Fan arrangement for a gas turbine engine
Abstract
A gas turbine engine for an aircraft including: engine core including a turbine; and fan including a plurality of fan blades extending radially from a hub, each fan blade having a leading and trailing edge. Turbine includes a lowest pressure turbine stage having a row of rotor blades each extending radially and having a leading and trailing edge. A fan-turbine radius difference is measured as radial distance between: a point on a circle swept by a radially outer tip of the trailing edge of each of the rotor blades of the lowest pressure stage of the turbine; and a point on a circle swept by a radially outer tip of the leading edge of each of fan blades; and a fan speed to fan-turbine radius ratio defined as: the maximum take - off rotational speed of the fan fan - turbine radius difference is in a range between 0.8 rpm/mm to 5 rpm/mm.
Claims
exact text as granted — not AI-modified1 . 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; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; a nacelle surrounding the gas turbine engine, the nacelle comprising an inner surface at least partly defining a bypass duct located radially outside of the engine core; and a bypass duct outlet guide vane extending radially across the bypass duct between an outer surface of the engine core and the inner surface of the nacelle, wherein the bypass duct outlet guide vane extends between a radially inner tip and a radially outer tip and has a leading edge and a trailing edge relative to the direction of gas flow through the bypass duct, an outer wall axis is defined joining the radially outer tip of the trailing edge of the bypass duct outlet guide vane and the rearmost tip of the inner surface of the nacelle, wherein the outer wall axis lies in a longitudinal plane containing the centreline of the gas turbine engine, an outer bypass duct wall angle is defined as the angle between the outer wall axis and the centreline, and the outer bypass duct wall angle is in a range from −15 to +1 degrees.
2 . The gas turbine engine of claim 1 , wherein the outer bypass duct wall angle is in a range between −5 degrees and −1 degrees.
3 . The gas turbine engine of claim 1 , wherein the outer bypass duct wall angle is in a range between −4.0 degrees and −1.0 degrees.
4 . The gas turbine engine of claim 1 , wherein the outer bypass duct wall angle is:
a) between −0.5 and −4, and optionally wherein a fan tip radius of the gas turbine engine is in the range from 110 cm to 150 cm; or b) is in a range between −2.5 degrees and −4 degrees, and optionally a fan tip radius of the gas turbine engine is in the range from 155 cm to 200 cm.
5 . The gas turbine engine of claim 1 , wherein a negative value of the outer bypass duct wall angle corresponds to the outer wall axis sloping towards the centreline of the gas turbine engine.
6 . The gas turbine engine of claim 1 , wherein a bypass duct outlet guide vane radius, measured radially between the engine centreline and the radially outer tip of the trailing edge of the bypass outlet guide vane, is in a range from 90 cm to 210 cm, and optionally:
a) the gas turbine engine has a fan tip radius in the range from 110 cm to 150 cm and the bypass duct outlet guide vane radius is in the range from 90 cm to 150 cm, optionally 110 cm to 135 cm; or b) the gas turbine engine has a fan tip radius in the range from 155 cm to 200 cm and the bypass duct outlet guide vane radius is in the range from 160 cm to 210 cm, optionally 170 cm to 200 cm.
7 . The gas turbine engine of claim 1 , wherein the rearmost inner tip of the nacelle inner wall is movable to provide a variable area bypass exhaust nozzle, and wherein the outer wall axis is defined based on the position of the rearmost tip of the inner surface of the nacelle during cruise conditions.
8 . The gas turbine engine of claim 1 , wherein the fan tip radius of the fan is measured between a centreline of the engine and an outermost tip of each fan blade at its leading edge; and the nacelle is arranged to surround the fan and the engine core and define a bypass exhaust nozzle, the bypass exhaust nozzle having an outer radius, and
an outer bypass to fan ratio of:
the
outer
radius
of
the
bypass
exhaust
nozzle
the
fan
tip
radius
is in the range from 0.6 to 1.05.
9 . The gas turbine engine of claim 8 , wherein the outer bypass to fan ratio is any one or more of:
i. in the range from 0.60 to 1.05; ii. in the range from 0.65 to 1.00; iii. lower than 1.05, optionally lower than 1.02, and further optionally lower than 1.00; and/or iv. in the range from 0.80 to 1.00.
