U.S. patent application number 16/718405 was filed with the patent office on 2020-07-23 for highly loaded inlet duct in a geared turbofan.
This patent application is currently assigned to ROLLS-ROYCE plc. The applicant listed for this patent is ROLLS-ROYCE plc. Invention is credited to Ian J. BOUSFIELD.
Application Number | 20200232392 16/718405 |
Document ID | / |
Family ID | 65528267 |
Filed Date | 2020-07-23 |
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United States Patent
Application |
20200232392 |
Kind Code |
A1 |
BOUSFIELD; Ian J. |
July 23, 2020 |
HIGHLY LOADED INLET DUCT IN A GEARED TURBOFAN
Abstract
A gas turbine engine includes an inlet duct to guide a core
engine flow to a compressor and an engine section stator arranged
in the inlet duct upstream of the compressor including vanes with
leading edges defining a first annulus area in the inlet duct, a
mid-span leading edge point of the engine section stator vanes
being arranged at a first radius. The compressor includes a first
rotor with a row of first blades with leading edges defining a
second annulus area; a mid-span leading edge point of the
compressor first blades being arranged at a second radius and an
axial distance from the engine section stator vane mid-span leading
edge point. The ratio of the second to the first annulus area is
equal or greater than 0.75, and the ratio of a difference between
the first and second radius to the axial distance is equal or
greater than 0.23.
Inventors: |
BOUSFIELD; Ian J.;
(Nottingham, GB) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
ROLLS-ROYCE plc |
London |
|
GB |
|
|
Assignee: |
ROLLS-ROYCE plc
London
GB
|
Family ID: |
65528267 |
Appl. No.: |
16/718405 |
Filed: |
December 18, 2019 |
Current U.S.
Class: |
1/1 |
Current CPC
Class: |
F02C 7/057 20130101;
F05D 2260/40311 20130101; F05D 2270/021 20130101; F02K 3/06
20130101; F02C 7/36 20130101; F02C 3/06 20130101; F05D 2220/327
20130101 |
International
Class: |
F02C 7/36 20060101
F02C007/36; F02C 7/057 20060101 F02C007/057; F02C 3/06 20060101
F02C003/06; F02K 3/06 20060101 F02K003/06 |
Foreign Application Data
Date |
Code |
Application Number |
Jan 18, 2019 |
GB |
1900733.5 |
Claims
1. A gas turbine engine comprising: a fan mounted to rotate about a
main longitudinal axis to produce a bypass flow and a core engine
flow; an engine core, comprising in axial flow series a compressor,
a combustor, and a turbine coupled to the compressor through a
shaft; an inlet duct to guide the core engine flow to the
compressor; a reduction gearbox that receives input from the shaft
and outputs drive to the fan so as to drive the fan at a lower
rotational speed than the shaft; an engine section stator arranged
in the inlet duct upstream of the compressor, the engine section
stator comprising a plurality of vanes with respective leading
edges defining a first annulus area A1 in the inlet duct, mid-span
leading edge points of the engine section stator vanes being
arranged at a first radius R1 from the main longitudinal axis;
wherein the compressor comprises a first rotor with a row of first
blades with respective leading edges defining a second annulus area
A2 in the inlet duct, mid-span leading edge points of the
compressor first blades being arranged at a second radius R2 from
the main longitudinal axis and at an axial distance L from the
engine section stator vane mid-span leading edge points, and
wherein a ratio of the second annulus area A2 to the first annulus
area A1 is equal or greater than 0.80, and a ratio .DELTA.R/L of a
difference .DELTA.R between the first radius and the second radius
to the axial distance L is equal or greater than 0.23.
2. The gas turbine engine according claim 1, wherein the ratio
.DELTA.R/L of the difference .DELTA.R between the first radius and
the second radius to the axial distance L is equal or greater than
0.25.
3. The gas turbine engine according to claim 1, wherein the ratio
.DELTA.R/L of the difference between the first radius and the
second radius to the axial distance L is equal or less than
0.70.
4. The gas turbine engine according to claim 1, wherein the ratio
.DELTA.R/L of the difference between the first radius and the
second radius to the axial distance L is equal or less than
0.68.
5. The gas turbine engine according to claim 1, wherein the ratio
of the second annulus area A2 to the first annulus area A1 is equal
or less than 1.10.
6. The gas turbine engine according to claim 1, wherein the ratio
of the second annulus area A2 to the first annulus area A1 is equal
or less than 1.05.
7. The gas turbine engine according to claim 1, comprising a
variable inlet guide vane arranged in the inlet duct between the
engine section stator and the first rotor of the compressor.
8. The gas turbine engine according to claim 7, comprising a strut
arranged in the inlet duct between the engine section stator and
the first rotor of the compressor.
9. The gas turbine engine according to claim 8, wherein the strut
is arranged between the engine section stator and the variable
inlet guide vane.
10. The gas turbine engine according to claim 1, wherein the fan
has a diameter greater than 230 cm.
11. The gas turbine engine according to claim 1, wherein the
reduction gearbox has a gear ratio comprised between 2.5 and
4.2.
12. The gas turbine engine according to claim 1, wherein the
compressor comprises 2 to 8 stages.
13. The gas turbine engine according to claim 1, wherein a bypass
ratio, defined as a mass ratio between the bypass flow and the core
engine flow, is higher than 10.
14. The gas turbine engine according to claim 1, wherein: the
compressor is an intermediate pressure compressor, the gas turbine
engine further comprising a high pressure compressor downstream of
the intermediate pressure compressor; the turbine is an
intermediate pressure turbine, the gas turbine engine further
comprising a high pressure turbine upstream of the intermediate
pressure turbine; and the shaft is a first shaft, the gas turbine
engine further comprising a second shaft coupling the high pressure
turbine to the high pressure compressor.
