Abstract: 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 (?p) with respect to the longitudinal centre line of the gas turbine engine.

Abstract: A system and a method of designing a nacelle for a gas turbine engine having a longitudinal centre line. The nacelle includes an air intake, an internal surface, an azimuthal angle, and a plurality of intake lines. The air intake comprises, in flow series, an intake lip, a throat and a diffuser. The internal surface at least partially defines the air intake. The azimuthal angle is defined about the longitudinal centre line. The intake lines extend along the internal surface of the nacelle at respective values of the azimuthal angle. Each intake line axially defines the air intake along the longitudinal centre line at the respective value of the azimuthal angle. The internal surface of the nacelle between the plurality of intake lines at a given axial location along the longitudinal centre line is analytically defined by an equation.

Abstract: A gas turbine engine (100) for an aircraft comprises a pylon attachment (112) and a shaft (108) defining an engine centreline (110). The engine centreline lies in an engine central plane (120) which intersects the pylon attachment. The gas turbine engine comprises an intake (104) having a non-axisymmetric geometry and a medial plane (130) defining left and right halves of the intake. The left and right halves are configured for at least one of optimum cross wind performance, optimum incidence performance and optimum cruise performance when the medial plane is aligned with a vertical plane. The intake is installed so that the medial plane is angularly offset with respect to the engine central plane. The engine may be installed on a wing of an aircraft with the medial plane closer to its optimal orientation than is the case for a conventional engine.

Abstract: 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 diffuser further includes a diffuser angle (?diff), indicating a degree of divergence of the diffuser relative to the longitudinal centre line. The diffuser angle (?diff) is from about 0 degrees to about 12 degrees.

Abstract: A nacelle for housing a fan within a gas turbine engine having a longitudinal centre line includes a leading edge and a trailing edge. A nacelle length (Lnac) is defined as an axial distance between the leading edge and the trailing edge. An azimuthal angle (?) is defined about the longitudinal centre line. The nacelle length (Lnac) varies azimuthally. The nacelle length (Lnac) decreases azimuthally from an inboard end of the nacelle to an outboard end of the nacelle.

Abstract: A nacelle for a gas turbine engine having a longitudinal centre line includes an intake lip disposed at an upstream end of the nacelle. The intake lip includes a crown and a keel. The crown includes a crown leading edge and the keel includes a keel leading edge. The crown leading edge and the keel leading edge define a scarf line therebetween. The scarf line forms a scarf angle (?scarf) relative to a reference line perpendicular to the longitudinal centre line. A fan casing is disposed downstream of the intake lip and includes a casing leading edge. The casing leading edge defines a droop line normal to the casing leading edge. The droop line forms a droop angle (?droop) relative to the longitudinal centre line. A relationship between the droop angle (?droop) and the scarf angle (?scarf) is given by: ?droop=?scarf/1.5±1 degree.

Abstract: 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 diffuser further includes a diffuser angle (?diff), indicating a degree of divergence of the diffuser relative to the longitudinal centre line. The diffuser angle (?diff) is from about 0 degrees to about 12 degrees.

Abstract: 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 (?p) with respect to the longitudinal centre line of the gas turbine engine.

Abstract: A nacelle for housing a fan within a gas turbine engine having a longitudinal centre line includes a leading edge and a trailing edge. A nacelle length (Lnac) is defined as an axial distance between the leading edge and the trailing edge. An azimuthal angle (?) is defined about the longitudinal centre line. The nacelle length (Lnac) varies azimuthally. The nacelle length (Lnac) decreases azimuthally from an inboard end of the nacelle to an outboard end of the nacelle.

Abstract: A gas turbine engine for an aircraft includes an engine core, fan, air intake and gearbox. The engine core includes a turbine, compressor, and core shaft connecting the turbine and compressor. The fan is located upstream of the core and includes a plurality of fan blades, the fan having a diameter greater than 2.0 m. The air intake is located upstream of the fan and has ratio of intake length to fan diameter of 0.20 to 0.60 and defines highlight, throat and diffuser areas. The gearbox receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft. The gas turbine engine has a bypass ratio greater than 10; and the air intake has an intake length and the ratio of the intake length to the fan diameter is from 0.20 to 0.60.

Abstract: A gas turbine engine for an aircraft includes an engine core, a fan, an air intake and a gearbox. The engine core includes a turbine, a compressor, and a core shaft connecting the turbine to the compressor. The fan is upstream of the engine core and includes a plurality of fan blades, and has a diameter greater than 2.0 m. The air intake is upstream of the fan and has ratio of intake length to fan diameter of 0.20 to 0.60. The gearbox receives an input from the core shaft and drives the fan at a lower rotational speed than the core shaft. The gas turbine engine has a bypass ratio greater than 10; and the air intake defines a highlight area, wherein the ratio of highlight area to throat area is from 1.15 to 1.35 and the ratio of diffuser area to throat area is from 0.85 to 1.15.

