Abstract: A bypass duct of a turbofan gas turbine engine includes a first shroud extending at least partially around an axis and a second shroud extending at least partially around the axis, the second shroud being radially spaced apart from the first shroud to define a bypass passage between the first and second shrouds. The bypass duct further includes a fairing disposed in the bypass passage. The fairing includes a first outer shell and a second outer shell defining therebetween an internal passage. The first and second outer shells extend radially from the first shroud to the second shroud. A plate extends across the internal passage between the first and second outer shells. A damper is engaged with the plate between the first and second outer shells.

Abstract: An apparatus for stabilizing a fan assembly of a gas turbine engine during transport includes a ring configured to engage a nose cone of the fan assembly; an outer bracket connected to the ring and configured to mount to a fan casing of the fan assembly; and a tensioning assembly between the ring and the outer bracket to move the ring along an axis of a shaft of the fan assembly. An assembly with the gas turbine engine and a method are also disclosed.

Abstract: An aircraft engine has a flammable fluid containment assembly including a component containing a flammable fluid. A transfer tube is fluidly connected to the component. A sealing ring is provided at an interface between the aircraft engine component and the transfer tube. A first thermal barrier is provided between the sealing ring and a first one of the aircraft engine component and the flammable fluid inside the transfer tube. The first thermal barrier includes a first annular cavity filled with a first thermal insulation medium and axially spanning the sealing ring.

Abstract: An aircraft engine has a flammable fluid containment assembly including a component containing a flammable fluid. A transfer tube is fluidly connected to the component. A sealing ring is provided at an interface between the aircraft engine component and the transfer tube. A first thermal barrier is provided between the sealing ring and a first one of the aircraft engine component and the flammable fluid inside the transfer tube. The first thermal barrier includes a first annular cavity filled with a first thermal insulation medium and axially spanning the sealing ring.

Abstract: A system for an aircraft engine, has: an aircraft component; and a heat exchanger having: a housing defining a first inlet, a first outlet, a second inlet, and a second outlet; first conduits within the housing, the first conduits fluidly connecting the first inlet to the first outlet; one or more second conduit within the housing, the one or more second conduits fluidly connecting the second inlet to the second outlet, the one or more second conduit in heat exchange relationship with the first conduits; a mixing chamber intersecting two or more of the first conduits, the mixing chamber having a peripheral wall extending around a mixing volume and a central axis; and protrusions extending from the peripheral wall into the mixing volume, the protrusions extending around the central axis and across a direction of the flow through the mixing chamber.

Abstract: A system for an aircraft engine, has: an aircraft component; and a heat exchanger having: a housing defining a first inlet, a first outlet, a second inlet, and a second outlet; first conduits within the housing, the first conduits fluidly connecting the first inlet to the first outlet; one or more second conduit within the housing, the one or more second conduits fluidly connecting the second inlet to the second outlet, the one or more second conduit in heat exchange relationship with the first conduits; and a mixing chamber intersecting two or more of the first conduits and separating the first conduits into upstream sections and downstream sections, the mixing chamber having a peripheral wall extending around a mixing volume, the peripheral wall defining a convergent section in which a flow circulating area of the mixing volume decreases in a downstream direction relative to the flow through the first conduits.

Abstract: A system for an aircraft engine, has: an aircraft component; and a heat exchanger having: a housing defining a first inlet, a first outlet, a second inlet, and a second outlet; first conduits within the housing, the first conduits fluidly connecting the first inlet to the first outlet; one or more second conduit within the housing, the one or more second conduits fluidly connecting the second inlet to the second outlet, the one or more second conduit in heat exchange relationship with the first conduits; a mixing chamber intersecting two or more of the first conduits, the mixing chamber having a peripheral wall extending around a mixing volume and a central axis; and fins protruding from the peripheral wall into the mixing volume, the fins extending in a direction having an axial component relative to the central axis, the fins defining flow passages interspaced between the fins.

Abstract: A system for an aircraft engine, has: an aircraft component; and a heat exchanger having: a housing defining a first inlet, a first outlet, a second inlet, and a second outlet; first conduits within the housing, the first conduits fluidly connecting the first inlet to the first outlet; one or more second conduit within the housing, the one or more second conduits fluidly connecting the second inlet to the second outlet, the one or more second conduit in heat exchange relationship with the first conduits; a mixing chamber intersecting two or more of the first conduits, the mixing chamber having a peripheral wall extending around a mixing volume and a central axis; and louvers mounted to the peripheral wall and extending transversally to the central axis, the louvers movable between a collapsed configuration and a deployed configuration and extending across the mixing volume in the deployed configuration.

Abstract: A system for an aircraft engine, has: an aircraft component; and a heat exchanger having: a housing defining a first inlet, a first outlet, a second inlet, and a second outlet; first conduits within the housing, the first conduits fluidly connecting the first inlet to the first outlet; one or more second conduit within the housing, the one or more second conduits fluidly connecting the second inlet to the second outlet, the one or more second conduit in heat exchange relationship with the first conduits; a mixing chamber intersecting two or more of the first conduits, the mixing chamber having a peripheral wall extending around a mixing volume and a central axis; and protrusions extending from the peripheral wall into the mixing volume, the protrusions extending around the central axis and across a direction of the flow through the mixing chamber.