15. A gas turbine engine comprising: a fan mounted to rotate about
a main longitudinal axis to produce a bypass flow and a core engine
flow; an engine core, comprising in axial flow series a compressor,
a combustor, and a turbine coupled to the compressor through a
shaft; an inlet duct to guide the core engine flow to the
compressor; a reduction gearbox that receives input from the shaft
and outputs drive to the fan so as to drive the fan at a lower
rotational speed than the shaft, the reduction gearbox having a
gear ratio comprised between 3.1 and 3.7; an engine section stator
arranged in the inlet duct upstream of the compressor, the engine
section stator comprising a plurality of vanes with respective
leading edges, mid-span leading edge points of the engine section
stator vanes being arranged at a first radius R1 from the main
longitudinal axis; wherein the compressor comprises a first rotor
with a row of first blades with respective leading edges, mid-span
leading edge points of the compressor first blades being arranged
at a second radius R2 from the main longitudinal axis and at an
axial distance L from the engine section stator vane mid-span
leading edge points, and wherein a ratio .DELTA.R/L of a difference
.DELTA.R between the first radius and the second radius to the
axial distance L is comprised between 0.40 and 0.70.
16. The gas turbine engine according to claim 15, wherein the
compressor comprises 2 to 8 stages.
17. A gas turbine engine comprising: a fan mounted to rotate about
a main longitudinal axis to produce a bypass flow and a core engine
flow; an engine core, comprising in axial flow series a compressor,
a combustor, and a turbine coupled to the compressor through a
shaft; an inlet duct to guide the core engine flow to the
compressor; a reduction gearbox that receives input from the shaft
and outputs drive to the fan so as to drive the fan at a lower
rotational speed than the shaft; an engine section stator arranged
in the inlet duct upstream of the compressor, the engine section
stator comprising a plurality of vanes with respective leading
edges defining a first annulus area A1 in the inlet duct, mid-span
leading edge points of the engine section stator vanes being
arranged at a first radius R1 from the main longitudinal axis;
wherein the compressor comprises a first rotor with a row of first
blades with respective leading edges defining a second annulus area
A2 in the inlet duct, mid-span leading edge points of the
compressor first blades being arranged at a second radius R2 from
the main longitudinal axis and at an axial distance L from the
engine section stator vane mid-span leading edge points, wherein a
ratio of the second annulus area A2 to the first annulus area A1 is
equal or greater than 0.60, and a ratio .DELTA.R/L of a difference
.DELTA.R between the first radius and the second radius to the
axial distance L is equal or greater than 0.23, and wherein the
ratio .DELTA.R/L of the difference between the first radius and the
second radius to the axial distance L is equal or less than
0.70.
18. The gas turbine engine according to claim 17, wherein the ratio
.DELTA.R/L of the difference between the first radius and the
second radius to the axial distance L is equal or less than 0.68.
Description
[0001] The present disclosure relates generally to gas turbine
engines, and more specifically to geared gas turbine engines.
[0002] Gas turbine engines are used to power aircraft, watercraft,
power generators, and the like. A gas turbine engine generally
comprises, in axial flow series, an air intake, a fan rotating
about a principal rotational axis, one or more compressors, a
combustor, one or more turbines, and an exhaust nozzle. Air
entering the air intake is accelerated by the fan to produce two
air flows: a first air flow (core flow) into a core and a second
air flow (bypass flow) which passes through a bypass duct to
provide propulsive thrust.
[0003] Air entering the core through an inlet duct is guided by an
engine section stator (ESS) or core inlet stator towards the
compressor where it is compressed, mixed with fuel and then fed
into the combustor, where combustion of the air/fuel mixture
occurs. The high temperature and high energy exhaust fluids are
then fed to the turbine, where the energy of the fluids is
converted to mechanical energy to drive the compressor in rotation
by suitable interconnecting shaft.
[0004] The mass ratio between the bypass flow and the core flow is
known as the bypass ratio. At subsonic aircraft velocities, a large
bypass ratio is desirable for high propulsive efficiency.
[0005] Recently, geared turbofans have been proposed. In a geared
turbofan, a compressor is coupled to the fan via a reduction
gearbox, which allows the fan to rotate slower than the compressor.
As the fan rotates slower, the fan diameter can be increased and so
the bypass ratio.
[0006] There are numerous competing requirements and parameters for
aircraft gas turbine engines. They must have low fuel burn, whilst
also being lightweight and have small dimensions.
[0007] This de-coupling of the fan speed from the compressor speed
opens the design space for the fan and the core considerably.
Consequently, the dimensions of conventional gas turbine engine
cores and fans may be non-optimal where a reduction gearbox is
introduced.
[0008] For example, as the compressor rotates faster, the blade
speed may exceed an optimal value leading to high Mach numbers,
poor efficiency and greater noise. Consequently, the compressor is
generally moved radially inward to reduce the blade speed.
[0009] Moreover, limited radius change is possible between the fan
trailing edge and the ESS. Because the EES radial position in
mainly determined by the fan size, as the fan diameter increases,
the ESS is moved radially outwardly to reduce the radius change
between the fan trailing edge and the ESS.
[0010] In other words, the difference in radius between the
compressor and the ESS increases in geared turbofan. To cope with
such increased difference in radius the inlet duct is made longer
and this results in a heavier, less efficient and more expensive
engine.
[0011] There is therefore a need for a geared gas turbine engine
with an improved inlet duct such as to overcome, at least in part,
the above mentioned disadvantages.
