Gas turbine engine
A gas turbine engine ( 10 ) for an aircraft comprises an engine core ( 11 ) comprising: a compressor system comprising a first, lower pressure, compressor ( 14 ), and a second, higher pressure, compressor ( 15 ); and an outer core casing ( 70 ) surrounding the compressor system. The gas turbine engine further comprises a fan ( 23 ) located upstream of the engine core ( 11 ), the fan comprising a plurality of fan blades. The outer core casing comprises: a first flange connection ( 60 ) arranged to allow separation of the outer core casing ( 70 ) at an axial position of the first flange connection ( 60 ), the first flange connection ( 60 ) having a first flange radius ( 104 ), wherein the first flange connection ( 60 ) is the first flange connection that is downstream of an axial position defined by the axial midpoint between the mid-span axial location on the trailing edge of the most downstream aerofoil of the first compressor ( 14 ) and the mid-span axial location on the leading edge of the most upstream aerofoil of the second compressor ( 15 ). A fan blade mass ratio of: first flange radius ( 104 ) mass of each fan blade is equal to or less than 19.0 mm/lb.
1. A gas turbine engine for an aircraft comprising:
an engine core comprising:
a compressor system with compressor blades comprising respective aerofoils, the compressor system comprising a first, lower pressure, compressor, and a second, higher pressure, compressor;
an outer core casing surrounding the compressor system and comprising:
a first flange connection arranged to allow separation of the outer core casing at an axial position of the first flange connection, the first flange connection having a first flange radius, the first flange connection being the first flange connection that is downstream of an axial position defined by an axial midpoint between a mid-span axial location on a trailing edge of a most downstream aerofoil of the first compressor and a mid-span axial location on a leading edge of a most upstream aerofoil of the second compressor; and
an inner core casing provided radially inwardly of the compressor blades of the compressor system, the inner core casing and the outer core casing defining a core working gas flow path (A) therebetween, a gas path radius being defined as an outer radius of the gas flow path (A) at the axial position of the first flange connection, and a gas path ratio of:
first
flange
radius
g
a
s
path
radius
being equal to or greater than 1.10 and less than or equal to 2.0; and
a fan located upstream of the engine core, the fan comprising a plurality of fan blades, a fan blade mass ratio of:
first
flange
radius
mass
of
each
fan
blade
being equal to or less than 19.0 mm/lb.
2. The gas turbine engine of claim 1 , wherein the fan blade mass ratio is equal to or greater than 5 mm/lb.
3. The gas turbine engine of claim 1 , wherein a blade set mass ratio of:
first
flange
radius
total
mass
of
the
plurality
of
fan
blades
is in a range between 0.95 mm/lb and 0.35 mm/lb.
4. The gas turbine engine of claim 1 , wherein each of the fan blades are at least partly formed from a metallic material, wherein optionally the metallic material is titanium or aluminium lithium alloy.
5. The gas turbine engine of claim 1 , wherein each of the fan blades are formed at least partly from a composite material.
6. The gas turbine engine of claim 1 , wherein the fan diameter is greater than 240 cm and less than or equal to 380 cm.
7. The gas turbine engine of claim 1 , wherein the fan diameter is between 330 cm and 380 cm.
8. The gas turbine engine of claim 1 , wherein one or both of the following is satisfied:
a) the number of fan blades is between 16 and 22; or
b) the mass of each fan blade is in a range between 20 lb and 70 lb.
9. The gas turbine engine of claim 1 , wherein the engine further comprises a gearbox that receives an input from a core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft, and optionally wherein a gear ratio of the gearbox is between 3.1 and 4.0.
10. The gas turbine engine of claim 1 , wherein the first flange connection is at, or axially downstream of, the leading edge of the most upstream aerofoil of the second compressor.
11. The gas turbine engine of claim 1 , wherein the first flange connection is at, or axially upstream of, the leading edge of the most upstream aerofoil of the second compressor.
12. The gas turbine engine of claim 1 , further comprising:
a first turbine, and a first core shaft connecting the first turbine to the first compressor; and
a second turbine and a second core shaft connecting the second turbine to the second compressor,
wherein the second turbine, second compressor, and second core shaft are arranged to rotate at a higher rotational speed than the first core shaft.
13. The gas turbine engine of claim 1 , wherein a fan diameter ratio of:
first
flange
radius
f
an
diameter
is equal to or greater than 0.125 and less than or equal to 0.17.
