such that each pin of the array that is closest to the structural protrusion is positioned the pin array separation distance from the structural protrusion. 1. A combustor panel for use in a gas turbine engine comprising:a panel body having a first side and a second side;a plurality of cooling pins extending from the first side, the plurality of cooling pins arranged in a pin array, wherein each cooling pin extends a first height from the first side of the panel body, has a pin diameter, and is separated from adjacent cooling pins of the pin array by a pin array separation distance;at least one structural protrusion extending from the first side of the panel body,wherein no pins of the pin array are located at a position within a flashing distance that is equal to a protrusion separation distance plus one half of the pin diameter, wherein the protrusion separation distance is a predetermined minimum distance between an exterior surface of the at least one structural protrusion and an exterior surface of a cooling pin, and wherein a location of the pin is measured from a center point of the cooling pin to a closest point on the exterior surface of the at least one structural protrusion; andat least one pin array scallop integrally formed with the at least one ...

Directional cooling arrangement for airfoils

An airfoil according to an example of the present disclosure includes, among other things, an internal wall and an external wall. The external wall defines pressure and suction sides between a leading edge and a trailing edge, and the airfoil section defines a mean camber line that extends between the leading and trailing edges to bisect a thickness of the airfoil section. A first cavity and a second cavity are separated by the internal wall. The second cavity is bounded by the external wall at the leading edge. At least one crossover passage within the internal wall connects the first cavity to the second cavity. The crossover passage defines a passage axis. The passage axis defines a passage angle with respect to the mean camber line such that the passage axis extends transversely from the mean camber line to intersect a surface of the second cavity.

WOVEN SKIN CORES FOR TURBINE AIRFOILS

An airfoil includes an airfoil body having a pressure side and a suction side that each radially extend between a first radial boundary end and a second radial boundary end and each axially extend between a leading edge and a trailing edge. The airfoil body defines a plurality of first skin core passages disposed proximate the suction side and radially extend from the first radial boundary end towards the second radial boundary end, and a plurality of second skin core passages that radially extend toward the second radial boundary end and circumferentially extend towards the suction side proximate the second radial boundary end. 1. An airfoil for a gas turbine engine , comprising:an airfoil body having a pressure side and a suction side that each radially extend between a first radial boundary end and a second radial boundary end and each axially extend between a leading edge and a trailing edge, a plurality of first skin core passages disposed proximate the suction side and radially extend from the first radial boundary end towards the second radial boundary end, and', 'a plurality of second skin core passages that radially extend toward the second radial boundary end and circumferentially extend towards the suction side proximate the second radial boundary end., 'the airfoil body defining2. The airfoil of claim 1 , wherein proximate the second radial boundary end claim 1 , a first skin core passage of the plurality of first skin core passages is axially disposed between a first and second skin core passage of the plurality of second skin core passages.3. The airfoil of claim 1 , wherein the airfoil body defines a first skin core inlet passage that is disposed proximate the first radial boundary end and the suction side and the plurality of first skin core passages extend from the first skin core inlet passage.4. The airfoil of claim 3 , wherein the airfoil body defines a second skin core inlet passage that is disposed proximate the first radial boundary end and is ...

Turbine Blade Trailing Edge Cooling Feed

A turbine blade has an attachment root and an airfoil. A cooling passageway system has a plurality of trunks extending from respective inlets along the root inner diameter end from a leading trunk near a first axial end to a trailing trunk near a second axial end; and a plurality of outlets along the airfoil including trailing edge outlets fed by the trailing trunk. Viewed normal to a root end-to-end centerplane: the trailing trunk has a turn passing forward and then rearward; an outside of the turn protrudes forward; and the outside of the turn has a tighter curvature than an inside of the turn.