[0012] According to a first aspect, there is provided a gas turbine
engine comprising: a fan mounted to rotate about a main
longitudinal axis to produce a bypass flow and a core flow; an
engine core, comprising in axial flow series a compressor, a
combustor, and a turbine coupled to the compressor through a shaft;
an inlet duct to guide the core flow to the compressor; a reduction
gearbox that receives an input from the shaft and outputs drive to
the fan so as to drive the fan at a lower rotational speed than the
shaft; an engine section stator arranged in the inlet duct upstream
of the compressor, the engine section stator comprising a plurality
of vanes with respective leading edges defining a first annulus
area A1 in the inlet duct, a mid-span leading edge point of the
engine section stator vanes being arranged at a first radius R1
from the main longitudinal axis; wherein the compressor comprises a
first rotor with a row of first blades with respective leading
edges defining a second annulus area A2 in the inlet duct; a
mid-span leading edge point of the compressor first blades being
arranged at a second radius R2 from the main longitudinal axis and
at an axial distance L from the engine section stator vane mid-span
leading edge point; wherein a ratio A2/A1 of the second annulus
area A2 to the first annulus area A1 is equal or greater than 0.60,
and a ratio .DELTA.R/L of a difference .DELTA.R between the first
radius R1 and the second radius R2 to the axial distance L is equal
or greater than 0.23.
[0013] The blades may comprise an aerofoil portion with a tip, and
a root, and a spanwise direction is a direction extending between
the tip and the root of the blades, and a chordwise direction is a
direction extending between the leading edge and the trailing edge
of the blades.
[0014] In the present disclosure, upstream and downstream are with
respect to the core flow through the compressor; and front and rear
is with respect to the gas turbine engine, i.e. the fan being in
the front and the turbine being in the rear of the engine.
[0015] The present inventor has found a unique parameter
combination for the inlet duct that allows optimizing efficiency,
weight and dimensions of the gas turbine engine, without negatively
affecting the core flow in the inlet duct in terms of flow
properties.
[0016] The ratio .DELTA.R/L describes a mean slope of the inlet
duct. In the present disclosure, such ratio is relatively high,
enabling a shorter duct length, and therefore a shorter and lighter
engine, while maintain an optimum compressor blade speed, gear
ratio, and bypass ratio. In other words, the fan size can be
optimised for maximum thrust and efficiency, and the compressor can
be arranged distanced to the main longitudinal axis to optimise the
compressor blade speed.
[0017] The ratio A2/A1 of the second annulus area A2 to the first
annulus area A1 describes the mean axial diffusion along the inlet
duct and a ratio less than 1 implies an accelerating duct. It is
usually the static pressure gradient along an inner annulus of the
inlet duct which determines how difficult it is for the core flow
to flow in the inlet duct. The static pressure gradient increases
with the mean axial diffusion, so the ratio A2/A1 of the second
annulus area A2 to the first annulus area A1 is an indication of
such difficulty. Also the ratio .DELTA.R/L has an influence on the
static pressure gradient along the inlet duct: the higher the ratio
.DELTA.R/L, the higher the static pressure gradient. The unique
combination of the ratio A2/A1 and the ratio .DELTA.R/L of the
present disclosure enables to keep the static pressure gradient at
an acceptable level while optimizing efficiency, weight and
dimensions of the gas turbine engine as explained above.
[0018] In some examples, the ratio .DELTA.R/L of the difference
between the first radius and the second radius to the axial
distance L may be equal or greater than 0.25, or equal or greater
than 0.27, or equal or greater than 0.30, or equal or greater than
0.31, or equal or greater than 0.32, or equal or greater than 0.33,
or equal or greater than 0.34, or equal or greater than 0.35, or
equal or greater than 0.35, or equal or greater than 0.36, or equal
or greater than 0.37, or equal or greater than 0.38, or equal or
greater than 0.39, or equal or greater than 0.40, or equal or
greater than 0.42, or equal or greater than 0.44, or equal or
greater than 0.46, or equal or greater than 0.48, or equal or
greater than 0.50. In some examples, the ratio .DELTA.R/L of the
difference between the first radius and the second radius to the
axial distance L may be equal or less than 0.70, or equal or less
than 0.68, or equal or less than 0.66, or equal or less than 0.64,
or equal or less than 0.62, or equal or less than 0.6, or equal or
less than 0.58, or equal or less than 0.56, or equal or less than
0.54, or equal or less than 0.52, or equal or less than 0.50. For
example, the ratio .DELTA.R/L may be comprised between 0.30 and
0.70, or between 0.30 and 0.50 or between 0.35 and 0.70, or between
0.40 and 0.70, or between 0.35 and 0.65, or between 0.35 and 0.60,
or between 0.35 and 0.55, or between 0.40 and 0.60, or between 0.40
and 0.55. By limiting the ratio .DELTA.R/L to values of, or less
than, 0.70 it may be possible to reduce the risk of flow separation
within the inlet duct due to large adverse pressure gradients.
[0019] The ratio A2/A1 of the first annulus area A1 to the second
annulus area A2 may be equal or greater than 0.65, or equal or
greater than 0.70, or equal or greater than 0.75, or equal or
greater than 0.80, or equal or greater than 0.85, or equal or
greater than 0.90, or equal or greater than 0.95. Moreover, the
ratio A2/A1 of the first annulus area A1 to the second annulus area
A2 may be equal or less than 1.10, or equal or less than 1.05, or
equal or less than 1, or equal or less than 0.95, or equal or less
than 0.90. For example, the ratio A2/A1 may be comprised between
0.60 and 1.10, or between 0.60 and 1.05, or between 0.60 and 1, or
between 0.60 and 0.90, or between 0.60 and 0.95, or between 0.60
and 0.90, or between 0.65 and 1.10, or between 0.65 and 1, or
between 0.65 and 0.90, or between 0.65 and 0.85, or between 0.70
and 1.10, or between 0.70 and 1, or between 0.70 and 0.90, or
between 0.75 and 1.10, or between 0.75 and 1, or between 0.75 and
0.95, or between 0.80 and 1.10, or between 0.80 and 1.
[0020] The gas turbine engine may comprise a variable inlet guide
vane arranged in the inlet duct between the engine section stator
and the first rotor of the compressor.
[0021] The gas turbine engine may comprise a strut arranged in the
inlet duct between the engine section stator and the first rotor of
the compressor.
[0022] The strut may be arranged between the engine section stator
and the variable inlet guide vane. In other words, the strut may be
arranged downstream of the engine section stator and upstream of
the variable inlet guide vane and the first rotor.
