A nacelle having a fan casing, a cowl that movable between an advanced position and a retracted position that opens an opening between a bypass duct and the outside, deflectors secured to the mobile cowl, wherein, in the advanced position, they are around the fan casing and wherein, in the retracted position, they are across the opening, and a fan ramp with a mounting base and flaps that are able to rotate on the mounting base between a stowed position and a deployed position. For each flap, the fan ramp has a return element that urges the flap) into the deployed position, and the deflectors have a stop in contact with the flap when the mobile cowl passes from the retracted position to the advanced position. Thus, in the advanced position, the flaps are folded back and their bulk is reduced.

An anti-icing system of a nacelle inlet of an engine of an aircraft includes first and second direct acting valves and first and second control valve assemblies fluidly connected to the nacelle inlet. The first direct acting valve includes a first inlet, outlet, valve chamber, and piston. The first piston is positioned in the first direct acting valve. The first control valve assembly is fluidly connected to the first valve. The second direct acting valve includes a second inlet, outlet, valve chamber, and piston. The second piston is positioned in the second direct acting valve. The second direct acting valve is fluidly connected to the first direct acting valve in a series configuration. The second control valve assembly is fluidly connected to the second valve chamber.

An assembly for an aircraft propulsion system includes a variable area inlet with a fixed structure and a moveable structure. The variable area inlet is configured to open and close an airflow inlet passage into the aircraft propulsion system. The moveable structure is configured to move axially along a centerline between an aft position and a forward position. The moveable structure includes an inlet lip structure and a deflector. When the moveable structure is in the aft position, the airflow inlet passage is closed, and the deflector is at least partially recessed into the fixed structure. When the moveable structure is in the forward position, the airflow inlet passage is opened axially between an aft end of the inlet lip structure and a forward end of the fixed structure, and a forward end of the deflector is disposed axially at the forward end of the fixed structure.

A nacelle for a gas turbine engine having a longitudinal centre line. The nacelle includes an air intake disposed at an upstream end of the nacelle. The air intake includes, in flow series, an intake lip, a throat and a diffuser. The nacelle further includes a protrusion extending radially inward from the air intake downstream of the intake lip. The protrusion extends circumferentially by a protrusion angle (θ.sub.p) with respect to the longitudinal centre line of the gas turbine engine.

A method for repairing a turbomachine air intake lip including a damaged area whose shape differs from the initial shape of the lip, includes identifying the damage; depositing a filler material on the damaged area so as to give it a shape identical to the initial shape; inspecting the shape of the lip; and inspecting the performed repair.

A nacelle forward part of an aircraft propulsion assembly, includes an air inlet lip at the front end, an internal structure and an external panel extending the lip, and an annular stiffening frame formed around an axis of revolution and comprising an internal peripheral edge linked to the internal structure and an external peripheral edge linked to the external panel. The stiffening frame presents an inclination relative to the axis of revolution, its slope being oriented towards the front end of the nacelle forward part. The stiffening frame is configured to deviate, outside the nacelle of the propulsion assembly, an object striking the stiffening frame. The stiffening frame so configured is capable of withstanding the impacts of objects without breaking while enabling the object striking the stiffening frame to be deviated towards the exterior of the nacelle.

A gas turbine engine has a compression system radius ratio defined as the ratio of the radius of the tip of a fan blade to the radius of the tip of the most downstream compressor blade in the range of from 5 to 9. This results in an optimum balance between installation benefits, operability, maintenance requirements and engine efficiency when the gas turbine engine is installed on an aircraft.

Nacelle inlet structures, engine assemblies and vehicles including the same, and related methods. A nacelle inlet structure of an engine assembly includes an inlet outer barrel and an inlet inner barrel with a tubular portion and an inlet attachment flange. The tubular portion of the inlet inner barrel extends at least partially along a direction parallel to the engine axis, while the inlet attachment flange extends from the tubular portion. The inlet attachment flange is integrally formed with at least a portion of the tubular portion of the inlet inner barrel. The inlet attachment flange is configured to be operatively coupled to an engine case of the engine assembly to operatively couple the nacelle inlet structure to the engine case. In examples, a method of manufacturing an inlet inner barrel includes forming a composite laminate with a composite manufacturing process.

Apparatuses and systems are provided herein for unducted propulsion systems. The system includes an aft housing for low drag for high subsonic sustained flight. A plurality of blades are affixed to the aft housing, wherein the housing defines a flowpath curve extending from the axial extent of the aft blade root to the aft end of the aft housing. The flowpath curve is described by an axial direction parallel to an axis of rotation and a radius from the axis of rotation. The flowpath curve includes first point having a first radius where the radius reaches a maximum aft of the aft blade root and a second point forward of the first point having a second radius where the radius stops decreasing. The ratio of the first radius to the second radius is greater than or equal to 1.081.

A ducted fan engine module and method of assembling the same is disclosed. In various embodiments, the ducted fan engine module includes a ducted fan engine having a ducted fan case configured to house a fan; a lower V-blade fitting secured to a lower frame; and an upper V-blade fitting secured to an upper frame.