FLUID TUBE ASSEMBLY FOR GAS TURBINE ENGINE

A fluid tube assembly includes a first fluid tube extending through a duct liner, a portion of the first fluid tube disposed radially outward of the duct liner and a portion of the first fluid tube disposed radially inward of the duct liner. Also included is a second fluid tube located radially inward of the duct liner. Further included is a clamp located radially inward of the duct liner and having an inner wall defining an interior space. Yet further included is a nut at least partially disposed within the interior space, the nut having an outer geometry corresponding to a geometry of a portion of the inner wall of the clamp, the first fluid tube and the second fluid tube each having a portion disposed within the nut, the clamp and the nut providing anti-rotation during assembly of the fluid tube assembly. 1. A fluid tube assembly comprising:a first fluid tube extending through a duct liner, a portion of the first fluid tube disposed radially outward of the duct liner and a portion of the first fluid tube disposed radially inward of the duct liner;a second fluid tube located radially inward of the duct liner;a clamp located radially inward of the duct liner and having an inner wall defining an interior space; anda nut at least partially disposed within the interior space, the nut having an outer geometry corresponding to a geometry of a portion of the inner wall of the clamp, the first fluid tube and the second fluid tube each having a portion disposed within the nut, the clamp and the nut providing anti-rotation during assembly of the fluid tube assembly.2. The fluid tube assembly of claim 1 , further comprising a clamp annular ring extending inwardly from the inner wall of the clamp claim 1 , the first fluid tube claim 1 , the second fluid tube claim 1 , and the nut free to move along a longitudinal direction of the first and second fluid tubes claim 1 , the annular ring defining a movement limit for the first fluid tube claim 1 , the second fluid tube claim 1 , ...

Impingement cooling arrangement for airfoils

An airfoil for a gas turbine engine according to an example of the present disclosure includes, among other things, an airfoil section that has an internal wall and an external wall. The external wall defines pressure and suction sides that extends in a chordwise direction between a leading edge and a trailing edge, a first impingement cavity and a second impingement cavity bounded by the external wall at a leading edge region that defines the leading edge. A first crossover passage within the internal wall is connected to the first impingement. The first crossover passage defines a first passage axis that intersects a surface of the first impingement cavity. A second crossover passage within the internal wall is connected to the second impingement cavity. The second crossover passage defines a second passage axis that intersects a surface of the second impingement cavity.

COMBUSTOR LINER PANEL END RAIL WITH DIFFUSED INTERFACE PASSAGE FOR A GAS TURBINE ENGINE COMBUSTOR

A liner panel for use in a combustor of a gas turbine engine, the liner panel including a cold side; and a rail that extends from the cold side, the rail includes a first diffusion interface passage surface and a second diffusion interface passage surface, the first diffusion interface passage surface angled with respect to the second diffusion interface passage surface. 120-. (canceled)21. A liner panel for use in a combustor of a gas turbine engine , the liner panel comprising:a rail that extends from a cold side, such that an end section thereof is adjacent to a support shell that supports the liner panel, the rail forms a first aft diffusion interface passage surface of a pre-diffuser section, and a second aft diffusion interface passage surface of a diffuser section, the first aft diffusion interface passage surface angled relative to the second aft diffusion interface passage surface, the diffuser section of the rail defines a height with respect to the cold side between 0.05 H-0.6 H, where H is the height of the rail and the diffuser section at least partially defining an expanding passage that extends at an angle between 1-15 degrees with respect to the first aft diffusion interface passage surface of the pre-diffuser section.22. The liner panel as recited in claim 21 , wherein the rail forms a portion of a periphery rail that defines an edge of the liner panel to abut the support shell.23. The liner panel as recited in claim 21 , wherein the rail includes a chiseled end section to at least partially form the diffusion interface passage.24. The liner panel as recited in claim 21 , wherein the rail is trapezoidal shaped.25. The liner panel as recited in claim 21 , wherein the rail is ramp shaped.26. A combustor for a gas turbine engine comprising:a support shell;a forward liner panel mounted to the support shell, the forward liner panel comprising an aft rail that extends from a cold side of the forward liner panel to a aft rail that contacts the support ...

Internally cooled engine components

A component of a gas turbine engine is provided. The component may include an external surface with a skin cavity adjacent the external surface. The skin cavity may be defined by a hot-side surface and a cool-side surface with the hot-side surface between the cool-side surface and the external surface. A flat portion may be disposed on the hot-side surface and may, for example, define a wall-thickness check region. A protrusion may be disposed on the cool-side surface opposite the flat portion on the hot-side surface. The hot-side surface and the cool-side surface may tend to increase a heat transfer coefficient along the flat portion of the hot-side surface, which may also define a wall thickness check region.

