Abstract: A heat shield panel for a gas turbine engine combustor may comprise a perimetrical rail extending at least partially around a perimeter of the heat shield panel and defining a main backside cavity, a plurality of air admittance holes, each air admittance hole extending along a hole axis, and a midrail dividing the main backside cavity into a first backside cavity and a second backside cavity, wherein at least a portion of the midrail is axially offset from at least one of the hole axes.

Abstract: A combustor panel may comprise a surface extending from a first end of the combustor panel to a second end of the combustor panel opposite the first end. A first rail located proximate the first end may extend from the surface of the combustor panel. The first rail may define a plurality of channels. The first rail may be formed a preselected distance away from the first end of the combustor panel.

Abstract: A gas-turbine engine is provided comprising a high-pressure turbine, a low-pressure turbine aft of the high-pressure turbine, and a housing around the low-pressure turbine and the high-pressure turbine. A stator may be disposed between the low-pressure turbine and the high-pressure turbine with a boss formed on the stator. A bushing may be retained in the boss with a pin coupled to the housing and partially within the bushing. A method of mounting a full-ring stator is also provided, the method including the steps of inserting a bushing into a boss of the full-ring stator by press fitting, and inserting a pin into the bushing of the full-ring stator.

Abstract: A disc of a gas turbine engine system with particular retaining ring placement and engagement implementation are provided. For example, the disc includes a disc bore including, a groove formed on an axially extending surface of the disc bore, wherein the groove includes a forward surface that extends radially into the disc bore to a groove floor that is cut into the disc bore, and wherein the groove extends axially to an aft surface that extends radially outward to at least the axially extending surface of the disc bore, and a disc web that extends radially outward from the disc bore, relative to an axis of rotation of the gas turbine engine.

Abstract: A heat shield panel for use in a gas turbine engine combustor is disclosed. In various embodiments, the heat shield panel includes a hot side, a cold side spaced from the hot side, a rail member disposed on the cold side proximate an outer perimeter, the rail member having an outer wall and an inner wall and an orifice extending through the rail member, from the inner wall to the outer wall, the orifice having an entrance portion having an entrance opening positioned on the inner wall and extending at least to an intermediate portion of the rail member and an exit portion having an exit opening positioned on the outer wall and extending at least to the intermediate portion of the rail member, the entrance portion of the orifice being angled relative to the exit portion of the orifice.

Abstract: A disc of a gas turbine engine system is provided. The disc includes a disc bore including, an interstage coupling disposed on an axially extending surface at a peripheral edge of the disc bore that includes a protrusion that extends axially beyond an outer surface of the disc bore, and a groove formed on the axially extending surface of the disc bore wherein an aft surface of the groove is also a forward surface of the interstage coupling, wherein the groove includes a forward surface that extends radially into the disc bore to a groove floor that is cut into the disc bore and extends axially to the aft surface that extends radially outward to at least the axially extending surface of the disc bore, and a disc web that extends radially outward from the disc bore, relative to an axis of rotation of the gas turbine engine.

Abstract: A heat shield panel for use in a gas turbine engine combustor is disclosed. The heat shield panel includes a hot side, a cold side and at least one attachment mechanism having a stud and a central axis extending through the stud and a plurality of standoff pins positioned circumferentially around the stud, the standoff pins having a radial extent, a circumferential extent that is greater than the radial extent, a radially outer surface having a radially convex shape and a radially inner surface having a radially concave shape.

Abstract: A heat shield panel for a gas turbine engine combustor is disclosed. The heat shield panel includes a hot side defining a first surface having an outer perimeter, a cold side defining a second surface spaced from the first surface and a plurality of holes, each hole including a central axis having vector components defined by a common vector.

Abstract: A heat shield panel for a gas turbine engine combustor is disclosed. The heat shield panel includes a hot side defining a first surface having an outer perimeter, a cold side defining a second surface spaced from the first surface, a rail member disposed on the cold side proximate a first portion of the outer perimeter, the rail member having an outer wall and an inner wall, an undercut portion within the second surface proximate the inner wall of the rail member and an orifice extending through the rail member, from the inner wall to the outer wall.

Abstract: A mid-turbine frame for a gas turbine engine includes an inner frame case that includes a bolt opening and at least one spoke for connecting an outer frame case to the inner frame case that includes an inlet passage that extends in a radial direction. A central bolt extends through the bolt opening for securing at least one spoke to the inner frame case.

