Void structures, systems and devices with void structures, and methods of fabricating void structures are disclosed. A void structure is disclosed with a repeating elongated-aperture pattern designed to provide negative Poisson's Ratio behavior under macroscopic stress and strain loading. The pattern can include horizontal and vertical elliptically shaped apertures that are arranged on horizontal and vertical lines in a way that the lines are equally spaced in both dimensions. The centers of each aperture is on a crossing point of two of the lines. The vertical and horizontal elliptically shaped apertures alternate on the vertical and horizontal lines such that any vertical aperture is surrounded by horizontal apertures along the lines (and vice versa), and the next vertical apertures are found on both diagonals. The voids can also act as cooling and/or damping holes and, due to their arrangement, also as stress reduction features.

A deflector panel for a combustor of a gas turbine engine includes an outer radial edge, an inner radial edge, and two circumferential edges extending between the outer radial edge and the inner radial edge. Each circumferential edge including a panel scallop at least partially defining a deflector opening receptive of a fuel nozzle of the combustor. A combustor of a gas turbine engine includes a combustor shell at least partially defining a combustion zone therein, and a deflector assembly operably connected to the combustor shell. The deflector assembly includes a plurality of circumferentially abutted deflector panels. Each deflector panel includes an outer radial edge, an inner radial edge, and two circumferential edges extending between the outer radial edge and the inner radial edge. Each circumferential edge includes a panel scallop at least partially defining a deflector opening. A fuel nozzle extends through the deflector opening.

The present invention is a combustor provided with a combustor transition pipe connected to a turbine while interposing a transition pipe seal in between, including a flange portion provided at an end portion on a downstream side in a fluid flow direction of the combustor transition pipe, the flange portion projecting to radially inside and extending in a circumferential direction. The flange portion includes a pin hole into which a pin to position the transition pipe seal is inserted, a circumferential slit portion either extending within a range in a radial direction where the pin hole is formed or being located on radially outside of the pin hole and extending in the circumferential direction, and a hole portion on which part of the circumferential slit portion abuts.

A gas turbine engine combustion chamber includes upstream and downstream ring structures and a plurality of circumferentially arranged combustion chamber segments. Each segment extends the full length of the combustion chamber and each segment is secured to the upstream ring structure and is mounted on the downstream ring structure. A frame structure at the downstream end of each segment has multiple spaced radially extending holes. The downstream end of each segment has an axially upstream extending groove and the downstream ring structure has an annular axially upstream extending hook which locates in the groove of each segment. A portion of the downstream ring structure abuts the frame structure of each segment. The downstream ring structure has multiple holes through the portion abutting the frame structure and segment is removably secured to the downstream ring structure by multiple fasteners locating in the holes in the segments and the downstream ring structure.

A fuel injector assembly for a gas turbine engine includes a centerbody extended along a lengthwise direction, the centerbody defining a first fuel nozzle, and an annular shroud defining a second fuel nozzle directly surrounding the centerbody and extended along the lengthwise direction. A passage is defined through the annular shroud and extended generally along the lengthwise direction. The passage defines an exit opening disposed at a downstream end adjacent to the combustion chamber and in fluid communication therewith. The annular shroud defines a fuel inlet opening disposed at an upstream end of the passage. The annular shroud defines an air inlet opening in fluid communication with the passage. The air inlet opening is disposed between the fuel inlet opening and the exit opening.

Dampers and a fuel delivery systems including such dampers are disclosed. The damper includes a first component and a dampening component. The first component is coupled to a first tubular element and includes a first extended hollow section. The dampening component is coupled to a second tubular element. The dampening component includes a first end portion including a plurality of slits. The first end portion is disposed within the first extended hollow section to frictionally couple the first end portion to the first extended hollow section.

A turbine component assembly is disclosed, including a first component, a second component, and a circumferentially oriented flat spring. The first component is arranged to be disposed adjacent to a hot gas path, and includes a ceramic matrix composite composition. The second component is adjacent to the first component and arranged to be disposed distal from the hot gas path across the first component. The circumferentially oriented flat spring is disposed on and directly contacting the second component and directly contacting and supporting the first component as a compliant contact interface between the first component and the second component. The circumferentially oriented flat spring provides a radial spring compliance between the first component and the second component.

A method of preventing spallation for a geometrically segmented thermally insulating top coat on an article, the method comprising forming a surface feature in a surface of the article; forming a crack deflection feature in the surface proximate the surface feature; disposing the thermally insulating topcoat over the surface feature; and forming segmented portions that are separated by faults extending through the thermally insulating topcoat from the surface feature.

A combustor for use in a gas turbine engine including: a radially inward heat shield panel; and a radially outward heat shield panel located radially outward of the radially inward heat shield panel, the radially inward heat shield panel and the radially outward heat shield panel being in a facing spaced relationship defining a combustion chamber therebetween, wherein at least one of the radially inward heat shield panel and the radially outward heat shield panel has a non-axisymmetric shape.

A system includes a turbine combustor. The turbine combustor has a combustor liner disposed about a combustion chamber, a flow sleeve, and a radial passageway. The flow sleeve disposed at an offset about the combustor liner to define a passage, wherein the passage is configured to direct an exhaust gas flow toward a head end of the turbine combustor. The radial passageway extends between the flow sleeve and the combustor liner, and the radial passageway is configured to isolate an oxidant flow through the radial passageway from the exhaust gas flow through the passage for a first operating condition and a second operating condition of the turbine combustor. The offset between the combustor liner and the flow sleeve at the first operating condition is greater than the offset between the combustor liner and the flow sleeve at the second operating condition.