Impingement cooling arrangement for airfoils

An airfoil for a gas turbine engine according to an example of the present disclosure includes, among other things, an airfoil section that has an internal wall and an external wall. The external wall defines pressure and suction sides that extends in a chordwise direction between a leading edge and a trailing edge, a first impingement cavity and a second impingement cavity bounded by the external wall at a leading edge region that defines the leading edge, a first feeding cavity separated from the first impingement cavity and from the second impingement cavity by the internal wall, and a first crossover passage within the internal wall that connects the first impingement cavity and the first feeding cavity. The first crossover passage defines a first passage axis that intersects a surface of the first impingement cavity. A second crossover passage within the internal wall connects to the second impingement cavity. The second crossover passage defines a second passage axis that intersects a surface of the second impingement cavity.

GAS TURBINE ENGINE COMPONENT WITH DISCHARGE SLOT HAVING A FLARED BASE

A component for a gas turbine engine includes a body portion that extends between a leading edge and a trailing edge of the component. The trailing edge includes a flared region and a non-flared region. At least one discharge slot is disposed at least partially within the flared region of the component. 1. A component for a gas turbine engine , comprising:a body portion that extends between a leading edge and a trailing edge, wherein said trailing edge includes a cooling circuit having a flared region and a non-flared region the trailing edge further including at least one discharge slot, and a first discharge slot of the at least one discharge slot is disposed at least partially within the flared region of the cooling circuit;and wherein the cooling circuit defines a first width in the flared region and a second width in the non-flared region, the first width being larger than the second width.2. The component as recited in claim 1 , wherein the at least one discharge slot has a first portion that includes an oval geometry.3. The component as recited in claim 1 , wherein said component is one of a vane and a blade.4. The component as recited in claim 1 , wherein said at least one discharge slot includes a plurality of second discharge slots.5. The component as recited in claim 1 , further comprising at least a second discharge slot disposed radially outward of the first discharge slot.6. The component as recited in claim 1 , further comprising at least a second discharge slot disposed radially inward of the first discharge slot.7. The component as recited in claim 1 , wherein a thickness of the flared region increases as the trailing edge is approached.8. (canceled)9. The component as recited in claim 1 , wherein at least a portion of the flared region is disposed within a filleted section of the body portion.10. The component as recited in claim 1 , wherein said at least one discharge slot is positioned at a position that is immediately adjacent to a platform of ...

COMBUSTOR LINER PANEL END RAIL WITH CURVED INTERFACE PASSAGE FOR A GAS TURBINE ENGINE COMBUSTOR

A combustor for a gas turbine engine includes a support shell; a first liner panel mounted to the support shell via a multiple of studs, the first liner panel including a first rail that extends from a cold side of the first liner panel such that the rail is non-perpendicular to the cold side and includes a concave surface to at least partially form a curved interface passage; and a second liner panel mounted to the support shell via a multiple of studs, the first liner panel including a second rail that extends from a cold side of the second liner panel and includes a convex surface to at least partially form the curved interface passage. 1. A liner panel for use in a combustor of a gas turbine engine , the liner panel comprising:a cold side; anda rail that extends from the cold side such that the rail is non-perpendicular to the cold side.2. The liner panel as recited in claim 1 , wherein the rail is curved.3. The liner panel as recited in claim 2 , wherein the rail is curved along a single radius.4. The liner panel as recited in claim 2 , wherein the rail is curved along a multiple of radii.5. The liner panel as recited in claim 1 , wherein the liner panel is at least one of a forward liner panel claim 1 , and an aft liner panel.6. The liner panel as recited in claim 1 , wherein the rail is a periphery rail that defines an end of the liner panel.7. A combustor for a gas turbine engine comprising:a support shell;a first liner panel mounted to the support shell via a multiple of studs, the first liner panel including a first rail that extends from a cold side of the first liner panel, the first rail is non-perpendicular to the cold side and includes a concave surface to at least partially form a curved interface passage; anda second liner panel mounted to the support shell via a multiple of studs, the second liner panel including a second rail that extends from a cold side of the second liner panel, the second rail includes a convex surface to at least partially ...

