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 combustion chamber for a turbine engine, in particular for an aircraft turbojet engine or turboprop engine. The combustion chamber includes a radially outer annular shroud, a radially inner annular shroud coaxial with the radially outer shroud, and an end wall connecting the radially outer shroud and the radially inner shroud. The combustion chamber further includes a first annular sealing member coaxial with said radially inner and outer shrouds. The first annular sealing member is radially interposed between the end wall and the radially outer shroud.

A combustor includes a liner having an outlet end to pass combustion gas and a liner flange protruding outward from the outlet end; a transition piece to discharge combustion gas from the liner to a turbine, the transition piece having an inlet end for coupling to the outlet end of the liner and a transition piece flange protruding outward from the inlet end to face the liner flange; and a first elastic support installed on the liner flange to protrude toward the transition piece flange. A force applied from the transition piece elastically deforms an elastic arch of the first elastic support, which includes a movable support that is spaced apart from the liner flange if the force applied from the transition piece does not primarily deform the elastic arch. An auxiliary elastic support installed inside the first elastic support elastically deforms if the force secondarily deforms the elastic arch.

In accordance with at least one aspect of this disclosure, a fuel seal for an aircraft fuel system includes a seal ring of a metallic material configured to retain a shape of the seal ring under compression, and a coating of elastomeric material, over molded onto the seal ring and configured to conform to a shape of a sealing surface. In embodiments, the seal ring and coating can be configured to provide drop-tight sealing at a temperature of at least 400 degrees F. In certain embodiments, the seal ring and coating can be configured to seal one or more fuel system components of an aircraft fuel system.

The present disclosure provides devices for guiding an element in a combustion chamber wall opening. Between the welds or brazes joining a bushing with a cup and the trajectory of the ring edge, an intermediate space is reserved in an internal annular groove for said welds or brazes to protrude into, so that said welds or brazes lie outside the trajectory of the ring edge.

A combustion chamber of a gas turbine having a peripheral support structure and a heat shield arranged therein. The support structure has a reduced cross-section on the downstream side and a stop element on the upstream side. There is a gap between the heat shield and the stop element to compensate for different strains and tolerances. For reduction of the cooling air consumption, a peripheral sealing groove which is open towards the heat shield and has a sealing element arranged therein is provided in the stop element, which sealing element rests against the heat shield and covers the gap.

A gas turbine engine including a core having a compressor section fluidly connected to a combustor via a primary flowpath and a turbine section connected to the combustor via the core flow path. An assembly is disposed within the gas turbine engine and includes a first part connected to a second part via a radial stack joint. The first part includes a radially inward facing surface contacting a corresponding radially outward facing surface of the second part. A fastener protrudes through the first part and the second part and is configured to maintain the relative positions of the first part and the second part. A channel is disposed on at least one of the radially inward facing surfaces and is positioned between the fastener and a circumferential edge of the first part. The channel is connected to at least one cooling air source.

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 combustor for a gas turbine engine includes a combustor shell, a burner seal, and a burner seal retainer. The combustor shell includes metallic materials and is formed to define an interior combustion space. The burner seal includes ceramic matrix composite materials and is configured to extend through apertures formed in the combustor shell and the heat shield. The burner seal retainer is configured to retain the burner seal to the combustor shell.

A combustor adapted for use in a gas turbine engine includes a combustor shell comprising metallic materials. The combustor shell is formed to define an internal space. The combustor further includes a heat shield mounted to an axially aft surface of the combustor shell within the internal space and a combustor liner arranged to extend along inner surfaces of the combustor shell within the internal space. The combustor liner cooperates with the heat shield to define a combustor chamber.