Abstract: A gas turbine engine for an aircraft has an engine core, fan, air intake and gearbox. The engine core has a turbine, compressor, and core shaft connecting them. The fan is upstream of the engine core and has fan blades, the fan having a diameter greater than 2.0 m. The air intake is upstream of the fan and has ratio of intake length to fan diameter of from 0.20 to 0.60. The gearbox receives an input from the core shaft and outputs drive to the fan to drive the fan at a lower rotational speed than the core shaft. The gas turbine engine has a bypass ratio greater than 10; and the air intake defines highlight, throat and diffuser areas, wherein the gas turbine engine has a contraction ratio from 1.10 to 1.35, the contraction ratio being the ratio of the highlight area to the throat area.

Abstract: A mounting arrangement for mounting an aircraft gas turbine engine to an aircraft includes an engine nacelle with a distal assembly including a part annular engine cowl, a gas turbine engine core housing surrounded by the cowl and a distal bifurcation extending between the engine core housing and engine cowl in a first direction to define a first axis. The mounting arrangement includes a proximal assembly having a mount configured to mount the proximal assembly to the engine core housing. The proximal assembly includes a pylon configured to mount the proximal assembly to mounting location such as a wing of the aircraft at an engine mounting location. The pylon extends in a line between the wing and the engine core housing to define a second axis which is normal to a distal surface of the wing at the engine mounting location and is non-parallel to the vertical axis.

Abstract: A gas turbine engine comprises a pylon attachment, a shaft defining an engine centreline which lies in an engine central plane intersecting the pylon attachment, a fan defining a fan plane normal to the engine centreline and an intake upstream of the fan plane. The geometric centreline of the intake coincides with the engine centreline at an axial position corresponding to the downstream end of the intake and curves away from the engine centreline upstream of said axial position. The engine may be mounted on one side of an aircraft such that the orientation of the highlight plane of the intake is aligned to the air flow field of the aircraft on that side during flight.

Abstract: A gas turbine engine (100) for an aircraft comprises a pylon attachment (112) and a shaft (108) defining an engine centreline (110). The engine centreline lies in an engine central plane (120) which intersects the pylon attachment. The gas turbine engine comprises an intake (104) having a non-axisymmetric geometry and a medial plane (130) defining left and right halves of the intake. The left and right halves are configured for at least one of optimum cross wind performance, optimum incidence performance and optimum cruise performance when the medial plane is aligned with a vertical plane. The intake is installed so that the medial plane is angularly offset with respect to the engine central plane. The engine may be installed on a wing of an aircraft with the medial plane closer to its optimal orientation than is the case for a conventional engine.

Abstract: A heat sink for location in a fluid flow, including a heat sink base and a plurality of heat dissipating elements, such as elongate fins, extending from the surface of the heat sink base. In certain arrangements the heat sink is provided with a diversion flow passageway for diverting a fraction of fluid flow away from the heat dissipating elements. In other arrangements there may be two arrays of elongate fins laterally offset. In yet a further arrangement the heat sink may be configured to promote the generation of at least one vortex.

Abstract: A gas turbine engine including an intake, a fan and an injector system. The intake has an inner wall which defines an intake passage for the fan. The injector system includes a cabin blower system including a cabin blower compressor arranged in use to compress fluid used in a cabin of an aircraft and by the injector system. The intake includes an injector of the injector system through which in use fluid from the cabin blower compressor is injected into a main airflow for flow control of air on the way to the fan.

Abstract: A fluid impingement arrangement comprising a supply manifold and at least one nozzle exit coupled to the supply manifold. The nozzle exit is arranged as a Coanda surface having a restriction and has at least one static pressure tapping that cross-connects two regions of the restriction to induce passive oscillation in a fluid jet passing through the nozzle exit.

Abstract: A boundary layer energizer (20) for energizing a boundary layer flow over a surface (22), the boundary layer energizer (20) comprising a plurality of passages (24), each passage terminating in a respective hole (26) provided on the surface (22), the holes being arranged in a cluster (23) on the surface, wherein the plurality of passages are angled with respect to one another at the surface such that, when in use, a vortex (24) is formed by a fluid flowing through the plurality of passages.

Abstract: A fluid transfer arrangement comprising a duct having a first end and a second end, a pulse generation mechanism located at the first end of the duct to direct fluid pulses towards the second end of the duct in use, and a baffle located at the second end of the duct that defines an aperture having sharp edges. The sharp edges generate ring vortex fluid flow from the aperture in use. Applications include impingement heating and cooling.