COMBUSTOR LINER PANEL END RAIL ANGLED COOLING INTERFACE PASSAGE FOR A GAS TURBINE ENGINE COMBUSTOR

A combustor for a gas turbine engine includes a support shell with an angled shell interface; a first liner panel mounted to the support shell via a multiple of studs, the first liner panel including a first rail that extends from a cold side of the first liner panel adjacent to the angled shell interface; and a second liner panel mounted to the support shell via a multiple of studs, the second liner panel including a second rail that extends from a cold side of the second liner panel adjacent to said first rail, the second rail adjacent to the angled shell interface. 1. A liner panel for use in a combustor of a gas turbine engine , the liner panel comprising:a cold side; anda rail that extends from the cold side, the rail includes an angled surface with respect to the cold side.2. The liner panel as recited in claim 1 , wherein the angled surface with respect to the cold side is non-perpendicular.2. The liner panel as recited in claim 1 , wherein the angled surface is between about 30-60 degrees.4. The liner panel as recited in claim 1 , wherein the liner panel is at least one of a forward liner panel claim 1 , and an aft liner panel.5. The liner panel as recited in claim 1 , wherein the angled surface is a forward surface of a forward rail of an aft liner panel.6. The liner panel as recited in claim 1 , wherein the angled surface is an angled surface of an aft rail of a forward liner panel.7. A combustor for a gas turbine engine comprising:a support shell with an angled shell interface;a first liner panel mounted to the support shell via a multiple of studs, the first liner panel including a first rail that extends from a cold side of the first liner panel adjacent to the angled shell interface; anda second liner panel mounted to the support shell via a multiple of studs, the second liner panel including a second rail that extends from a cold side of the second liner panel adjacent to said first rail, the second rail adjacent to the angled shell interface.8. The ...

COMBUSTOR LINER PANEL END RAIL COOLING ENHANCEMENT FEATURES FOR A GAS TURBINE ENGINE COMBUSTOR

A combustor for a gas turbine engine includes a support shell; a first liner panel mounted to the support shell via a multiple of studs, the first liner panel including a first rail that extends from a cold side of the first liner panel; a second liner panel mounted to the support shell via a multiple of studs, the second liner panel including a second rail that extends from a cold side of the second liner panel adjacent to the first rail to form an interface passage; and at least one heat transfer feature within the interface passage. 1. A liner panel for use in a combustor of a gas turbine engine , the liner panel comprising:a cold side; anda rail that extends from the cold side, the rail includes a surface with at least one heat transfer feature.2. The liner panel as recited in claim 1 , wherein the rail at least partially forms an interface passage.3. The liner panel as recited in claim 2 , wherein the rail is angled with respect to the cold side.4. The liner panel as recited in claim 1 , wherein the liner panel is at least one of a forward liner panel claim 1 , and an aft liner panel.5. The liner panel as recited in claim 1 , wherein the rail is a forward rail of an aft liner panel.6. The liner panel as recited in claim 1 , wherein the rail is an aft rail of a forward liner panel.7. The liner panel as recited in claim 1 , wherein the rail is a periphery rail.8. The liner panel as recited in claim 1 , wherein the heat transfer features are arranged in rows.9. The liner panel as recited in claim 1 , wherein the heat transfer features are shaped as at least one of pins claim 1 , circles claim 1 , ovals claim 1 , and racetracks.10. A combustor for a gas turbine engine comprising:a support shell;a first liner panel mounted to the support shell via a multiple of studs, the first liner panel including a first rail that extends from a cold side of the first liner panel;a second liner panel mounted to the support shell via a multiple of studs, the second liner panel ...

COMBUSTOR LINER PANEL END RAIL COOLING INTERFACE PASSAGE FOR A GAS TURBINE ENGINE COMBUSTOR

A combustor for a gas turbine engine includes a support shell; a first liner panel mounted to the support shell via a multiple of studs, the first liner panel including a first rail that extends from a cold side of the first liner panel; and a second liner panel mounted to the support shell via a multiple of studs, the second liner panel including a second rail that extends from a cold side of the second liner panel adjacent to said first rail, the second rail includes a discontinuous rail end surface. 1. A liner panel for use in a combustor of a gas turbine engine , the liner panel comprising:a cold side; anda rail that extends from the cold side, the rail includes a discontinuous rail end surface.2. The liner panel as recited in claim 1 , wherein the rail is non-perpendicular to the cold side.3. The liner panel as recited in claim 1 , wherein the rail is angled with respect to the cold side.4. The liner panel as recited in claim 1 , wherein the liner panel is at least one of a forward liner panel claim 1 , and an aft liner panel.5. The liner panel as recited in claim 1 , wherein the rail is a forward rail of an aft liner panel.6. The liner panel as recited in claim 1 , wherein the rail is an aft rail of a forward liner panel.7. The liner panel as recited in claim 1 , wherein the rail is a periphery rail.8. The liner panel as recited in claim 1 , wherein the discontinuous rail end surface includes a circular passage.9. The liner panel as recited in claim 1 , wherein the discontinuous rail end surface includes a semi-circular passage.10. The liner panel as recited in claim 1 , wherein the discontinuous rail end surface includes an oval passage.11. The liner panel as recited in claim 1 , wherein the discontinuous rail end surface includes a “castle” shaped passage.12. A combustor for a gas turbine engine comprising:a support shell;a first liner panel mounted to the support shell via a multiple of studs, the first liner panel including a first rail that extends from a ...