[0023] The present disclosure may relate to relatively large gas
turbine engines, i.e. to gas turbine engines with relatively large
fan diameter. For example, the fan may have a diameter greater than
230 cm, or greater than 235 cm, or greater than 240 cm, or greater
than 245 cm, or greater than 250 cm.
[0024] The reduction gearbox may have a gear ratio comprised
between 2.5 and 4.2, for example between 3.0 and 4.0, or between
3.1 and 3.7, or between 3.5 and 3.8.
[0025] The compressor may comprise a plurality of stages, each
stage comprising a rotor and a stator. The compressor may comprise
two or more stages, for example three or four stages. The
compressor may comprise less than ten stages, or less than nine
stages. For example, the compressor may comprise 2 to 8 stages.
[0026] The compressor may be an intermediate pressure compressor
and the gas turbine engine may comprise a high pressure compressor
downstream of the intermediate pressure compressor.
[0027] Moreover, the turbine may be an intermediate pressure
turbine and the gas turbine engine may comprise a high pressure
turbine upstream of the intermediate pressure compressor.
[0028] The shaft may be a first shaft and the gas turbine engine
may further comprise a second shaft coupling the high pressure
turbine to the high pressure compressor. The first shaft may couple
the intermediate pressure compressor and the intermediate pressure
turbine.
[0029] The gas turbine engine may have a bypass ratio, defined as a
mass ratio between the bypass flow and the core engine flow, higher
than 10.
[0030] According to a second aspect, there is provided a gas
turbine engine comprising a fan mounted to rotate about a main
longitudinal axis to produce a bypass flow and a core engine flow;
an engine core, comprising in axial flow series a compressor, a
combustor, and a turbine coupled to the compressor through a shaft;
an inlet duct to guide the core engine flow to the compressor; a
reduction gearbox that receives an input from the shaft and outputs
drive to the fan so as to drive the fan at a lower rotational speed
than the shaft, the reduction gearbox having a gear ratio comprised
between 3.1 and 3.7; and an engine section stator arranged in the
inlet duct upstream of the compressor, the engine section stator
comprising a plurality of vanes with respective leading edges, a
mid-span leading edge point of the engine section stator vanes
being arranged at a first radius R1 from the main longitudinal
axis. The compressor comprises a first rotor with a row of first
blades with respective leading edges; a mid-span leading edge point
of the compressor first blades being arranged at a second radius R2
from the main longitudinal axis and at an axial distance L from the
engine section stator vane mid-span leading edge point. A ratio
.DELTA.R/L of a difference .DELTA.R between the first radius and
the second radius to the axial distance L is comprised between 0.23
and 0.70, optionally between 0.25 and 0.70.
[0031] As said before, arrangements of the present disclosure may
be particularly beneficial for fans that are driven via a gearbox.
Accordingly, the gas turbine engine may comprise 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. The input to the gearbox may be directly from the core
shaft, or indirectly from the core shaft, for example via a spur
shaft and/or gear. The core shaft may rigidly connect the turbine
and the compressor, such that the turbine and compressor rotate at
the same speed (with the fan rotating at a lower speed).
[0032] The gas turbine engine as described and/or claimed herein
may have any suitable general architecture. For example, the gas
turbine engine may have any desired number of shafts that connect
turbines and compressors, for example one, two or three shafts.
Purely by way of example, the turbine connected to the core shaft
may be a first turbine, the compressor connected to the core shaft
may be a first compressor, and the core shaft may be a first core
shaft. The engine core may further comprise a second turbine, a
second compressor, and a second core shaft connecting the second
turbine to the second compressor. The second turbine, second
compressor, and second core shaft may be arranged to rotate at a
higher rotational speed than the first core shaft.
[0033] In such an arrangement, the second compressor may be
positioned axially downstream of the first compressor. The second
compressor may be arranged to receive (for example directly
receive, for example via a generally annular duct) flow from the
first compressor.
[0034] The gearbox may be arranged to be driven by the core shaft
that is configured to rotate (for example in use) at the lowest
rotational speed (for example the first core shaft in the example
above). For example, the gearbox may be arranged to be driven only
by the core shaft that is configured to rotate (for example in use)
at the lowest rotational speed (for example only be the first core
shaft, and not the second core shaft, in the example above).
Alternatively, the gearbox may be arranged to be driven by any one
or more shafts, for example the first and/or second shafts in the
example above.
[0035] The gearbox is a reduction gearbox (in that the output to
the fan is a lower rotational rate than the input from the core
shaft). Any type of gearbox may be used. For example, the gearbox
may be a "planetary" or "star" gearbox, as described in more detail
elsewhere herein. The gearbox may have any desired reduction ratio
(defined as the rotational speed of the input shaft divided by the
rotational speed of the output shaft), for example greater than
2.5, for example in the range of from 3 to 4.2, for example on the
order of or at least 3, 3.1, 3.2, 3.3, 3.4, 3.5, 3.6, 3.7, 3.8,
3.9, 4, 4.1 or 4.2. The gear ratio may be, for example, between any
two of the values in the previous sentence. A higher gear ratio may
be more suited to "planetary" style gearbox. In some arrangements,
the gear ratio may be outside these ranges.
[0036] In any gas turbine engine as described and/or claimed
herein, a combustor may be provided axially downstream of the fan
and compressor(s). For example, the combustor may be directly
downstream of (for example at the exit of) the second compressor,
where a second compressor is provided. By way of further example,
the flow at the exit to the combustor may be provided to the inlet
of the second turbine, where a second turbine is provided. The
combustor may be provided upstream of the turbine(s).
[0037] The or each compressor (for example the first compressor and
second compressor as described above) may comprise any number of
stages, for example multiple stages. Each stage may comprise a row
of rotor blades and a row of stator vanes, which may be variable
stator vanes (in that their angle of incidence may be variable).
The row of rotor blades and the row of stator vanes may be axially
offset from each other.
