An aircraft propulsion system includes a propulsive fan assembly configured for assembly into an aircraft structure, the propulsive fan assembly that includes a fan rotatable about a fan axis, an inlet duct assembly disposed within the aircraft fuselage, the inlet duct assembly that includes an upper inlet duct with an upper inlet opening and a lower inlet duct with a lower inlet opening. The upper inlet duct and the lower inlet duct merge into a common inlet duct forward of the propulsive fan assembly, and an outlet duct is disposed aft of the propulsive fan assembly.

Various embodiments of an airfoil and machines with airfoils are disclosed. The airfoils include a thicker leading airfoil portion and a thinner trailing airfoil portion. In one embodiment, the leading airfoil portion is formed by bending a body of the airfoil back toward itself. In another embodiment, the leading airfoil portion has a solid geometry and includes two elliptic surfaces. To prevent detachment of airflow, the leading airfoil portion includes at least two arc portions or surfaces that act to direct the airflow down to the trailing airfoil portion in a manner that stabilizes vortexes that may form in the region of changing thickness.

An integrated combustor nozzle includes an inner liner segment; an outer liner segment; and a panel extending radially between the inner and outer liner segments. The panel includes a forward end, an aft end, and a side walls extending axially from the forward end to the aft end. The aft end defines a turbine nozzle having a trailing edge circumferentially offset from the forward end. The inner liner segment has a pair of sealing surfaces, each of which defines a first continuous curve in the circumferential direction. The outer liner segment has a pair of sealing surfaces, each of which defines a second continuous curve in the circumferential direction. In some instances, the curves are monotonic in the circumferential direction. A segmented annular combustor including an array of such integrated combustor nozzles is also provided.

A damped turbine blade assembly for a gas turbine engine is disclosed. The damped turbine blade assembly includes a damper positioned within a first small slot of a first turbine blade and a second large slot of the second turbine blade. A portion of the damper can slidably mate with the second large slot providing a radial and angular connection between the first turbine blade and second turbine blade while allowing movement in a direction tangent to a radial of a center axis of the gas turbine engine. The tangential movement is resisted by friction between the damper contacting the second large slot and provides friction damping against vibrations felt by the turbine blades during operation of the gas turbine engine. The damper can be shaped and/or pre-stressed to control the normal force component of the friction between the damper and the second large slot.

A turbine wheel for a turbine of an exhaust gas turbocharger. The turbine wheel has a base body comprising a back wall and a plurality of turbine blades, connected to the base body, which each comprise a leading edge. The leading edges are designed to be wave shaped.

An exhaust turbine includes a turbine wheel having a plurality of rotor blades, an exhaust outlet where the exhaust outlet is arranged downstream of the rotor blades and is delimited radial to the outside by an axial turbine diffusor. An exhaust mass flow can be output in an axial flow direction by the exhaust outlet duct and the axial turbine diffusor opens downstream, radially to the outside at a non-constant diffusor-opening angle, such that at a first diffusor-opening angle deviates from a second diffusor-opening angle by at least 1°. The exhaust turbine further includes a waste gate duct, the outlet region of which opens into the diffusor or directly downstream of the diffusor into the exhaust outlet duct. The waste gate duct is designed to generate a substantially axial flow direction of a waste gate mass flow at the outlet region of the waste gate duct.

A cascade type thrust reverser device for a turbomachine of an aircraft, comprising a thrust reverser cascade and a casing, the cascade including first cavities, and the casing comprising an opening defining a housing wherein said cascade can be inserted in a first direction, and the casing and said cascade being in relative translation with respect to one another in the first direction between a first position of the device in which the cascade is entirely positioned in the housing and a second position of the device in which said cascade is at least partially outside said housing. The casing comprises an acoustic treatment panel including second cavities extending in a second plane parallel to the first plane, each first cavity facing a second cavity when the device is in the first position to form an acoustic treatment cell.

Various embodiments of an airfoil and machines with airfoils are disclosed. The airfoils include a thicker leading airfoil portion and a thinner trailing airfoil portion. In one embodiment, the leading airfoil portion is formed by bending a body of the airfoil back toward itself. In another embodiment, the leading airfoil portion has a solid geometry and includes two elliptic surfaces. To prevent detachment of airflow, the leading airfoil portion includes at least two arc portions or surfaces that act to direct the airflow down to the trailing airfoil portion in a manner that stabilizes vortexes that may form in the region of changing thickness.

An engine component includes a body having an internal surface and an external surface, the internal surface at least partially defining an internal cooling circuit. The component further includes a plurality of cooling holes formed in the body and extending between the internal cooling circuit and the external surface of the body. The plurality of cooling holes includes a first cooling hole with a metering portion and a diffuser portion extending from the metering portion to the external surface of the body.

An airfoil vane includes an airfoil section including an outer wall that defines an internal cavity; and a baffle situated in the internal cavity, the baffle including a baffle wall that defines a central cavity having a leading end and a trailing end corresponding to a leading end and a trailing end of the airfoil section, and a tail extending from the baffle wall, the tail including at least one feature configured to disturb an airflow surrounding the tail. A baffle for the airfoil vane assembly and a method of assembling a ceramic matrix composite airfoil vane are also disclosed.