A fuel delivery system for a gas turbine engine comprises a cryogenic fuel tank, a first fuel line for connection to the cryogenic fuel tank, a fuel pump connected to receive fuel via the first fuel line, a plurality of fuel lines connecting the fuel pump to a combustor of the gas turbine engine, a controller configured to operate the fuel delivery system, a purge gas tank connected to the first fuel line and configured to store a purge gas for purging the plurality of fuel lines and a fuel gas tank connected to the first fuel line and configured to store a fuel gas for flushing purge gas from the plurality of fuel lines.

In accordance with at least one aspect of the present disclosure, there is provided a fuel system of an aircraft engine, comprising: a fuel conduit interface connecting a fuel conduit to a component of the fuel system; and a sweep line structure. The sweep line structure includes an inner surface facing toward, extending around, and defining a cavity around the fuel conduit interface, and an outer surface opposite the inner surface, the cavity being fluidly sealed relative to the outer surface.

A transfer tube assembly comprises a transfer tube slidably engaged in sealing engagement with a first component. The transfer tube has a shoulder engageable with a stopper for limiting relative axial movement between the transfer tube and the first component. The shoulder has an abutment surface facing a corresponding bore surface of the first bore of the first component. The abutment surface and the bore surface are configured to generate axially opposing reaction forces in response to the abutment surface and the bore surface contacting each other.

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 hydraulic snubber insert can have an elongated stem and at least one segment extending transversally from the stem, each segment having a size and shape mating a cross-sectional size and shape of the liquid carrying line, and at least one aperture, the insert being configured for the at least one segment to be pushable snugly into the liquid carrying line and pullable out from the liquid carrying line via the stem.

Methods and systems of operating a gas turbine engine in a low-power condition are provided. In one embodiment, the method includes supplying fuel to a combustor by supplying fuel to a first fuel manifold and a second fuel manifold of the gas turbine engine. The method also includes, while supplying fuel to the combustor by supplying fuel to the first fuel manifold: stopping supplying fuel to the second fuel manifold; and supplying pressurized air to the second fuel manifold to flush fuel in the second fuel manifold into the combustor and hinder coking in the second fuel manifold and associated fuel nozzles.

A fuel system for an aircraft engine comprises a fuel metering unit and a separate flow divider. The flow divider has an inlet port fluidly connected to the fuel metering unit via a fuel line. A primary and a secondary fuel manifold are fluidly connected to the flow divider. The fuel metering unit and the flow divider have a fuel supply mode in which fuel is allowed to flow in a first direction through the fuel line from the fuel metering unit to the flow divider to feed the primary and secondary fuel manifolds, and an ecology mode in which fuel is allowed to flow in a second direction through the same fuel line from the flow divider towards the fuel metering unit. A same fuel line is thus used as a fuel supply line and an ecology line.

A system includes a combustor. The combustor has a combustor wall with a combustor dome at an upstream end of the combustor wall, and an outlet at a downstream end of the combustor wall opposite the upstream end. The combustor wall includes an inner wall portion and an outer wall portion defining an interior of the combustor therebetween. Each of the inner wall portion and outer wall portion extends from the combustor dome to the downstream end of the combustor wall. The combustor wall includes an air cooling passage embedded inside at least one of the inner wall portion and the outer wall portion. The air cooling passage extends from the upstream end of the combustor wall to the downstream end of the combustor wall.

Turbine engine systems are described. The turbine engine systems include a combustor arranged along a core flow path of the turbine engine, a cryogenic fuel tank configured to supply a fuel to the combustor, a fuel supply line having a first flow supply line and a second flow supply line, the first flow supply line fluidly connecting the cryogenic fuel tank to the combustor through a first core flow path heat exchanger, and the second flow supply line fluidly connecting the cryogenic fuel tank to the combustor through a second core flow path heat exchanger, and a flow controller arranged along the fuel supply line and configured to respectively control a flow of fuel into the first flow supply line and the second flow supply line.

A flow diverter of a coke catching element is provided. The flow diverter includes a body disposable in a flow field and defining an interior, an upstream end, which is defined relative to the flow field, and a downstream end, which is downstream from the upstream end. At least one of the upstream end and the downstream end is open to the interior and the other of the upstream end and the downstream end is closed. The interior diverges at the one of the upstream end and the downstream end that is open and converges at the other of the upstream end and the downstream end.