A combustor assembly includes a volute wall extended annularly around a combustor centerline, extended at least partially as a spiral curve from a circumferential reference line around the combustor centerline, and defining a combustion chamber therewithin, an annular inner wall extended at least partially along a lengthwise direction from the volute wall, an annular outer wall extended at least partially along the lengthwise direction from the volute wall, the annular inner wall and the annular outer wall being separated along a radial direction from the combustor centerline, and a primary flow passage being defined between the annular inner wall and the annular outer wall in fluid communication with the combustion chamber, and a flow passage wall extending between a portion of the volute wall and a portion of the annular outer wall.

A section for a gas turbine engine according to an example of the present disclosure includes, among other things, a rotor including a row of blades extending in a radial direction outwardly from a hub. The row of blades deliver flow to a bypass flow path, an intermediate flow path, and a core flow path. A first case surrounds the row of blades to establish the bypass flow path. A first flow splitter divides flow between the bypass flow path and a second duct. An aftmost row of guide vanes extends in the radial direction across the bypass flow path. A second flow splitter radially inboard of the first flow splitter divides flow from the second duct between the intermediate flow path and the core flow path. A bypass port interconnects the intermediate and bypass flow paths.

A section for a gas turbine engine according to an example of the present disclosure includes, among other things, a rotor including a row of blades extending in a radial direction outwardly from a hub. The row of blades deliver flow to a bypass flow path, an intermediate flow path, and a core flow path. A first case surrounds the row of blades to establish the bypass flow path. A first flow splitter divides flow between the bypass flow path and a second duct. An aftmost row of guide vanes extends in the radial direction across the bypass flow path. A second flow splitter radially inboard of the first flow splitter divides flow from the second duct between the intermediate flow path and the core flow path. A bypass port interconnects the intermediate and bypass flow paths.

A method of controlling plume exhaust heat and/or noise radiation from a turbofan engine assembly having a short nacelle. A mixer duct shell is supported such that a downstream edge of the short nacelle overlays an upstream portion of the mixer duct shell. A first portion of fan exhaust may be routed through the mixer duct shell between its inner surface and an outer surface of a core engine shroud. A second portion of fan exhaust may be routed over an outer surface of the mixer duct shell. At least one of the inner surface and an outer surface of the mixer duct shell may have an acoustic lining including a honeycomb core structure.

A method of controlling plume exhaust heat and/or noise radiation from a turbofan engine assembly having a short nacelle. A mixer duct shell is supported such that a downstream edge of the short nacelle overlays an upstream portion of the mixer duct shell. A first portion of fan exhaust may be routed through the mixer duct shell between its inner surface and an outer surface of a core engine shroud. A second portion of fan exhaust may be routed over an outer surface of the mixer duct shell. At least one of the inner surface and an outer surface of the mixer duct shell may have an acoustic lining including a honeycomb core structure.

A heat management system for a turbomachine may include a first heat exchanger configured and arranged to receive a first fluid stream from a first duct, a second heat exchanger configured and arranged to receive the first fluid stream after discharging from the first heat exchanger, and a hatch configured to provide fluid communication from a second duct to the first duct so as to introduce a second fluid stream from the second duct to the first duct. A method of cooling fluid streams may include directing a first fluid stream from a first duct across or through a first heat exchanger, directing the first fluid stream across or through a second heat exchanger after discharging from the first heat exchanger, and directing a second fluid stream from a second duct to the first duct, with the second fluid stream flowing through a hatch configured to provide fluid communication from the second duct to the first duct.

A supersonic gas turbine engine for an aircraft that comprises a nacelle, a fan, an engine core including a primary duct configured to guide a core airflow through the engine core, a bypass duct extending between the engine core and an engine casing and configured to guide a bypass airflow through the bypass duct, an intake located upstream of the fan, and a tertiary airflow duct extending between the engine casing and the nacelle and configured to guide a tertiary airflow. The intake is configured to extract air from the intake and guide it to the tertiary airflow duct in which the extracted air flows as tertiary airflow. It is provided that at least one heat exchanger is mounted in the tertiary airflow duct.

Systems and methods are provided for an air-driven augmentor fan equipped aircraft propulsor. The augmentor fan may increase the effective bypass ratio of the aircraft propulsor and reduce fuel consumption and carbon emissions of the aircraft. The augmentor fan may be driven by air energized by a ducted fan powered by the core engine of the aircraft propulsor. The energized air may be received by an inlet, flowed through a flow path, and exhausted out the outlet to drive the augmentor fan. The exhausted energized air may impart a torque on the augmentor fan or blades of the augmentor fan. One or more of the inlet, flow path, or outlet may be variable in size to control the volume of air flowed through the flow path.

Systems and methods are provided for an air-driven augmentor fan equipped aircraft propulsor. The augmentor fan may increase the effective bypass ratio of the aircraft propulsor and reduce fuel consumption and carbon emissions of the aircraft. The augmentor fan may be driven by air energized by a ducted fan powered by the core engine of the aircraft propulsor. The energized air may be received by an inlet, flowed through a flow path, and exhausted out the outlet to drive the augmentor fan. The exhausted energized air may impart a torque on the augmentor fan or blades of the augmentor fan. One or more of the inlet, flow path, or outlet may be variable in size to control the volume of air flowed through the flow path.

A gas turbine engine defining a centerline and a circumferential direction, the gas turbine engine including: a turbomachine comprising a compressor section, a combustion section, and a turbine section arranged in serial flow order, the turbomachine defining a working gas flowpath and a fan duct flowpath; a primary fan driven by the turbomachine defining a primary fan tip radius R.sub.1 and a primary fan hub radius R.sub.2; a secondary fan located downstream of the primary fan and driven by the turbomachine, at least a portion of an airflow from the primary fan configured to bypass the secondary fan, the secondary fan defining a secondary fan tip radius R.sub.3 and a secondary fan hub radius R.sub.4, wherein the secondary fan is configured to provide a fan duct airflow through the fan duct flowpath during operation to generate a fan duct thrust, wherein the fan duct thrust is equal to % Fn.sub.3S of a total engine thrust during operation of the gas turbine engine at a rated speed during standard day operating conditions; wherein a ratio of R.sub.1 to R.sub.3 equals $\sqrt{\left(\mathrm{EFP}\right)\frac{\left(1-Rq{R}_{\mathrm{Sec}.-{\mathrm{Fan}}^2}\right)}{\left(1-Rq{R}_{\mathrm{Prim}.-\mathrm{Fa}{n}^2}\right)}\left(\frac{1}{\%{\mathrm{Fn}}_{3S}}-1\right)};$