The jet engine comprising a ramjet air path extending from an intake, into a combustion chamber, and out an exhaust nozzle, a fuel inlet leading into the combustion chamber, an oxidizer inlet leading into the combustion chamber and a partition being operable to selectively close the ramjet air path upstream of the combustion chamber to allow operation of the jet engine in rocket mode and open the ramjet air path to allow operation of the jet engine in ramjet mode.

A diffuser case for a gas turbine engine including: a pre-diffuser; a diffuser case defining a dump region, an inner shroud, and an outer shroud, the pre-diffuser being fluidly connected to the inner shroud and the outer shroud through the dump region; a tangential onboard injector module fluidly connected to the inner shroud through inlet orifice located in the diffuser case proximate an aft end of the inner shroud; and an inlet extender initiating at the inlet orifice of the tangential onboard injector module, extends through the inner shroud, and terminates at a distal end proximate the pre-diffuser.

A tubular combustion chamber system for a gas turbine unit includes a plurality of annularly arranged transition lines, which are designed to be connected at the upstream ends thereof to respective burners and to conduct hot gas produced by the burners to a turbine. The tubular combustion chamber system has a hot gas distributor, which is designed to be connected to the turbine and defines a ring channel, which is open to the turbine and into which the downstream ends of the transition lines lead. A gas turbine unit includes a plurality of annularly arranged burners, a turbine and a tubular combustion chamber system which connects the burners to the turbine.

A tubular combustion chamber system for a gas turbine unit includes a plurality of annularly arranged transition lines, which are designed to be connected at the upstream ends thereof to respective burners and to conduct hot gas produced by the burners to a turbine. The tubular combustion chamber system has a hot gas distributor, which is designed to be connected to the turbine and defines a ring channel, which is open to the turbine and into which the downstream ends of the transition lines lead. A gas turbine unit includes a plurality of annularly arranged burners, a turbine and a tubular combustion chamber system which connects the burners to the turbine.

Heatshields for installation within gas turbine engines are described. The heatshields include a metal body having a first end, a second end, a first side, and a second side, wherein the first side and the second side define parallel sides extending from the first end to the second end, an engagement portion formed along the first side and arranged to engage with a portion of a case, a shielding portion formed along the second side, and a mid-body portion extending between the engagement portion and the shielding portion and has an arcuate shape in cross-section. The metal body is configured to form a hoop, split-ring structure with the first end attached to the second end.

Heatshields for installation within gas turbine engines are described. The heatshields include a metal body having a first end, a second end, a first side, and a second side, wherein the first side and the second side define parallel sides extending from the first end to the second end, an engagement portion formed along the first side and arranged to engage with a portion of a case, a shielding portion formed along the second side, and a mid-body portion extending between the engagement portion and the shielding portion and has an arcuate shape in cross-section. The metal body is configured to form a hoop, split-ring structure with the first end attached to the second end.

A fuel injection system includes an outer support defining a fuel manifold and an inner support, with a feed arm extending radially between the inner support and the outer support. A plurality of outlet openings extending in an axial direction from the feed arm for feeding respective injection nozzles. The feed arm defines a plurality of fuel passages therethrough in fluid communication with the fuel manifold and outlet openings to supply fuel from the fuel manifold to the outlet openings. A heat shield extends from the outer support to the inner support and extends about the outer support and the feed arm to provide heat shielding to the fuel manifold and the fuel passages.

A fuel injection system includes an outer support defining a fuel manifold and an inner support, with a feed arm extending radially between the inner support and the outer support. A plurality of outlet openings extending in an axial direction from the feed arm for feeding respective injection nozzles. The feed arm defines a plurality of fuel passages therethrough in fluid communication with the fuel manifold and outlet openings to supply fuel from the fuel manifold to the outlet openings. A heat shield extends from the outer support to the inner support and extends about the outer support and the feed arm to provide heat shielding to the fuel manifold and the fuel passages.

A combustor section of a gas turbine engine includes a combustor having a combustor inlet, and a combustor bypass passage having a passage inlet located upstream of the combustor inlet. The combustor bypass passage is configured to divert a selected bypass airflow around the combustor. A combustor bypass valve is located at the combustor bypass passage to control the selected bypass airflow along the combustor bypass passage. A method of operating a gas turbine engine, includes urging a core airflow from a compressor section toward a combustor section, flowing a first portion of the core airflow into the combustor section via a combustor inlet, and flowing a second portion of the core airflow into a combustor bypass passage via a combustor bypass valve, thereby bypassing the combustor with the second portion of the core airflow.

A combustor section of a gas turbine engine includes a combustor having a combustor inlet, and a combustor bypass passage having a passage inlet located upstream of the combustor inlet. The combustor bypass passage is configured to divert a selected bypass airflow around the combustor. A combustor bypass valve is located at the combustor bypass passage to control the selected bypass airflow along the combustor bypass passage. A method of operating a gas turbine engine, includes urging a core airflow from a compressor section toward a combustor section, flowing a first portion of the core airflow into the combustor section via a combustor inlet, and flowing a second portion of the core airflow into a combustor bypass passage via a combustor bypass valve, thereby bypassing the combustor with the second portion of the core airflow.