[0038] The or each turbine (for example the first turbine and
second turbine as described above) may comprise any number of
stages, for example multiple stages. Each stage may comprise a row
of rotor blades and a row of stator vanes. The row of rotor blades
and the row of stator vanes may be axially offset from each
other.
[0039] Each fan blade may be defined as having a radial span
extending from a root (or hub) at a radially inner gas-washed
location, or 0% span position, to a tip at a 100% span position.
The ratio of the radius of the fan blade at the hub to the radius
of the fan blade at the tip may be less than (or on the order of)
any of: 0.4, 0.39, 0.38, 0.37, 0.36, 0.35, 0.34, 0.33, 0.32, 0.31,
0.3, 0.29, 0.28, 0.27, 0.26, or 0.25. The ratio of the radius of
the fan blade at the hub to the radius of the fan blade at the tip
may be in an inclusive range bounded by any two of the values in
the previous sentence (i.e. the values may form upper or lower
bounds). These ratios may commonly be referred to as the hub-to-tip
ratio. The radius at the hub and the radius at the tip may both be
measured at the leading edge (or axially forwardmost) part of the
blade. The hub-to-tip ratio refers, of course, to the gas-washed
portion of the fan blade, i.e. the portion radially outside any
platform.
[0040] The radius of the fan may be measured between the engine
centreline and the tip of a fan blade at its leading edge. The fan
diameter (which may simply be twice the radius of the fan) may be
greater than (or on the order of) any of: 220 cm, 230 cm (around
90.5 inches), 235 cm (around 92.5 inches), 240 cm, 250 cm (around
100 inches), 260 cm, 270 cm (around 105 inches), 280 cm (around 110
inches), 290 cm (around 115 inches), 300 cm (around 120 inches),
310 cm, 320 cm (around 125 inches), 330 cm (around 130 inches), 340
cm (around 135 inches), 350 cm, 360 cm (around 140 inches), 370 cm
(around 145 inches), 380 (around 150 inches) cm, 390 cm (around 155
inches), 400 cm, 410 cm (around 160 inches) or 420 cm (around 165
inches). The fan diameter may be in an inclusive range bounded by
any two of the values in the previous sentence (i.e. the values may
form upper or lower bounds).
[0041] The rotational speed of the fan may vary in use. Generally,
the rotational speed is lower for fans with a higher diameter.
Purely by way of non-limitative example, the rotational speed of
the fan at cruise conditions may be less than 2500 rpm, for example
less than 2300 rpm. Purely by way of further non-limitative
example, the rotational speed of the fan at cruise conditions for
an engine having a fan diameter in the range of from 220 cm to 300
cm (for example 230 cm to 270 cm) may be in the range of from 1700
rpm to 2500 rpm, for example in the range of from 1800 rpm to 2300
rpm, for example in the range of from 1900 rpm to 2100 rpm. Purely
by way of further non-limitative example, the rotational speed of
the fan at cruise conditions for an engine having a fan diameter in
the range of from 320 cm to 380 cm may be in the range of from 1200
rpm to 2000 rpm, for example in the range of from 1300 rpm to 1800
rpm, for example in the range of from 1400 rpm to 1600 rpm.
[0042] In use of the gas turbine engine, the fan (with associated
fan blades) rotates about a rotational axis. This rotation results
in the tip of the fan blade moving with a velocity U.sub.tip. The
work done by the fan blades on the flow results in an enthalpy rise
dH of the flow. A fan tip loading may be defined as
dH/U.sub.tip.sup.2, where dH is the enthalpy rise (for example the
1-D average enthalpy rise) across the fan and U.sub.tip is the
(translational) velocity of the fan tip, for example at the leading
edge of the tip (which may be defined as fan tip radius at leading
edge multiplied by angular speed). The fan tip loading at cruise
conditions may be greater than (or on the order of) any of: 0.28,
0.29, 0.3, 0.31, 0.32, 0.33, 0.34, 0.35, 0.36, 0.37, 0.38, 0.39 or
0.4. The fan tip loading may be in an inclusive range bounded by
any two of the values in the previous sentence (i.e. the values may
form upper or lower bounds).
[0043] Gas turbine engines in accordance with the present
disclosure may have any desired bypass ratio, where the bypass
ratio is defined as the ratio of the mass flow rate of the flow
through the bypass duct to the mass flow rate of the flow through
the core at cruise conditions. In some arrangements the bypass
ratio may be greater than (or on the order of) any of the
following: 10, 10.5, 11, 11.5, 12, 12.5, 13, 13.5, 14, 14.5, 15,
15.5, 16, 16.5, 17, 17.5, 18, 18.5, 19, 19.5 or 20. The bypass
ratio may be in an inclusive range bounded by any two of the values
in the previous sentence (i.e. the values may form upper or lower
bounds). The bypass duct may be substantially annular. The bypass
duct may be radially outside the core engine. The radially outer
surface of the bypass duct may be defined by a nacelle and/or a fan
case.
[0044] The overall pressure ratio of a gas turbine engine as
described and/or claimed herein may be defined as the ratio of the
stagnation pressure upstream of the fan to the stagnation pressure
at the exit of the highest pressure compressor (before entry into
the combustor). By way of non-limitative example, the overall
pressure ratio of a gas turbine engine as described and/or claimed
herein at cruise may be greater than (or on the order of) any of
the following: 35, 40, 45, 50, 55, 60, 65, 70, 75. The overall
pressure ratio may be in an inclusive range bounded by any two of
the values in the previous sentence (i.e. the values may form upper
or lower bounds).
[0045] Specific thrust of an engine may be defined as the net
thrust of the engine divided by the total mass flow through the
engine. At cruise conditions, the specific thrust of an engine
described and/or claimed herein may be less than (or on the order
of) any of the following: 110 Nkg.sup.-1s, 105 Nkg.sup.-1s, 100
Nkg.sup.-1s, 95 Nkg.sup.-1s, 90 Nkg.sup.-1s, 85 Nkg.sup.-1s or 80
Nkg.sup.-1s. The specific thrust may be in an inclusive range
bounded by any two of the values in the previous sentence (i.e. the
values may form upper or lower bounds). Such engines may be
particularly efficient in comparison with conventional gas turbine
engines.