TURBINE BLADE WITH TIP VORTEX CONTROL AND TIP SHELF

A blade for a gas turbine engine includes an airfoil which includes pressure and suction side surfaces joined at leading and trailing edges to provide an exterior airfoil surface. The airfoil extends from a 0% span position to a 100% span position at a tip. A mean chamber line is provided at the tip and extends from the leading edge to the trailing edge. The mean chamber line has a leading edge tangent line at the leading edge at the tip and a trailing edge tangent line at the trailing edge at the tip. The leading edge tangent line and the trailing edge tangent line intersect to provide a total camber angle in a range of −35° to 35°. The pressure side includes a tip shelf. 1. A blade for a gas turbine engine , comprising:an airfoil including pressure and suction side surfaces joined at leading and trailing edges to provide an exterior airfoil surface, the airfoil extending from a 0% span position to a 100% span position at a tip, a mean chamber line is provided at the tip and extends from the leading edge to the trailing edge, the mean chamber line has a leading edge tangent line at the leading edge at the tip and a trailing edge tangent line at the trailing edge at the tip, the leading edge tangent line and the trailing edge tangent line intersect to provide a total camber angle in a range of −35° to 35°, wherein the pressure side includes a tip shelf.2. The blade of claim 1 , wherein the tip shelf wraps around the leading edge and includes cooling holes.3. The blade of claim 2 , wherein the cooling holes include a forward-most hole nearest the leading edge claim 2 , the forward-most hole arranged aft of the leading edge.4. The blade of claim 3 , wherein the airfoil includes a stagnation point at the tip shelf aft of the leading edge on the pressure side claim 3 , the forward-most hole arranged between the leading edge and the stagnation point.5. The blade of claim 2 , wherein the tip shelf includes a shelf radius near the leading edge claim 2 , and the leading ...

AIRFOILS HAVING TAPERED TIP FLAG CAVITY AND CORES FOR FORMING THE SAME

Core assemblies for manufacturing airfoils and airfoils for gas turbine engines are described. The core assemblies include a tip flag cavity core having an upstream portion, a tapering portion, and a downstream portion, with the tapering portion located between the upstream portion and the downstream portion and the downstream portion defines an exit in a formed airfoil. The upstream portion has a first radial height H, the downstream portion has a second radial height Hthat is less than the first radial height H, the tapering portion transitions from the first radial height Hat an upstream end to the second radial height Hat a downstream end, and at least one metering pedestal aperture is located within the tapering portion.

COMBUSTOR LINER PANEL END RAIL MATCHING HEAT TRANSFER FEATURES

A combustor section of a turbine engine includes a first liner panel including a first end rail. The first end rail includes a protruding heat transfer feature. A second liner panel includes a second end rail disposed proximate the first end rail. The second end rail includes a recess matching the protruding heat transfer feature of the first end rail. A turbine engine and a method of assembling a combustor assembly of a gas turbine engine are also disclosed. 1. A combustor section of a turbine engine comprising:a first liner panel including a first end rail, the first end rail including a protruding heat transfer feature;a second liner panel including a second end rail disposed proximate the first end rail, the second end rail including a recess matching the protruding heat transfer feature of the first end rail.2. The combustor section as recited in claim 1 , wherein the first end rail and the second end rail define an interface between the first liner panel and the second liner panel.3. The combustor section as recited in claim 2 , wherein the interface extends in a direction parallel to a combustor longitudinal axis.4. The combustor section as recited in claim 2 , wherein the interface extends in a direction transverse to a combustor longitudinal axis.5. The combustor section as recited in claim 1 , wherein the first end rail and the second end rail are angled relative to a hot side of respective ones of the first liner panel and the second liner panel.6. The combustor section as recited in claim 1 , wherein the protruding heat transfer feature of the first end rail fits at least partially within the recess on the second end rail.7. The combustor section as recited in claim 1 , wherein the first end rail is disposed at a periphery of the first liner panel and the second end rail is disposed at a periphery of the second liner panel.8. The combustor section as recited in claim 7 , wherein each of the first end rail and the second end rail space the corresponding ...

