Rolls-royce Corporation

Delta Biotechnology Limited

Gooee Limited

Mark J Wilson

Benedict R Phelps

Stephane M M Baralon

Gabriel Gonzalez-Gutierrez

Nigel H S Smith

Kashmir S Johal

Marco Barale

Craig W Bemment

Benjamin Mohankumar

Cesare A Hall

Shmuel Silverman

Richard A Johnson

Simon Coombes

Jonathan Couch

Steven James Burton

John Rodney Woodrow

Alan Victor Quirk

Stephen Berezenko

Patricia Carol Wood

Dan Andrei Tutunea

David St Coghlan

Alan Robert Maguire

Darrell Sleep

Peter Ireland

Glenn A Knight

Christopher Benson

Benjamin J Sellers

Yi Wang

Andrew Robert Goodey

Edward Hinchliffe

Michael O Hales

Miles Thomas Trumper

Daniel Robinson

Stephane Baralon

Nigel Smith

Robert C Chambers

George Bostock

Diane Joan Cousens

Susan Mary Kingsman

Hendrik Van Urk

Benedict Phelps

Cesare Hall

Glenn Knight

Miles Trumper

Robert Chambers

An aircraft engine comprising a fan, the fan having a diameter D and including a plurality of fan blades, the fan blades having a sweep metric S, each fan blade having a leading edge, and a forward-most portion on the leading edge of each fan blade being in a first reference plane. The aircraft engine further comprises a nacelle, comprising an intake portion forward of the fan, a forward edge on the intake portion being in a second reference plane, wherein the intake portion has a length L measured along an axis of the aircraft engine between the first reference plane and the second reference plane, the aircraft engine having a cruise design point condition M, wherein Mis between 0.4 and 0.93, L/D is between 0.2 and 0.45 and Sis from &#x2212;1 to 0.1.

Aircraft engine

An aircraft engine comprising a fan, the fan having a diameter D and including a plurality of fan blades, the fan blades having a sweep metric S, each fan blade having a leading edge, and a forward-most portion on the leading edge of each fan blade being in a first reference plane. The aircraft engine further comprises a nacelle, comprising an intake portion forward of the fan, a forward edge on the intake portion being in a second reference plane, wherein the intake portion has a length L measured along an axis of the aircraft engine between the first reference plane and the second reference plane, the aircraft engine having a cruise design point condition M, wherein Mis between 0.4 and 0.93, and L/D is between 0.2 and 0.45.

A fan blade for a gas turbine engine has a covered passage. A cross section through the fan blade at a point along the blade span is defined as having particular change in angle (&#x3b1;3&#x2212;&#x3b1;1) of the camber line between the leading edge and the trailing edge and/or between the leading edge and the point on the camber line that corresponds to the start of the covered passage.

Low speed fan up camber

A gas turbine engineis provided in which a fan having fan bladesin which the camber distribution relative to covered passage of the fanallows the gas turbine engine to operate with improved efficiency when compared with conventional engines, whilst retaining an acceptable flutter margin.

Gas turbine engine with optimized fan blade geometry

A gas turbine engine for an aircraft including: an engine core including a turbine, compressor, and core shaft connecting turbine to the compressor; a fan located upstream of engine core, the fan containing a plurality of fan blades mounted for rotation about an engine axis, each blade of the plurality of fan blades having a leading edge and trailing edge extending across a span of an airflow duct from blade root to tip; and a gearbox that receives an input from core shaft and outputs drive to fan so as to drive the fan at a lower rotational speed than the core shaft, wherein the trailing edge of each blade is characterized by a metric M defined as a rate of change of an angle of trailing edge between 0.4 and 0.8 of the span divided by an area averaged trailing edge angle, M being not less than around 4.

Gas turbine engine

A gas turbine engine has a cycle operability parameter &#x3b2; in a defined range to achieve improved overall performance, taking into account fan operability and/or bird strike requirements as well as engine efficiency. The defined range of cycle operability parameter &#x3b2; may be particularly beneficial for gas turbine engines in which the fan is driven by a turbine through a gearbox.

Aircraft engine operability

A gas turbine engineis provided in a fan root to tip pressure ratio, defined as the ratio of the mean total pressure of the flow at the fan exit that subsequently flows through the engine core (P) to the mean total pressure of the flow at the fan exit that subsequently flows through the bypass duct (P), is no greater than a certain value. The gas turbine enginemay provide improved efficiency when compared with conventional engines, whilst retaining an acceptable flutter margin.

Pressure ratio distributions for a gas turbine engine

A gas turbine engineis provided in which a fan having fan blades in which the camber distribution along the span allows the gas turbine engine to operate with improved efficiency when compared with conventional engines, whilst retaining an acceptable flutter margin.

A fan blade is provided with an aerofoil portion for which, for cross-sections through the aerofoil portion at radii between 15% and 25% of the blade span from the root radius, the average leading edge thickness is greater than the leading edge thickness at the tip. The geometry of the fan blade may result in a lower susceptibility to flutter.

Gas turbine engine fan blade

A fan blade is provided with an aerofoil portion for which, at radii between 20% and 40% of the blade span, the location of the position of maximum thickness along the camber line is at less than a defined percentage of the total length of the camber line. For all cross-sections through the aerofoil portion at radii greater than 70% of the blade span, the location of the position of maximum thickness along the camber line is at more than a defined percentage of the total length of the camber line. The geometry of the fan blade may result in a lower susceptibility to flutter.

A gas turbine engine comprising a compressor, a compressor case surrounding the compressor and a compressor tip injector system is disclosed. The compressor tip injector system comprises a cabin blower system comprising a cabin blower compressor arranged in use to compress air used in a cabin of an aircraft and by the compressor tip injector system. The compressor case comprises one or more injectors of the compressor tip injector system through which in use air from the cabin blower compressor is injected towards blade tip ends of blades of the compressor as they rotate.

Compressor tip injector

The disclosed devices, systems, and methods may be used to automatically identify, locate, and assign luminaires into groups such that lighting systems may be more efficiently configured, used, and maintained especially in large buildings, etc. For example, a Lucas-Kanade function may be used with a system of sensors which are capable of detecting light patterns from individual and groups of luminaires to form virtual maps of luminaire locations which may be correlated with actual luminaire floor plans to efficiently identify, locate, and group the luminaires.

Automated luminaire identification and group assignment devices, systems, and methods using dimming function

The disclosed devices, systems, and methods may be used to automatically identify, locate, and assign luminaires into groups such that lighting systems may be more efficiently configured, used, and maintained especially in large buildings, etc. For example, a wink function may be used with a system of sensors which are capable of detecting light patterns from individual and groups of luminaires to form virtual maps of luminaire locations which may be correlated with actual luminaire floor plans to efficiently identify, locate, and group the luminaires.

Automated luminaire identification and group assignment

Growing community of inventors

Nottingham, United Kingdom

Mark J Wilson