Rolls-Royce Corporation

David G MacManus

Christopher T J Sheaf

Robert E Christie

Fernando L Tejero Embuena

Christopher Tj Sheaf

Zahid Muhammad Hussain

Andrew J Mullender

Brian Handley

Angus Roy Smith

Christopher Thomas John Sheaf

Giles Anthony Richardson

A nacelle for a gas turbine engine includes a leading edge at an upstream side of the nacelle. The nacelle further includes a trailing edge at a downstream side of the nacelle. The nacelle further includes an outer surface disposed between the leading edge and the trailing edge. The nacelle further includes a concave section continuous with the outer surface and disposed proximal to the trailing edge. The concave section includes an upstream end and a downstream end spaced apart from the upstream end. The concave section is curved radially inwards relative to the outer surface between the upstream end and the downstream end.

Nacelle for gas turbine engine and aircraft comprising the same

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 (&#x3b8;) with respect to the longitudinal centre line of the gas turbine engine.

Nacelle for gas turbine engine

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.

Gas turbine engine nacelle and method of designing same

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 (&#x3b8;), indicating a degree of divergence of the diffuser relative to the longitudinal centre line. The diffuser angle (&#x3b8;) is from about 0 degrees to about 12 degrees.

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 (L) is defined as an axial distance between the leading edge and the trailing edge. An azimuthal angle (&#x3b8;) is defined about the longitudinal centre line. The nacelle length (L) varies azimuthally. The nacelle length (L) decreases azimuthally from an inboard end of the nacelle to an outboard end of the nacelle.

An intake centre fairing for a gas turbine engine includes a body. The body includes an outer surface, an apex point and a base. The apex point is at a first end of the body and the base is at a second end of the body. The base includes a base centre. The body defines a longitudinal axis along its length, a radial direction relative to the longitudinal axis and a circumferential direction relative to the longitudinal axis. The outer surface of the body is tapered from the base to the apex point along the longitudinal axis. The apex point is radially offset relative to the base centre along the radial direction. The apex point is further circumferentially offset relative to the base centre along the circumferential direction.

Intake centre fairing for a gas turbine engine

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 (&#x3b8;) 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 (&#x3b8;) relative to the longitudinal centre line. A relationship between the droop angle (&#x3b8;) and the scarf angle (&#x3b8;) is given by: &#x3b8;=&#x3b8;/1.5&#xb1;1 degree.

Nacelle for a gas turbine engine

A nacelle for a turbofan gas turbine engine, having in flow series an intake lip, a diffuser and a fan casing. The diffuser has, in flow series, a main section and a recessed section adjacent to the fan casing. A transition region, such as a step, slope or curve, is provided between the main section and the recessed section. In some arrangements the transition region may be configured to promote flow separation and the formation of a separation bubble in the recessed section.

Nacelle

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.

Fluid impingement arrangement

A gas turbine engine having a rotational axis, an intake and a compressed gas source; the intake includes a lining having a facing which defines an inlet surface and an array of holes; the array of holes includes at least a first set of holes and a second set of holes, the holes of the first set of holes are angled a relative to a radial line and the holes of the second set of holes are angled &#x3b2; relative to the radial line; the angles &#x3b1; and &#x3b2; are different. An active flow control arrangement including a compressed gas supply pipe, a valve arrangement, a controller, a compressed gas distribution pipe may be provided. Compressed gas may be provided to prevent the formation of separation of a main gas flow through the intake and prevent or remove ice accretion.

Intake liner for a gas turbine engine

An arrangement for reducing vorticity downstream of a turbine nozzle guide vanein a gas turbine engine. The arrangement including directing cooling air through the guide vaneand out through outletsin directions to counteract the wave vorticity produced downstream of the guide vane

Growing community of inventors

Olney, United Kingdom

David G MacManus