COMBUSTOR PANEL ENDRAIL INTERFACE

A combustor of a gas turbine engine may include a combustor shell, a first combustor panel coupled to the combustor shell, and a second combustor panel coupled to the combustor shell. The first combustor panel may have a first endrail and the second combustor panel may have a second endrail. An annular cooling cavity may be defined between the combustor shell and the first and second combustor panels and a channel may be defined between the first endrail and the second endrail, wherein direct line-of-sight through the channel from the annular cooling cavity to a combustor chamber is obstructed. Said differently, the interface between the adjacent endrails may be non-linear, in a direction from the annular cooling cavity to the combustor chamber. 1. A combustor of a gas turbine engine , the combustor defining a combustor chamber and comprising:a first combustor panel comprising a first endrail; anda second combustor panel comprising a second endrail;wherein an annular cooling cavity is defined at least partially by the first and second combustor panels and a channel is defined between the first endrail and the second endrail, wherein direct line-of-sight through the channel from the annular cooling cavity to the combustor chamber is obstructed.2. The combustor of claim 1 , further comprising a combustor shell claim 1 , wherein the first combustor panel and the second combustor panel are coupled to the combustor shell.3. The combustor of claim 1 , wherein the first endrail and the second endrail have complementary geometries that form a shiplap interface.4. The combustor of claim 1 , wherein a centerline axis of the channel extending from the annular cooling cavity to the combustor chamber is non-linear.5. The combustor of claim 4 , wherein the channel comprises a first bend and a second bend.6. The combustor of claim 5 , wherein a first minor angle of the first bend is about 90 degrees and a second minor angle of the second bend is greater than about 90 degrees.7. ...

REDUNDANT ENDRAIL FOR COMBUSTOR PANEL

A combustor panel may include a body having an edge, a first endrail disposed along the edge of the body, and a second endrail disposed adjacent the first endrail, wherein a volume is defined between the first endrail and the second endrail. The first endrail may have a first height, measured from the body to a first tip of the first endrail, and the second endrail may have a second height, measured from the body to a second tip of the second endrail, wherein the first height is the same as the second height. The second endrail may include an endrail hole extending through the second endrail. 1. A combustor panel comprising:a body comprising an edge;a first endrail disposed along the edge of the body; anda second endrail disposed adjacent the first endrail, wherein a gap is defined between the first endrail and the second endrail.2. The combustor panel of claim 1 , wherein the first endrail has a first height claim 1 , measured from the body to a first tip of the first endrail claim 1 , and the second endrail has a second height claim 1 , measured from the body to a second tip of the second endrail claim 1 , wherein the first height is substantially the same as the second height.3. The combustor panel of claim 1 , wherein the second endrail defines an endrail hole extending through the second endrail.4. The combustor panel of claim 3 , wherein the endrail hole is canted relative to the combustor panel.5. The combustor panel of claim 3 , wherein the first endrail is solid and has a continuous surface.6. The combustor panel of claim 1 , wherein a distance between the first endrail and the second endrail is between about 0.05 inches and about 1.0 inch.7. The combustor panel of claim 1 , wherein a distance between the first endrail and the second endrail is between about 0.10 inches and about 0.50 inches.8. A combustor of a gas turbine engine claim 1 , the combustor comprising:a combustor shell; and a body comprising an edge;', 'a first endrail disposed along the edge ...