[0046] A gas turbine engine as described and/or claimed herein may
have any desired maximum thrust. Purely by way of non-limitative
example, a gas turbine as described and/or claimed herein may be
capable of producing a maximum thrust of at least (or on the order
of) any of the following: 160 kN, 170 kN, 180 kN, 190 kN, 200 kN,
250 kN, 300 kN, 350 kN, 400 kN, 450 kN, 500 kN, or 550 kN. The
maximum thrust may be in an inclusive range bounded by any two of
the values in the previous sentence (i.e. the values may form upper
or lower bounds). The thrust referred to above may be the maximum
net thrust at standard atmospheric conditions at sea level plus 15
deg C. (ambient pressure 101.3 kPa, temperature 30 deg C.), with
the engine static.
[0047] In use, the temperature of the flow at the entry to the high
pressure turbine may be particularly high. This temperature, which
may be referred to as TET, may be measured at the exit to the
combustor, for example immediately upstream of the first turbine
vane, which itself may be referred to as a nozzle guide vane. At
cruise, the TET may be at least (or on the order of) any of the
following: 1400K, 1450K, 1500K, 1550K, 1600K or 1650K. The TET at
cruise may be in an inclusive range bounded by any two of the
values in the previous sentence (i.e. the values may form upper or
lower bounds). The maximum TET in use of the engine may be, for
example, at least (or on the order of) any of the following: 1700K,
1750K, 1800K, 1850K, 1900K, 1950K or 2000K. The maximum TET may be
in an inclusive range bounded by any two of the values in the
previous sentence (i.e. the values may form upper or lower bounds).
The maximum TET may occur, for example, at a high thrust condition,
for example at a maximum take-off (MTO) condition.
[0048] A fan blade and/or aerofoil portion of a fan blade described
and/or claimed herein may be manufactured from any suitable
material or combination of materials. For example at least a part
of the fan blade and/or aerofoil may be manufactured at least in
part from a composite, for example a metal matrix composite and/or
an organic matrix composite, such as carbon fibre. By way of
further example at least a part of the fan blade and/or aerofoil
may be manufactured at least in part from a metal, such as a
titanium based metal or an aluminium based material (such as an
aluminium-lithium alloy) or a steel based material. The fan blade
may comprise at least two regions manufactured using different
materials. For example, the fan blade may have a protective leading
edge, which may be manufactured using a material that is better
able to resist impact (for example from birds, ice or other
material) than the rest of the blade. Such a leading edge may, for
example, be manufactured using titanium or a titanium-based alloy.
Thus, purely by way of example, the fan blade may have a
carbon-fibre or aluminium based body (such as an aluminium lithium
alloy) with a titanium leading edge.
[0049] A fan as described and/or claimed herein may comprise a
central portion, from which the fan blades may extend, for example
in a radial direction. The fan blades may be attached to the
central portion in any desired manner. For example, each fan blade
may comprise a fixture which may engage a corresponding slot in the
hub (or disc). Purely by way of example, such a fixture may be in
the form of a dovetail that may slot into and/or engage a
corresponding slot in the hub/disc in order to fix the fan blade to
the hub/disc. By way of further example, the fan blades maybe
formed integrally with a central portion. Such an arrangement may
be referred to as a bladed disc or a bladed ring. Any suitable
method may be used to manufacture such a bladed disc or bladed
ring. For example, at least a part of the fan blades may be
machined from a block and/or at least part of the fan blades may be
attached to the hub/disc by welding, such as linear friction
welding.
[0050] The gas turbine engines described and/or claimed herein may
or may not be provided with a variable area nozzle (VAN). Such a
variable area nozzle may allow the exit area of the bypass duct to
be varied in use. The general principles of the present disclosure
may apply to engines with or without a VAN.
[0051] The fan of a gas turbine as described and/or claimed herein
may have any desired number of fan blades, for example 14, 16, 18,
20, 22, 24 or 26 fan blades.
[0052] As used herein, cruise conditions may mean cruise conditions
of an aircraft to which the gas turbine engine is attached. Such
cruise conditions may be conventionally defined as the conditions
at mid-cruise, for example the conditions experienced by the
aircraft and/or engine at the midpoint (in terms of time and/or
distance) between top of climb and start of decent.
[0053] Purely by way of example, the forward speed at the cruise
condition may be any point in the range of from Mach 0.7 to 0.9,
for example 0.75 to 0.85, for example 0.76 to 0.84, for example
0.77 to 0.83, for example 0.78 to 0.82, for example 0.79 to 0.81,
for example on the order of Mach 0.8, on the order of Mach 0.85 or
in the range of from 0.8 to 0.85. Any single speed within these
ranges may be the cruise condition. For some aircraft, the cruise
conditions may be outside these ranges, for example below Mach 0.7
or above Mach 0.9.
[0054] Purely by way of example, the cruise conditions may
correspond to standard atmospheric conditions at an altitude that
is in the range of from 10000 m to 15000 m, for example in the
range of from 10000 m to 12000 m, for example in the range of from
10400 m to 11600 m (around 38000 ft), for example in the range of
from 10500 m to 11500 m, for example in the range of from 10600 m
to 11400 m, for example in the range of from 10700 m (around 35000
ft) to 11300 m, for example in the range of from 10800 m to 11200
m, for example in the range of from 10900 m to 11100 m, for example
on the order of 11000 m. The cruise conditions may correspond to
standard atmospheric conditions at any given altitude in these
ranges.
[0055] Purely by way of example, the cruise conditions may
correspond to: a forward Mach number of 0.8; a pressure of 23000
Pa; and a temperature of -55 deg C.
