A turbofan gas turbine engine configured to reduce hotspots within combustors. The engine includes an axis and a combustor that is circumferentially disposed about the axis. The combustor includes an annular combustor liner that includes a front portion and a rear portion. The annular combustor liner is joined to an annular combustor dome via front portion and defines a chamber and a nozzle is mounted within the annular combustor dome and is configured to inject fuel into a plurality of swirlers. At least one or more dilution openings is circumferentially distributed around the liner such that a region is fluidly connected through the annular combustor liner to the chamber. Each one of the pluralities of dilution openings includes an opening and a radial support wall that is positioned aft of the opening such that the radial support wall extends into the chamber.

A gas turbine combustion duct includes a duct body and a fluid intake hole. The duct body includes a duct wall defining a plenum for routing a flow of combustion products from a combustor downstream through the gas turbine combustion duct to a turbine section. The fluid intake hole extends from an outward-facing surface to an inward-facing surface through the duct wall for receiving an outside fluid flow into the plenum, and is laterally circumscribed about its entire periphery by a lateral-facing surface. The lateral-facing surface includes a curved surface portion along a shortest path from the inward-facing surface to the outward-facing surface and is free of angled surface transitions along the shortest path between the inward-facing surface and the outward-facing surface. The fluid intake hole is wider at the outward-facing surface than at the inward-facing surface. A boss may define the fluid intake hole and the lateral-facing surface.

A gas turbine engine that includes a combustor configured for efficient combustion of fuel for the generation of combustion gases. The engine includes an annular combustor that includes an annular wall that defines a channel configured to conduct hot combustion gases along a flow-path of combustion gases. The annular wall has an exterior first surface and an interior second surface. A plurality of dilution holes is defined through the wall. Each dilution hole is defined by a dilution hole surface that extends between an entry end and an exit end. The entry end is defined by the first surface and the exit end is defined by the second surface such that the dilution hole has a convergent and then divergent cross-sectional profile. The entry end has a first geometric shape and the exit end has a second geometric shape that is different than the second geometric shape.

A method of reducing noise from a combustor of a gas turbine engine includes the steps of establishing a maximum noise limit that may be for a particular frequency range. A primary fuel flow percentage, which may be emitted from a fuel nozzle arrangement having various groupings of simplex and duplex nozzles, is then established. An immersion depth measured between an aft rim of a swirler and a distal tip of the fuel nozzles may then be reduced thereby reducing the noise amplitude.

A liner cooling device for cooling a liner of a gas turbine is provided. The liner cooling device may include a support portion disposed between a liner and a transition piece of a gas turbine and configured to include a cooling flow passage through which cooling air moves to the transition piece. The support portion includes a support member disposed between the liner and the transition piece and an auxiliary support member disposed in the cooling flow passage and having a hole through which the cooling air passes.

A manufacturing method is provided. During this method, a preform component is provided for a turbine engine. The preform component includes a substrate. A meter section of a cooling aperture is formed in the substrate. An internal coating is applied onto a surface of the meter section. An external coating is applied over the substrate. A diffuser section of the cooling aperture is formed in the external coating and the substrate to provide the cooling aperture.

Methods for manufacturing combustor panels of gas turbine engines and combustor panels are described. The methods include defining a particle deposit near-steady state for at least a portion of a combustor panel, the particle deposit near-steady state representative of a build-up of particles on the at least a portion of the combustor panel during use, generating a template based on the defined particle deposit near-steady state, wherein the template includes one or more augmentation elements based on the representative of build-up of particles, and forming a combustor panel based on the template, wherein the formed combustor panel includes one or more augmentation elements defined in the template.

A plurality of cooling passages extending in an axial direction are formed in a transition piece so as to be aligned in a circumferential direction and the axial direction. One or more downstream side passages are formed in a downstream side region (Rd) within one circumferential region. One or more upstream side passages are formed in an upstream side region Ru within the circumferential region. The total cross-sectional area per unit circumferential length of the one or more downstream side passages is larger than the total cross-sectional area per unit circumferential length of the one or more upstream side passages.

A self-cooled orifice structure that may be for a combustor of a gas turbine engine, and may further be a dilution hole structure, includes a hot side panel, a cold side panel spaced from the hot side panel, and a continuous inner wall extending between the hot and cold side panels and defining an orifice having a centerline and communicating axially through the hot and cold side panels. A plurality of end walls of the structure are in a cooling cavity that is defined in-part by the hot and cold side panels and the inner wall. Each end wall extends between and are engaged to the hot and cold side panels and are circumferentially spaced from the next adjacent end wall. A plurality of inlet apertures extend through the cold side panel and are in fluid communication with the cavity, and each one of the plurality of inlet apertures are proximate to a first side of a respective one of the plurality of end walls. A plurality of outlet apertures extend through the hot side panel and are in fluid communication with the cavity, and each one of the plurality of outlet apertures are associated with an opposite second side of a respective one of the plurality of end walls.

The described reverse flow combustor of a gas turbine engine includes inner and outer combustor liners defining a combustor chamber therewithin. A large exit duct and a small exit duct are disposed at downstream ends of the outer and inner liner respectively. The small exit duct includes an annular ring removably mounted to a support element of the gas turbine engine and includes a plurality of cooling elements integrally formed with the annular ring and projecting therefrom into impingement airflow. The cooling elements increase the effective surface area of the inner surface of the annular ring, which is adapted to be cooled by the impingement airflow.