COMBUSTOR LINER PANEL END RAIL WITH DIFFUSED INTERFACE PASSAGE FOR A GAS TURBINE ENGINE COMBUSTOR

A liner panel for use in a combustor of a gas turbine engine, the liner panel including a cold side; and a rail that extends from the cold side, the rail includes a first diffusion interface passage surface and a second diffusion interface passage surface, the first diffusion interface passage surface angled with respect to the second diffusion interface passage surface. 1. A liner panel for use in a combustor of a gas turbine engine , the liner panel comprising:a cold side; anda rail that extends from the cold side, the rail includes a first diffusion interface passage surface and a second diffusion interface passage surface, the first diffusion interface passage surface angled with respect to the second diffusion interface passage surface.2. The liner panel as recited in claim 1 , wherein the first diffusion interface passage surface at least partially defines a pre-diffuser section and the second diffusion interface passage surface at least partially defines a diffuser section along one side of a diffusion interface passage axis.3. The liner panel as recited in claim 2 , wherein the diffuser section of the rail defines a height with respect to the cold side between 0.05 H-0.6 H claim 2 , where H is the height of the rail.4. The liner panel as recited in claim 3 , wherein the diffuser section at least partially defines an expanding passage that extends at an angle between 1-15 degrees with respect to a surface that forms the pre-diffuser section.5. The liner panel as recited in claim 1 , wherein the liner panel is at least one of a forward liner panel claim 1 , and an aft liner panel.6. The liner panel as recited in claim 1 , wherein the rail is a periphery rail that defines an edge of the liner panel.7. The liner panel as recited in claim 1 , wherein the rail includes a ramped end section.8. The liner panel as recited in claim 1 , wherein the rail is trapezoidal shaped.9. The liner panel as recited in claim 1 , wherein the rail is ramp shaped.10. A combustor for ...

Combustor liner panel end rail cooling interface passage for a gas turbine engine combustor

A combustor for a gas turbine engine includes a support shell; a first liner panel mounted to the support shell via a multiple of studs, the first liner panel including a first rail that extends from a cold side of the first liner panel; and a second liner panel mounted to the support shell via a multiple of studs, the second liner panel including a second rail that extends from a cold side of the second liner panel adjacent to said first rail, the second rail includes a discontinuous rail end surface.

Combustor liner panel end rail with curved interface passage for a gas turbine engine combustor

A combustor for a gas turbine engine includes a support shell; a first liner panel mounted to the support shell via a multiple of studs, the first liner panel including a first rail that extends from a cold side of the first liner panel such that the rail is non-perpendicular to the cold side and includes a concave surface to at least partially form a curved interface passage; and a second liner panel mounted to the support shell via a multiple of studs, the first liner panel including a second rail that extends from a cold side of the second liner panel and includes a convex surface to at least partially form the curved interface passage.

Combustor liner panel end rail angled cooling interface passage for a gas turbine engine combustor

A combustor for a gas turbine engine includes a support shell with an angled shell interface; a first liner panel mounted to the support shell via a multiple of studs, the first liner panel including a first rail that extends from a cold side of the first liner panel adjacent to the angled shell interface; and a second liner panel mounted to the support shell via a multiple of studs, the second liner panel including a second rail that extends from a cold side of the second liner panel adjacent to said first rail, the second rail adjacent to the angled shell interface.

Combustor liner panel end rail cooling enhancement features for a gas turbine engine combustor

A combustor for a gas turbine engine includes a support shell; a first liner panel mounted to the support shell via a multiple of studs, the first liner panel including a first rail that extends from a cold side of the first liner panel; a second liner panel mounted to the support shell via a multiple of studs, the second liner panel including a second rail that extends from a cold side of the second liner panel adjacent to the first rail to form an interface passage; and at least one heat transfer feature within the interface passage.

Combustor liner panel end rail cooling interface passage for a gas turbine engine combustor

A combustor for a gas turbine engine includes a support shell (68, 70), a first liner panel (72A, 72B) mounted to the support shell (68, 70) via a multiple of studs, the first liner panel (72A, 72B) including a first rail (124a) that extends from a cold side of the first liner panel (72A, 72B), and a second liner panel (74A, 74B) mounted to the support shell (68, 70) via a multiple of studs, the second liner panel (74A, 74B) including a second rail (122b) that extends from a cold side of the second liner panel (74A, 74B) adjacent to said first rail (124a), the second rail (122b) includes a discontinuous rail end surface (150).

Gas turbine engine component with a trailing edge discharge slot

Airfoils and corresponding method of manufacturing

An airfoil (100; 300) for a gas turbine engine (20) may comprise an airfoil body (120) having an outer diameter (OD) end (142; 342) extending between a leading edge (124; 324) and a trailing edge (126; 326) and having an ID end (132) located opposite the airfoil body (120) from the OD end (142; 342). The airfoil body (120) defines a helical skin cooling passage (158; 358) extending between the ID end (132) of the airfoil (100; 300) and the OD end (142; 342) of the airfoil (100; 300). The airfoil body (120) may further define a main body core (154; 354). The helical skin cooling passage (158; 358) may thermally shield the main body core (154; 354).

