An air inlet section for an enclosure for an aircraft engine is provided that includes an inner barrel panel, an outer barrel panel, a lipskin and a forward bulkhead. The lipskin extends between an inner barrel end and an outer barrel end. The inner barrel end is disposed proximate the forward end of the inner barrel panel and the outer barrel end is disposed proximate the forward end of the outer barrel panel. The forward bulkhead has a panel that extends between an outer radial end and an inner radial end. The inner barrel panel, the outer barrel panel, and the lipskin define an interior annular region, and the forward bulkhead defines a sub-portion of interior annular region. The outer radial end of the forward bulkhead panel is disposed forward of the inner radial end of the forward bulkhead panel.

Hybrid propulsion is obtained by arranging, around a conventional turbomachine, a reversible electric machine linked to the low-pressure shaft via a mechanical movement transmission, a one-way clutch being arranged on the low-pressure shaft between the fan and a low-pressure compressor. The low-pressure shaft includes two aligned portions which can be separated and reconnected. A casing includes a first portion able to tilt around a horizontal transverse axis of the aircraft relative to a second portions of the casing. One portion of the low-pressure shaft, the fan and the reversible electric machine are mounted on the first portion of the casing, and the other portion of the low-pressure shaft is mounted on the second portion of the casing. This device facilitates electric taxiing of the aircraft, using the machine as a motor, whereas the rear of the low-pressure shaft remains stationary.

A gas turbine engine that includes a fan nacelle and a core nacelle that provide a bypass flow path radially between. The fan nacelle has an inlet including a throat. The inlet has an inlet forward-most point. A fan is arranged in the bypass flow path and rotatable about an axis. The fan has a leading edge recessed from the inlet forward-most point an inlet length in an axial direction. A spinner has a spinner length from a spinner forward-most point to the leading edge. A ratio of the spinner length to inlet length is equal to or greater than about 0.5.

A nacelle for a turbomachine is provided and includes an inner wall that is annular about a longitudinal axis of the nacelle, the annular inner wall being designed to surround part of the turbomachine. The nacelle further includes an annular outer wall surrounding the annular inner wall. The annular outer wall includes a first acoustically porous part, and the annular inner wall including a second acoustically porous part. The first and second porous parts are arranged facing one another so as to allow soundwaves, emitted by the turbomachine housed in the nacelle, to pass through the annular inner wall and then the annular outer wall so as to escape from the nacelle.

A variable geometry inlet system of an aircraft engine includes an inlet duct. The inlet duct includes at least first and second sections moveable between extended and retracted positions such that the inlet duct defines a variable axial length of an inlet passage for selective flight conditions. The inclusion of acoustic treatment may assist in controlling noise.

An aircraft comprises a machine body. The machine body encloses a turbofan gas turbine engine and a plurality of ancillary systems. The turbofan gas turbine engine comprises, in axial flow sequence, a heat exchanger module, a fan assembly, a compressor module, a combustor module, a turbine module, and an exhaust module. The machine body comprises a single fluid inlet aperture, with the fluid inlet aperture being configured to allow a fluid cooling flow to enter the machine body and to pass through the heat exchanger module. The heat exchanger module is configured to transfer a waste heat load from the gas turbine engine and the ancillary systems to the fluid cooling flow prior to an entry of the entire fluid cooling flow into the fan module.

A turbofan gas turbine engine includes, in axial flow sequence, a heat exchanger module, a fan assembly, a compressor module, a turbine module, and an exhaust module. The fan assembly includes a plurality of fan blades defining a fan diameter (D). The heat exchanger module is in fluid communication with the fan assembly by an inlet duct, and the heat exchanger module includes a plurality of radially-extending hollow vanes arranged in a circumferential array with a channel extending axially between each pair of adjacent hollow vanes. An airflow entering the heat exchanger module is divided between a set of vane airflows through each of the hollow vanes and a set of channel airflows through each of the channels.

An aircraft turbojet engine nacelle comprising an air intake assembly of substantially annular shape, the air intake assembly comprising an external wall and an internal wall which are connected to one another upstream by an air intake lip and downstream by an annular rear structure. The annular rear structure comprises, arranged concentrically, one or more monolithic panels forming a first annular zone and comprising one or more hollow casings forming at least a second annular zone, and at least one connecting ring which provides the connection with at least one of the external and internal walls. The annular rear structure thus has increased structural mechanical strength.

An aircraft includes a machine body that encloses a turbofan gas turbine engine and a plurality of ancillary systems. The turbofan gas turbine engine includes, in axial flow sequence, a first heat exchanger module, a fan assembly, a compressor module, a combustor module, a turbine module, and an exhaust module. The aircraft includes a second heat exchanger module. The machine body comprises a single fluid inlet aperture, with the fluid inlet aperture being configured to allow a fluid cooling flow to enter the machine body and to pass through the first heat exchanger module. When a temperature of the fluid cooling flow is less than a temperature of a fluid to be cooled, the fluid to be cooled is directed to the first heat exchanger module, and when a temperature of the fluid cooling flow is greater than a temperature of the fluid to be cooled, the fluid to be cooled is directed to the second heat exchanger module and cooled using a fuel supply for the gas turbine engine.

A nacelle of an aircraft propulsion assembly includes a front section and a main section. The nacelle also includes a junction flange between the front section and the main section including a plurality of deformable damper elements distributed along the inner peripheral edge of a stiffening frame of the nacelle. The damper element allows significant stresses to be absorbed between the engine flange and the stiffening frame. This configuration is particularly adapted to large nacelles. An aircraft propulsion assembly including such a nacelle is also described.