A propulsion system for an aircraft (10) is provided having an aft engine configured to be mounted to the aircraft (10) at an aft end of the aircraft (10). The aft engine includes a fan (304) rotatable about a central axis of the aft engine having a plurality of fan blades (310). The aft engine also includes a nacelle (306) encircling the plurality of fan blades (310) with one or more structural members (308) extending between the nacelle (306) and the mean line (15) of the aircraft (10) at a location forward of the plurality of fan blades (310) when the aft engine is mounted to the aircraft (10). The aft engine may increase a net thrust of the aircraft (10) when mounted to the aircraft (10).

Aft engine for an aircraft

According to an example embodiment, a gas turbine engine assembly includes, among other things, a fan that has a plurality of fan blades. A diameter of the fan has a dimension D that is based on a dimension of the fan blades. Each fan blade has a leading edge. An inlet portion is situated forward of the fan. A length of the inlet portion has a dimension L between a location of the leading edge of at least some of the fan blades and a forward edge on the inlet portion. A dimensional relationship of L/D is between about 0.2 and about 0.45.

Low pressure ratio fan engine having a dimensional relationship between inlet and fan size

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.

Gas turbine engine inlet

An inlet (18) of a nacelle (16) for channeling inlet airflow to a fan (28) of a gas turbine engine (12) includes a forward lip (40) and an inner surface (46). The forward lip extends around a centerline axis (20) of the engine and is adapted to accommodate a flow angle (β) of incoming airflow. The inner surface extends from the forward lip along the centerline axis (20) of the engine to adjacent the fan. The inner surface has a first profile (52) above the centerline axis of the engine and a second profile (54) below the centerline axis of the engine. The first profile and the second profile are adapted to define an inlet centerline axis (56) that extends below the centerline axis of the engine at a face (50) of the fan.

Cambered aero-engine inlet

An aircraft including a fuselage and an aft engine is provided. The fuselage extends from a forward end of the aircraft towards an aft end of the aircraft. The aft engine is mounted to the fuselage proximate the aft end of the aircraft and includes a fan and a nacelle. The fan is rotatable about a central axis of the aft engine and includes a plurality of fan blades. The nacelle of the aft engine surrounds the plurality of fan blades and defines a bottom portion having a forward end. Additionally, the nacelle defines a curved surface at the forward end of the bottom portion, the curved surface including a reference point where the curved surface defines the smallest radius of curvature. The nacelle further defines a normal reference line extending normal from the reference point. The normal reference line defines an angle with the central axis of the aft engine greater than zero to, e.g., allow for a maximum amount of airflow into the aft engine.

Aft engine nacelle shape for an aircraft

A nacelle for a turbofan gas turbine engine is provided. An inner wall of the nacelle defines an air intake which directs air into the fan section of the engine. The intake has, in flow series, an intake lip, a throat and a diffuser. The diffuser has, in flow series, first and second flow conditioning sections over both of which the flow cross-sectional area of the diffuser increases with increasing downstream distance from the throat. In addition, over the second section the nacelle inner wall lies substantially on a surface of an oblique circular cone having an apex which is offset from the centreline of the engine.

Gas turbine engine nacelle

A method of designing a nacelle for a gas turbine engine includes defining an initial geometry for the nacelle, pivoting, at each of one or more positions on a highlight and in a plane of a respective longitudinal section, a front portion of a section of an intake lip lying in the respective longitudinal section about the highlight, and obtaining an altered geometry for the nacelle by adjusting surfaces of the intake neighboring the pivoted front portion to smoothly blend the surfaces of the pivoted front portion to unadjusted surfaces of the intake further removed from the pivoted front portion.

Gas turbine engine nacelle having a symmetric flowpath and design method thereof

An initial geometry for a nacelle provides an intake at the upstream end of the nacelle and an exhaust at the downstream end of the nacelle. The intake has in flow series, an intake lip and a diffuser. The intake lip has a highlight defining the boundary between the inner and outer surfaces of the intake lip. On all longitudinal sections containing the axis of the engine (i) the inner and outer surfaces are tangency matched at the highlight, and (ii) the inner and outer surfaces have their maximum curvatures at the highlight.

Dimitris Arzoglou

Papers

Gas turbine engine nacelle

Gas turbine engine nacelle having a symmetric flowpath and design method thereof

Gas turbine engine nacelle having a symmetric flowpath