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 method of assembling a mid-turbine frame includes engaging at least two tabs in a radially inner end of at least one spoke in a groove in an inner frame case. The at least one spoke is secured to the inner frame case with a central bolt extending through the inner frame case into the at least one spoke.

Abstract: A heat shield panel for use in 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, and an orifice extending through the rail member, from the inner wall to the outer wall, the orifice having a cross sectional area that increases from the inner wall to the outer wall.

Abstract: A gas turbine engine includes a combustor disposed about an engine central axis and a circumferential row of vanes adjacent an exit region of the combustor. The combustor includes a combustor wall and a combustion chamber. The combustor wall has a lip at the exit region. Each vane includes a platform and an airfoil section extending from the platform. The forward edges of the platforms are adjacent the lip of the combustor wall. The lip defines a first annular slot and the forward edges of the vanes collectively define a second annular slot. The first and second annular slots together define an annular seal slot. An annular feather seal is entrapped in the annular seal slot between the combustor wall and the platform.

Abstract: According to one embodiment, a mini-disk of a gas turbine engine having an axis is provided. The mini-disk includes a bore, a web extending radially with respect to the axis of the gas turbine engine from the bore, a base extending axially with respect to the axis of the gas turbine engine from the bore, a connector located on an end of the base, the connector configured to connect with a hub arm of the gas turbine engine, and an expansion feature configured in the base and located between the bore and the connector, the expansion feature configured to reduce an axial stiffness of the base.

Abstract: A vane for use in a gas turbine engine has an airfoil extending between a leading edge and a trailing edge, and is generally hollow. The vane has a radially outer platform and a radially inner platform. At least one of the radially inner and radially outer platforms has thicker portions circumferentially located beyond one of the leading edge and the trailing edge, and thinner portions circumferentially beyond circumferential edges of the airfoil. A mid-turbine frame and a gas turbine engine are also disclosed.

Abstract: A turbine vane has an airfoil extending between an inner platform and an outer platform. The airfoil is hollow. A hollow within the airfoil extends between an inner leading edge to an inner trailing edge. An inner radius is defined at the inner trailing edge in a circumferential direction measured between walls which extend from the inner leading edge to the inner trailing edge. The inner radius varies along a radial length defined between the inner and outer platforms. A mid-turbine frame and a gas turbine engine 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 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: A combustor for a gas turbine engine includes an annular combustor shell, the annular combustor shell defining a combustion chamber, and a fuel injector extending at least partially into the combustion chamber, and configured to deliver a flow of fuel and a flow of combustion air into the combustion chamber for combustion. The fuel injector includes a swirler with a swirler exit having a circumferential width along a circumferential axis greater than a radial width along a radial axis. A swirler for a gas turbine engine includes a swirler entrance and a swirler exit. The swirler exit has a circumferential width along a curvilinear circumferential axis greater than a radial width along a radial axis.

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: Methods of manufacturing toothed components for gas turbine engines are provided. The methods include forming a first tooth in the component with a top land, a bottom land, a side wall extending therebetween, and a fillet radius transitioning between the side wall and the bottom land, forming a second tooth in the component adjacent the first tooth, the second tooth having a top land, a bottom land, a side wall extending therebetween and facing the first tooth, and a fillet radius transitioning between the side wall and the bottom land, the bottom land of the second tooth extending toward the bottom land of the first tooth, wherein the bottom lands define a gable area of the component, and forming a stress relief feature in the gable area such that the stress relief feature reduces a stress concentration near the gable area during operation of the toothed component.

Abstract: A mid-turbine frame for a gas turbine engine according to an example of the present disclosure includes, among other things, a first frame case, a flange coupled to the first frame case, and a heat shield adjacent to the flange. A method of cooling a portion of a gas turbine engine is also disclosed.

Abstract: A compliant seal assembly that may be for an MTF of a gas turbine engine includes a housing orientated about an axis with a circumferentially extending groove in the housing opened radially outward. An annular carrier of the assembly is constructed and arranged to move radially within the groove, and includes a circumferentially extending channel that is opened radially outward for receipt of a piston ring that may seal to a surrounding cylindrical wall. In operation, the carrier moves radially with respect to the housing to compensate for radial displacement between the housing and the wall, and the piston ring moves axially along the wall to compensate for axial displacement between the housing and the wall.

Abstract: A method of controlling fuel injection into a combustor of a gas turbine engine including: applying a first electrical charge to fuel such that the fuel becomes a charged fuel; and applying a second electrical charge to a component of the combustor, wherein the first electrical charge is applied to the fuel at a first frequency and the second electrical charge is applied to the component at a second frequency such that at least one of a selected tone, a selected screech, and a selected noise is produced by spraying the charged fuel through the component and into a combustion chamber of the combustor from a fuel nozzle.

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, and a rail disposed about a periphery of the cold side, the rail including a first rail member having a first length extending along the panel, a first height extending from the cold side, and a first thickness that varies along the first length.

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: According to one embodiment, a mini-disk of a gas turbine engine having an axis is provided. The mini-disk includes a bore, a web extending radially with respect to the axis of the gas turbine engine from the bore, a base extending axially with respect to the axis of the gas turbine engine from the bore, a connector located on an end of the base, the connector configured to connect with a hub arm of the gas turbine engine, and an expansion feature configured in the base and located between the bore and the connector, the expansion feature configured to reduce an axial stiffness of the base.

Abstract: According to one or more embodiments, a cover plate for a gas turbine engine is provided. The cover plate includes a cover plate body that extends from a disc blade attachment radially along an outside surface of a disc web to a disc bore of a disc creating a cavity between the cover plate body and the disc web of the disc, and an attachment protrusion extending from the inner most end of the cover plate body toward the disc bore, and wherein the attachment protrusion is configured to engage with a retention ring.

Abstract: A method of controlling fuel injection into a combustor of a gas turbine engine including: applying a first electrical charge to fuel such that the fuel becomes a charged fuel; and applying a second electrical charge to a component of the combustor, wherein the first electrical charge is applied to the fuel at a first frequency and the second electrical charge is applied to the component at a second frequency such that at least one of a selected tone, a selected screech, and a selected noise is produced by spraying the charged fuel through the component and into a combustion chamber of the combustor from a fuel nozzle.