14. The gas turbine engine of claim 1 , further comprising a front mount arranged to be connected to a pylon.
15. The gas turbine engine of claim 14 , wherein a front mount position to fan diameter ratio of:
the
axial
distance
between
the
first
flange
connection
and
the
front
mount
the
fan
diameter
is less than or equal to 0.145.
16. The gas turbine engine of claim 14 , wherein the front mount is a core mount.
17. The gas turbine engine of claim 1 , further comprising:
a nacelle surrounding the engine core and defining a bypass duct between the engine core and the nacelle; and
a fan outlet guide vane (OGV) extending radially across the bypass duct between an outer surface of the engine core and an inner surface of the nacelle, the fan OGV having a radially inner edge and a radially outer edge, wherein an axial midpoint of the radially inner edge is defined as a fan OGV root centrepoint.
18. The gas turbine engine according to claim 1 , wherein the fan blade mass ratio is equal to or less than 18.0 mm/lb.
19. The gas turbine engine according to claim 1 , wherein the fan blade mass ratio is equal to or less than 17.0 mm/lb.
20. The gas turbine engine according to claim 1 , wherein the fan blade mass ratio is equal to or less than 16.0 mm/lb.
21. The gas turbine engine according to claim 1 , wherein the fan blade mass ratio is equal to or less than 15.0 mm/lb.
22. A gas turbine engine for an aircraft comprising:
an engine core comprising:
a compressor system with compressor blades comprising respective aerofoils, the compressor system comprising a first, lower pressure, compressor, and a second, higher pressure, compressor; and
an outer core casing surrounding the compressor system and comprising:
a first flange connection arranged to allow separation of the outer core casing at an axial position of the first flange connection, the first flange connection having a first flange radius, the first flange connection being the first flange connection that is downstream of an axial position defined by an axial midpoint between a mid-span axial location on a trailing edge of a most downstream aerofoil of the first compressor and a mid-span axial location on a leading edge of a most upstream aerofoil of the second compressor;
a fan located upstream of the engine core, the fan comprising a plurality of fan blades, a fan blade mass ratio of:
first
flange
radius
mass
of
each
fan
blade
being equal to or less than 19.0 mm/lb; and
a front mount arranged to be connected to a pylon;
wherein a front mount position ratio of:
axial
distance
between
the
first
flange
connection
and
the
front
mount
first
flange
rad
ius
is equal to or less than 1.18.
23. A gas turbine engine for an aircraft comprising:
an engine core comprising:
a compressor system with compressor blades comprising respective aerofoils, the compressor system comprising a first, lower pressure, compressor, and a second, higher pressure, compressor; and
an outer core casing surrounding the compressor system and comprising:
a first flange connection arranged to allow separation of the outer core casing at an axial position of the first flange connection, the first flange connection having a first flange radius, the first flange connection being the first flange connection that is downstream of an axial position defined by an axial midpoint between a mid-span axial location on a trailing edge of a most downstream aerofoil of the first compressor and a mid-span axial location on a leading edge of a most upstream aerofoil of the second compressor;
a fan located upstream of the engine core, the fan comprising a plurality of fan blades, a fan blade mass ratio of:
first
flange
radius
mass
of
each
fan
blade
being equal to or less than 19.0 mm/lb;
a nacelle surrounding the engine core and defining a bypass duct between the engine core and the nacelle; and
a fan outlet guide vane (OGV) extending radially across the bypass duct between an outer surface of the engine core and an inner surface of the nacelle, the fan OGV having a radially inner edge and a radially outer edge, an axial midpoint of the radially inner edge being defined as a fan OGV root centrepoint;
wherein any one or more of the following is satisfied:
(i) a fan OGV root position ratio of:
axial
distance
between
the
first
flange
connection
and
the
fan
O
G
V
root
centrepoint
first
flange
radius
is equal to or less than 2.6,
(ii) a fan OGV root position to fan diameter ratio of:
the
axial
distance
between
the
first
flange
connection
and
the
fan
O
G
V
r
o
ot
centrepoint
the
fan
diameter
is less than or equal to 0.33,
(iii) a fan OGV tip position ratio of:
axial
distance
between
the
first
flange
connection
and
a
fan
O
G
V
tip
centrepoint
first
flange
radius
is equal to or less than 1.8, and
(iv) a fan OGV tip position to fan diameter ratio of:
the
axial
distance
between
the
first
flange
connection
and
the
fan
O
G
V
tip
centrepoint
the
fan
diameter
is less than or equal to 0.22.