10 . The gas turbine engine of claim 8 , wherein the outer bypass to fan ratio is in the range from 0.9 to 1.0, and optionally in the range from 0.90 to 1.00, and further optionally:
(i) the gas turbine engine has a fan tip radius in the range from 110 cm to 150 cm and the outer bypass to fan ratio is in the range from 0.95 to 1.00; or (ii) the gas turbine engine has a fan tip radius in the range from 155 cm to 200 cm and the outer bypass to fan ratio is equal to or around 0.95, for example being in the range from 0.91 to 0.98.
11 . The gas turbine engine of any of claim 8 , wherein the bypass exhaust nozzle has an exit plane, and the outer radius of the bypass exhaust nozzle is measured at the axial position of the exit plane of the bypass exhaust nozzle.
12 . The gas turbine engine of any of claim 8 , wherein the outer radius of the bypass exhaust nozzle is measured at the axial position of the rearmost tip of the nacelle.
13 . The gas turbine engine of any of claim 8 , wherein the outer radius of the bypass exhaust nozzle is the radial distance between the centreline of the engine and an inner surface of the nacelle at the axial position of the rearmost tip of the nacelle.
14 . The gas turbine engine of any of claim 8 , wherein the outer radius of the bypass exhaust nozzle is in the range of 100 cm to 200 cm, and optionally 100 cm to 190 cm, and further optionally:
(i) the fan tip radius is in the range from 95 cm to 150 cm and the outer radius of the bypass exhaust nozzle is in the range from 100 cm to 145 cm; or (ii) the fan tip radius is in the range from 155 cm to 200 cm and the outer radius of the bypass exhaust nozzle is in the range from 145 cm to 190 cm.
15 . The gas turbine engine of claim 8 , wherein the bypass exhaust nozzle has an inner radius, and an inner bypass to fan ratio of:
the
inner
radius
of
the
bypass
exhaust
nozzle
the
fan
tip
radius
is in the range from 0.4 to 0.65.
16 . The gas turbine engine of claim 15 , wherein the inner bypass to fan ratio is any one or more of:
i. in the range from 0.5 to 0.6, and optionally in the range from 0.50 to 0.60; ii. in the range from 0.40 to 0.65; iii. lower than 0.65, and optionally lower than 0.64, and optionally lower than 0.62; and/or iv. in the range from 0.54 to 0.64.
17 . The gas turbine engine of claim 15 , wherein:
(i) the gas turbine engine has a fan tip radius in the range from 110 cm to 150 cm and the inner bypass to fan ratio is in the range from 0.57 to 0.63, for example being in the range from 0.58 to 0.60; or (ii) the gas turbine engine has a fan tip radius in the range from 155 cm to 200 cm and the inner bypass to fan ratio is in the range from 0.5 to 0.6, and optionally from 0.52 to 0.58.
18 . The gas turbine engine of any of claim 15 , wherein the bypass exhaust nozzle has an exit plane and the inner radius of the bypass exhaust nozzle is measured at the axial position of the exit plane of the bypass exhaust nozzle.
19 . The gas turbine engine of any of claim 15 , wherein the inner radius of the bypass exhaust nozzle is:
(i) measured at the axial position of the rearmost tip of the nacelle; and/or (ii) the radial distance between the centreline of the engine and an outer surface of the engine core at the axial position of the rearmost tip of the nacelle.
20 . The gas turbine engine of claim 15 , wherein the inner radius of the bypass exhaust nozzle is in the range from 50 cm to 125 cm, and optionally from 65 cm to 110 cm, and further optionally:
(i) the gas turbine engine has a fan tip radius in the range from 110 cm to 150 cm and the inner radius of the bypass exhaust nozzle is in the range from 65 cm to 90 cm; or (ii) the gas turbine engine has a fan tip radius in the range from 155 cm to 200 cm and the inner radius of the bypass exhaust nozzle is in the range from 80 cm to 110 cm.Join the waitlist — get patent alerts
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