Combustor liner panel and rail with diffused interface passage for a gas turbine engine combustor

A liner panel for use in a combustor of a gas turbine engine, the liner panel including a cold side; and a rail that extends from the cold side, the rail includes a first diffusion interface passage surface and a second diffusion interface passage surface, the first diffusion interface passage surface angled with respect to the second diffusion interface passage surface.

Mateface for blade outer air seals in a gas turbine engine

A gas turbine engine includes one of a turbine section and a compressor section having multiple stages. At least one of the stages defines an outer diameter comprised of a plurality of circumferentially arranged blade outer air seals. Each blade outer air seal is spaced from each adjacent blade outer air seal in the plurality of circumferentially arranged blade outer air seals via a mateface gap. The mateface gap is oblique to a radius of the gas turbine engine, such that air entering the mateface gap is directed to an inner diameter surface of at least one of the blade outer air seals in the plurality of blade outer air seals.

Combustor liner panel assembly and method for cooling the same

Gas turbine engine and cooling method

Turbine blade trailing edge cooling feed

A turbine blade has an attachment root and an airfoil. A cooling passageway system has a plurality of trunks extending from respective inlets along the root inner diameter end from a leading trunk near a first axial end to a trailing trunk near a second axial end; and a plurality of outlets along the airfoil including trailing edge outlets fed by the trailing trunk. Viewed normal to a root end-to-end centerplane: the trailing trunk has a turn passing forward and then rearward; an outside of the turn protrudes forward; and the outside of the turn has a tighter curvature than an inside of the turn.

Turbine blade trailing edge cooling feed

A turbine blade has an attachment root and an airfoil. A cooling passageway system has a plurality of trunks extending from respective inlets along the root inner diameter end from a leading trunk near a first axial end to a trailing trunk near a second axial end; and a plurality of outlets along the airfoil including trailing edge outlets fed by the trailing trunk. Viewed normal to a root end-to-end centerplane: the trailing trunk has a turn passing forward and then rearward; an outside of the turn protrudes forward; and the outside of the turn has a tighter curvature than an inside of the turn.

Blade for a gas turbine engine with a tip shelf

Turbine blade trailing edge cooling feed and corresponding using method

Turbine blade trailing edge cooling feed

A turbine blade has an attachment root and an airfoil. A cooling passageway system has a plurality of trunks extending from respective inlets along the root inner diameter end from a leading trunk near a first axial end to a trailing trunk near a second axial end; and a plurality of outlets along the airfoil including trailing edge outlets fed by the trailing trunk. Viewed normal to a root end-to-end centerplane: the trailing trunk has a turn passing forward and then rearward; an outside of the turn protrudes forward; and the outside of the turn has a tighter curvature than an inside of the turn.

Impingement cooling arrangement for airfoils

Internally cooled gas turbine engine components

Blade outer air seal including cooling trench

A gas turbine engine includes a compressor, a combustor fluidly connected to the compressor via a core flowpath, and a turbine fluidly connected to the combustor via the core flowpath. The turbine includes at least one stage having a plurality of rotors and a plurality of vanes. An outer diameter of the core flowpath at at least one stage is at least partially defined by a set of circumferentially arranged blade outer air seals. Each blade outer air seal includes a platform. An internal cooling cavity is defined within the platform. At least one mateface of the platform includes a cooling trench, and a first set of cooling holes connecting the internal cavity to the cooling trench.

Internally cooled engine components

A component of a gas turbine engine (20) is provided. The component may include an external surface with a skin cavity (120) adjacent the external surface. The skin cavity (120) may be defined by a hot-side surface (124) and a cool-side surface (122) with the hot-side surface (124) between the cool-side surface (122) and the external surface. A flat portion may be disposed on the hot-side surface (124) and may, for example, define a wall-thickness check region (144). A protrusion (152) may be disposed on the cool-side surface (122) opposite the flat portion on the hot-side surface (124). The hot-side surface (124) and the cool-side surface (122) may tend to increase a heat transfer coefficient along the flat portion of the hot-side surface (124), which may also define a wall thickness check region.

Gas turbine combustor with liner panel end rails

Combustor liner panel end rail with curved interface passage for a gas turbine engine combustor

Turbine airfoil having cooling hole arrangement

A turbine airfoil includes a body that has inner and outer platforms and an airfoil section that extends between the inner and outer platforms. Cooling holes define external breakout points from the body. The external breakout points are located in accordance with Cartesian coordinates of at least points 1 through 11, 193 through 201, and 268 through 275 set forth in Table 1 herein.