US 2020/0400080 A1Published application

Gas turbine engine with highly efficient fan

AssigneeROLLS-ROYCE plc

InventorsCraig W. BEMMENT, Pascal DUNNING

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Abstract

A gas turbine engine for an aircraft includes an engine core with a turbine, compressor, and core shaft connecting the two; and a fan upstream of the core with a plurality of blades extending from a hub each with a leading and trailing edge, wherein fan tip radius is between the engine centreline and each blade's leading edge outermost tip and hub radius is between the engine centreline and the hub's outer surface at each blade's leading edge radial position, the ratio of hub to tip radius between 0.2 and 0.285. A fan rotor entry temperature is the average temperature of airflow across the leading edge of each blade at cruise conditions and a fan rotor exit temperature is an average temperature of airflow across a radially outer portion of each blade at the trailing edge at cruise conditions, the ratio of entry to exit temperature between 1.11 and 1.05.

Claims22

Dependent claims

A gas turbine engine for an aircraft comprising:

an engine core comprising a turbine, a compressor, a core shaft connecting the turbine to the compressor, and an annular splitter at which the flow is divided between a core flow that flows through the engine core, and a bypass flow that flows along a bypass duct, wherein stagnation streamlines around the circumference of the engine, stagnating on a leading edge of the annular splitter, form a streamsurface forming a radially inner boundary of a streamtube that contains all of the bypass flow;
and
a fan located upstream of the engine core, the fan comprising a plurality of fan blades extending from a hub, each fan blade having a leading edge and a trailing edge, wherein a fan tip radius of the fan is defined between a centreline of the engine and an outermost tip of each fan blade at its leading edge and a hub radius is defined between the centreline of the engine and an outer surface of the hub at a blade root at the radial position of the leading edge of each fan blade, each fan blade having a radially outer portion lying within the streamtube that contains the bypass flow, and wherein a fan rotor entry temperature is defined as an average temperature of airflow across the leading edge of each fan blade at cruise conditions and a fan rotor exit temperature is defined as an average temperature of airflow across a radially outer portion of each fan blade at the trailing edge at cruise conditions;
wherein a fan hub to tip ratio of:
the
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fan
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hub
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radius
the
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fan
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tip
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radius
is in the range from 0.2 to 0.285;
and
a fan tip temperature rise of:
the
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fan
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tip
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rotor
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exit
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temperature
the
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fan
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rotor
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entry


temperature
is in the range from 1.11 to 1.05, and wherein
a fan tip loading at cruise conditions is defined as dH/Utip2, where dH is the enthalpy rise across the fan and Utip is the (translational) velocity of the fan tip, and the fan tip loading is in the range from 0.25 to 0.4.

55 more elements

A gas turbine engine for an aircraft comprising:

an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor, wherein the engine core has a core radius defined between the centreline of the engine and a forwardmost tip of the engine core;
and
a fan located upstream of the engine core, the fan comprising a plurality of fan blades extending from a hub, each fan blade having a leading edge and a trailing edge, wherein a fan tip radius of the fan is defined between a centreline of the engine and an outermost tip of each fan blade at its leading edge and a hub radius is defined between the centreline of the engine and an outer surface of the hub at a blade root at the radial position of the leading edge of each fan blade, a fan rotor entry temperature is defined as an average temperature of airflow across the leading edge of each fan blade at cruise conditions and a fan tip rotor exit temperature is defined as an average temperature of airflow across a radially outer portion of each fan blade at the trailing edge at cruise conditions, wherein the radially outer portion of each fan blade is or comprises the portion of each fan blade at a distance from the centreline of the engine greater than the core radius;
wherein a fan hub to tip ratio of:
the


fan


hub


radius
the


fan


tip


radius
is in the range from 0.2 to 0.285;
and
a fan tip temperature rise of:
the


fan


tip


rotor


exit


temperature
the


fan


rotor


entry


temperature
is in the range from 1.11 to 1.05, and wherein
a fan tip loading at cruise conditions is defined as dH/Utip2, where dH is the enthalpy rise across the fan and Utip is the (translational) velocity of the fan tip, and the fan tip loading is in the range from 0.25 to 0.4.

55 more elements

Dependencies are inferred from the claim text.

Facts

Application: 16/545,129Filed 20 Aug 2019
Priority: GB 1908971.324 Jun 2019
Published: 24 Dec 2020
Term ends: Not estimatedThe source records don’t support an estimate.

Classification12 codes

F02C 9/20
Control of working fluid flow: by throttling; by adjusting vanes
F02K 3/06
The plant including ducted fans, i.e. fans with high volume, low pressure outputs, for augmenting the jet thrust, e.g. of double-flow type: with…
F02C 9/28
Control of fuel supply: regulating systems responsive to plant or ambient parameters, e.g. temperature, pressure, rotor speed
F05D 2220/323
In gas turbines: for aircraft propulsion, e.g. jet engines

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