Abstract: A minidisk for a rotor system may comprise a balance flange defining a hole array, which may include a first hole having a first width. The first hole may be configured to receive a balance weight. A second hole and a third hole may have a second width. The second hole and the third hole may be disposed adjacent to the first hole. The second width may be greater than the first width.

Abstract: A vane pack for a gas turbine engine includes an annular arrangement of vanes. A ring is secured around the vanes and extends proud of an axial end of the vanes.

Abstract: A disc of a gas turbine engine system with particular retaining ring placement and engagement implementation are provided. For example, the disc includes a disc bore including, a groove formed on an axially extending surface of the disc bore, wherein the groove includes a forward surface that extends radially into the disc bore to a groove floor that is cut into the disc bore and extends axially to an aft surface that extends radially outward to at least the axially extending surface of the disc bore, and a scallop formed along at least one surface of the groove, wherein the scallop is configured to provide a flow path, and a disc web that extends radially outward from the disc bore, relative to an axis of rotation of the gas turbine engine.

Abstract: A rotor disc assembly includes a rotor disc and a minidisc. The rotor disc has a first extension member, a first finger, and a second finger. The first extension member axially extends from a disc body disposed about an axis. The first finger extends axially from the first extension member. The second finger is circumferentially spaced apart from the first finger. The second finger extends axially from the first extension member. Each of the first finger and the second finger has a first portion and a second portion that extends radially from a distal end of the first portion. The minidisc is operatively connected to the rotor disc. The minidisc has an interlocking finger that radially extends from a minidisc body and is disposed between the first finger and the second finger. The interlocking finger, the first portion, and second portion define a ring groove.

Abstract: A combustor panel may include an attachment feature having a central longitudinal axis and extending from a cold side of the combustor panel. The attachment feature may include a tip portion and a base portion and the attachment feature may be configured to extend through a combustor shell such that the tip portion is disposed outward of a diffuser-facing side of the combustor shell. The base portion may be configured to be disposed between the cold side of the combustor panel and the diffuser-facing side of the combustor shell. The attachment feature may define a core passage extending from an inlet opening defined on the tip portion and extending partially through the attachment feature to terminate within the attachment feature. The attachment feature may also define an offshoot passage extending from the core passage to an outlet opening defined on the base portion.

Abstract: Combustor panels for use in gas turbine engines having a panel body having a first side and a second side, and a plurality of cooling pins extending from the first side, the plurality of cooling pins arranged in a pin array pattern, wherein at least some of the cooling pins extend a first height from the first side of the panel body, each cooling pin having a pin width, and is separated from adjacent cooling pins of the pin array by a pin array separation distance. The pin array separation distance is equal to or greater than the pin width.

Abstract: A gas turbine engine includes a combustor. A turbine section is in fluid communication with the combustor. The turbine section includes a first vane stage aft of the combustor. A seal assembly is disposed between the combustor and the first vane stage. The seal assembly includes a plurality of openings communicating cooling airflow into a gap between an aft end of the combustor and the first vane stage. A combustor assembly and method are also disclosed.

Abstract: According to one embodiment, a heat shield panel for a combustor of a gas turbine engine is provided. The heat shield comprising: a panel body having a first surface configured to be oriented toward a combustion zone of a combustor, and a second surface opposite the first surface, the second surface being configured to be oriented toward a combustor liner of the combustor; a plurality of first pin fins projecting from the second surface of the panel body, wherein each of the plurality of first pin fins has a rounded top opposite the second surface; and one or more second pin fins projecting from the second surface of the panel body, wherein each of the one or more second pin fins has a flat top opposite the second surface.

Abstract: A gas turbine engine includes a combustor. A turbine section is in fluid communication with the combustor. The turbine section includes a first vane stage aft of the combustor. A seal assembly is disposed between the combustor and the first vane stage. The seal assembly includes a first plurality of openings and the first vane stage includes a second plurality of openings communicating cooling airflow into a gap between an aft end of the combustor and the first vane stage. A first vane stage assembly and a method are also disclosed.

Abstract: A sealing arrangement includes a turbine static structure with a seal ring having groove with a contact surface. Also included is a bearing compartment with a second contact surface, and a piston seal positioned between the turbine static structure and the bearing compartment. The piston seal includes sides configured to contact the contact surfaces and is positioned in the groove.