[0056] As used anywhere herein, "cruise" or "cruise conditions" may
mean the aerodynamic design point. Such an aerodynamic design point
(or ADP) may correspond to the conditions (comprising, for example,
one or more of the Mach Number, environmental conditions and thrust
requirement) for which the fan is designed to operate. This may
mean, for example, the conditions at which the fan (or gas turbine
engine) is designed to have optimum efficiency.
[0057] In use, a gas turbine engine described and/or claimed herein
may operate at the cruise conditions defined elsewhere herein. Such
cruise conditions may be determined by the cruise conditions (for
example the mid-cruise conditions) of an aircraft to which at least
one (for example 2 or 4) gas turbine engine may be mounted in order
to provide propulsive thrust.
[0058] The skilled person will appreciate that except where
mutually exclusive, a feature or parameter described in relation to
any one of the first and second aspects may be applied to any other
aspect. Furthermore, except where mutually exclusive, any feature
or parameter described herein may be applied to any aspect and/or
combined with any other feature or parameter described herein.
[0059] Embodiments will now be described by way of example only,
with reference to the Figures, in which:
[0060] FIG. 1 is a sectional side view of a gas turbine engine;
[0061] FIG. 2 is a close up sectional side view of an upstream
portion of a gas turbine engine;
[0062] FIG. 3 is a partially cut-away view of a gearbox for a gas
turbine engine;
[0063] FIG. 4 is a schematic, sectional view of an inlet duct of a
gas turbine engine.
[0064] FIG. 1 illustrates a gas turbine engine 10 having a
principal rotational axis 9. The engine 10 comprises an air intake
12 and a propulsive fan 23 that generates two airflows: a core
airflow A and a bypass airflow B. The gas turbine engine 10
comprises a core 11 that receives the core airflow A. The engine
core 11 comprises, in axial flow series, a low pressure compressor
14, a high-pressure compressor 15, combustion equipment 16, a
high-pressure turbine 17, a low pressure turbine 19 and a core
exhaust nozzle 20. A nacelle 21 surrounds the gas turbine engine 10
and defines a bypass duct 22 and a bypass exhaust nozzle 18. The
bypass airflow B flows through the bypass duct 22. The fan 23 is
attached to and driven by the low pressure turbine 19 via a shaft
26 and an epicyclic gearbox 30.
[0065] The fan 23 may have a diameter of at least 390 cm, for
example a diameter of 410 cm.
[0066] In use, the core airflow A is accelerated and compressed by
the low pressure compressor 14 and directed into the high pressure
compressor 15 where further compression takes place. The compressed
air exhausted from the high pressure compressor 15 is directed into
the combustion equipment 16 where it is mixed with fuel and the
mixture is combusted. The resultant hot combustion products then
expand through, and thereby drive, the high pressure and low
pressure turbines 17, 19 before being exhausted through the nozzle
20 to provide some propulsive thrust. The high pressure turbine 17
drives the high pressure compressor 15 by a suitable
interconnecting shaft 27. The fan 23 generally provides the
majority of the propulsive thrust. The epicyclic gearbox 30 is a
reduction gearbox.
[0067] An exemplary arrangement for a geared fan gas turbine engine
10 is shown in FIG. 2. The low pressure turbine 19 (see FIG. 1)
drives the shaft 26, which is coupled to a sun wheel, or sun gear,
28 of the epicyclic gear arrangement 30. Radially outwardly of the
sun gear 28 and intermeshing therewith is a plurality of planet
gears 32 that are coupled together by a planet carrier 34. The
planet carrier 34 constrains the planet gears 32 to precess around
the sun gear 28 in synchronicity whilst enabling each planet gear
32 to rotate about its own axis. The planet carrier 34 is coupled
via linkages 36 to the fan 23 in order to drive its rotation about
the engine axis 9. Radially outwardly of the planet gears 32 and
intermeshing therewith is an annulus or ring gear 38 that is
coupled, via linkages 40, to a stationary supporting structure
24.
[0068] Note that the terms "low pressure turbine" and "low pressure
compressor" as used herein may be taken to mean the lowest pressure
turbine stages and lowest pressure compressor stages (i.e. not
including the fan 23) respectively and/or the turbine and
compressor stages that are connected together by the
interconnecting shaft 26 with the lowest rotational speed in the
engine (i.e. not including the gearbox output shaft that drives the
fan 23). In some literature, the "low pressure turbine" and "low
pressure compressor" referred to herein may alternatively be known
as the "intermediate pressure turbine" and "intermediate pressure
compressor". Where such alternative nomenclature is used, the fan
23 may be referred to as a first, or lowest pressure, compression
stage.
[0069] The epicyclic gearbox 30 is shown by way of example in
greater detail in FIG. 3. Each of the sun gear 28, planet gears 32
and ring gear 38 comprise teeth about their periphery to intermesh
with the other gears. However, for clarity only exemplary portions
of the teeth are illustrated in FIG. 3. There are four planet gears
32 illustrated, although it will be apparent to the skilled reader
that more or fewer planet gears 32 may be provided within the scope
of the claimed invention. Practical applications of a planetary
epicyclic gearbox 30 generally comprise at least three planet gears
32.
[0070] The epicyclic gearbox 30 illustrated by way of example in
FIGS. 2 and 3 is of the planetary type, in that the planet carrier
34 is coupled to an output shaft via linkages 36, with the ring
gear 38 fixed. However, any other suitable type of epicyclic
gearbox 30 may be used. By way of further example, the epicyclic
gearbox 30 may be a star arrangement, in which the planet carrier
34 is held fixed, with the ring (or annulus) gear 38 allowed to
rotate. In such an arrangement the fan 23 is driven by the ring
gear 38. By way of further alternative example, the gearbox 30 may
be a differential gearbox in which the ring gear 38 and the planet
carrier 34 are both allowed to rotate.