Abstract: An aircraft engine has a flammable fluid containment assembly including a component containing a flammable fluid. A transfer tube is fluidly connected to the component. A sealing ring is provided at an interface between the aircraft engine component and the transfer tube. A first thermal barrier is provided between the sealing ring and a first one of the aircraft engine component and the flammable fluid inside the transfer tube. The first thermal barrier includes a first annular cavity filled with a first thermal insulation medium and axially spanning the sealing ring.

Abstract: The heat exchanger can have a first fluid circuit extending from a first inlet to a first outlet, one or both having a first mouth dividing internally into a number of passages within a periphery, a second fluid circuit fluidly distinct from the first fluid circuit, extending from a second inlet to a second outlet, the second fluid circuit in thermal exchange contact with the number of passages in a heat transfer region of the heat exchanger, and an adapter having a second mouth secured to the first mouth in a manner to establish internal fluid flow communication therewith.

Abstract: Recent literature on SARS-CoV-2 pathogenesis has suggested that the induction of substantial acute respiratory distress phenotypes is driven by a mismatched inflammatory response together with broad vascular dysfunction. While several detailed reports implementing multi-omic approaches have provided insight into the immune cell phenotypes involved in these processes, risk stratifying markers specific to COVID-19 and the vasculature have not been explored. Provided herein is a comprehensive, multi-omics-based description of the molecular antecedents to COVID-19 mortality, yielding new insights pertaining to the vasculature while highlighting the urgent need for clinical translation of novel biomarkers for disease prognosis.

Abstract: A gas turbine engine includes a fan assembly, a compressor assembly, a combustion chamber, a turbine assembly, a bypass duct conveying rearward a bypass airstream driven by the fan assembly when the gas turbine engine is in use, a fairing extending across at least a portion of the bypass duct downstream of the fan assembly, and a heat exchanger having an inlet fluidly connected to the compressor assembly and an outlet fluidly connected to a pneumatic actuator of the gas turbine engine. The fairing has a leading edge and a trailing edge. The heat exchanger is disposed adjacent the trailing edge of the fairing.

Abstract: The heat exchanger can have a first fluid circuit extending from a first inlet to a first outlet, one or both having a first mouth dividing internally into a number of passages within a periphery, a second fluid circuit fluidly distinct from the first fluid circuit, extending from a second inlet to a second outlet, the second fluid circuit in thermal exchange contact with the number of passages in a heat transfer region of the heat exchanger, and an adapter having a second mouth secured to the first mouth in a manner to establish internal fluid flow communication therewith.

Abstract: A turbofan engine comprising an outer bypass duct, an annular bypass flow path between the outer bypass duct and a core engine, an engine component forming an airflow obstruction adjacent the outer bypass duct, an air cooler having a tube, the tube having at least a sinuous portion extending in the annular bypass flow path, the sinuous portion extending along the outer bypass duct, downstream of the airflow obstruction, the sinuous portion configured for exchanging heat between a fluid circulating in the tube and air circulating in the bypass flow path during operation of the turbofan engine.

Abstract: A turbofan engine comprising an outer bypass duct, an annular bypass flow path between the outer bypass duct and a core engine, an engine component forming an airflow obstruction adjacent the outer bypass duct, an air cooler having a tube, the tube having at least a sinuous portion extending in the annular bypass flow path, the sinuous portion extending along the outer bypass duct, downstream of the airflow obstruction, the sinuous portion configured for exchanging heat between a fluid circulating in the tube and air circulating in the bypass flow path during operation of the turbofan engine.

Abstract: A gas turbine engine includes a fan assembly, a compressor assembly, a combustion chamber, a turbine assembly, a bypass duct conveying rearward a bypass airstream driven by the fan assembly when the gas turbine engine is in use, a fairing extending across at least a portion of the bypass duct downstream of the fan assembly, and a heat exchanger having an inlet fluidly connected to the compressor assembly and an outlet fluidly connected to a pneumatic actuator of the gas turbine engine. The fairing has a leading edge and a trailing edge. The heat exchanger is disposed adjacent the trailing edge of the fairing.

Abstract: A mounting method and system for supporting and positioning an internal fuel manifold relative to a combustor in a fuel injection system for a gas turbine engine. The method includes positioning the fuel manifold within a gas generator casing, applying a heat shield around a radially extending fuel inlet connected to the fuel manifold, and using the heat shield to at least partially support the fuel manifold within the gas generator casing.

Abstract: A fuel manifold assembly for a gas turbine engine fuel system comprises a fuel manifold at least partly enclosed by a heat shield, the internal manifold being made of a first material having a first coefficient of thermal expansion, and the heat shield being made of a second material having a second coefficient of thermal expansion that is lower than the first coefficient of thermal expansion.

Abstract: An access conduit for access to the interior of a combustor for the washing of the turbine components of a gas turbine engine. The access conduit is secured between a gas generator case and a combustion chamber liner of the combustor. The access conduit has an open outer end accessible from an outer face of the gas generator case and an open inner end exiting in a combustion chamber through the liner. The access conduit has one or more openings disposed in a combustor gap between the gas generator case and the liner of the combustor. A plug is removably secured in the open outer end and extends into the access conduit to obstruct the one or more openings.