[0071] The gearbox 30 may have a reduction gear ratio comprised
between 2.5 and 4.2, for example between 3.6 and 3.8. In other
words, the interconnecting shaft 26, which is the input of the
gearbox 30, rotates between 2.5 and 4.2, for example 3.6, times
faster than the fan 23, either that the fan 23 is rotated by the
planet carrier 34 or the ring gear 38 depending on the type of
epicyclic gearbox 30.
[0072] It will be appreciated that the arrangement shown in FIGS. 2
and 3 is by way of example only, and various alternatives are
within the scope of the present disclosure. Purely by way of
example, any suitable arrangement may be used for locating the
gearbox 30 in the engine 10 and/or for connecting the gearbox 30 to
the engine 10. By way of further example, the connections (such as
the linkages 36, 40 in the FIG. 2 example) between the gearbox 30
and other parts of the engine 10 (such as the input shaft 26, the
output shaft and the fixed structure 24) may have any desired
degree of stiffness or flexibility. By way of further example, any
suitable arrangement of the bearings between rotating and
stationary parts of the engine (for example between the input and
output shafts from the gearbox and the fixed structures, such as
the gearbox casing) may be used, and the disclosure is not limited
to the exemplary arrangement of FIG. 2. For example, where the
gearbox 30 has a star arrangement (described above), the skilled
person would readily understand that the arrangement of output and
support linkages and bearing locations would typically be different
to that shown by way of example in FIG. 2.
[0073] Accordingly, the present disclosure extends to a gas turbine
engine having any arrangement of gearbox styles (for example star
or planetary), support structures, input and output shaft
arrangement, and bearing locations.
[0074] Optionally, the gearbox may drive additional and/or
alternative components (e.g. the intermediate pressure compressor
and/or a booster compressor).
[0075] Other gas turbine engines to which the present disclosure
may be applied may have alternative configurations. For example,
such engines may have an alternative number of compressors and/or
turbines and/or an alternative number of interconnecting shafts. By
way of further example, the gas turbine engine shown in FIG. 1 has
a split flow nozzle 18, 20 meaning that the flow through the bypass
duct 22 has its own nozzle 18 that is separate to and radially
outside the core engine nozzle 20. However, this is not limiting,
and any aspect of the present disclosure may also apply to engines
in which the flow through the bypass duct 22 and the flow through
the core 11 are mixed, or combined, before (or upstream of) a
single nozzle, which may be referred to as a mixed flow nozzle. One
or both nozzles (whether mixed or split flow) may have a fixed or
variable area. Whilst the described example relates to a turbofan
engine, the disclosure may apply, for example, to any type of gas
turbine engine, such as an open rotor (in which the fan stage is
not surrounded by a nacelle) or turboprop engine, for example. In
some arrangements, the gas turbine engine 10 may not comprise a
gearbox 30.
[0076] The geometry of the gas turbine engine 10, and components
thereof, is defined by a conventional axis system, comprising an
axial direction (which is aligned with the rotational axis 9), a
radial direction (in the bottom-to-top direction in FIG. 1), and a
circumferential direction (perpendicular to the page in the FIG. 1
view). The axial, radial and circumferential directions are
mutually perpendicular.
[0077] FIG. 4 illustrates an inlet duct 50 through which the core
flow A enters the core of the gas turbine engine 10.
[0078] Arranged in the inlet duct 50, there are in flow series an
engine stator section (ESS) 52 with a plurality of vanes 54, a
strut 56, variable inlet guide vanes (VIGV) 58, and a first rotor
60 of the low pressure compressor 14 with a plurality of first
blades 62.
[0079] In alternative, not illustrated embodiments, the strut 56
may be omitted.
[0080] The vanes 54 of the ESS 52 feature respective leading edges
64, a mid-span point 66 of which is arranged at a first radius R1
from the main longitudinal axis 9. A first annulus area A1 is
defined at the leading edge 64.
[0081] The first blades 62 are arranged downstream of the ESS 52
and feature respective leading edges 68, a mid-span point 70 of
which is arranged at a second radius R2 from the main longitudinal
axis 9. A second annulus area A2 is defined at the leading edge 68
of the first blades 62. In the illustrated embodiment the second
annulus area A2 is larger than the first annulus area A1, leading
to a ratio A2/A1 larger than 1. In non-illustrated embodiments, the
second annulus area A2 may be less than the first annulus area A1,
leading to a ratio A2/A1 less than 1 and an accelerating inlet
duct. In embodiments, the ratio A2/A1 may be equal to 1, or equal
to 0.95.
[0082] The ESS leading edge mid-span point 66 and the first blade
leading edge mid-span point 70 are separated by an axial distance
L. L represents an axial length of the inlet duct 50. Moreover, the
ESS leading edge mid-span point 66 and the first blade leading edge
mid-span point 70 are separated by a radial distance, which is the
difference .DELTA.R between the first radius R1 and the second
radius R2.
[0083] The ratio of the difference .DELTA.R between the first
radius R1 and the second radius R2 to the axial distance L
represents a mean slope of the inlet duct 50 and is related to the
inlet duct curvature. In the present embodiment, .DELTA.R/L is
greater than 0.3, for example between 0.35 and 0.55, for example
about 0.42. In other words, the inlet duct 50 has a relatively high
mean slope allowing the first radius R1 to be relatively large,
such that the ESS may be optimised for a relatively large fan
diameter with high bypass ratio, and the second radius R2 to be
relatively short, such that the compressor blade tip does not
rotate at high Mach numbers that would lead to poor efficiency and
greater noise, and turbine temperatures may be increased leading to
increased efficiency. Moreover, a relatively high mean slope allows
at the same time a relatively short (and therefore light) inlet
duct 50 (and therefore gas turbine engine).
[0084] It will be understood that the invention is not limited to
the embodiments above-described and various modifications and
improvements can be made without departing from the concepts
described herein. Except where mutually exclusive, any of the
features may be employed separately or in combination with any
other features and the disclosure extends to and includes all
combinations and sub-combinations of one or more features described
herein.
* * * * *