U.S. patent application number 17/196345 was filed with the patent office on 2021-11-11 for high pressure ratio gas turbine engine.
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, Michael O. HALES.
Application Number | 20210348555 17/196345 |
Document ID | / |
Family ID | 1000005782232 |
Filed Date | 2021-11-11 |
United States Patent
Application |
20210348555 |
Kind Code |
A1 |
BOUSFIELD; Ian J. ; et
al. |
November 11, 2021 |
HIGH PRESSURE RATIO GAS TURBINE ENGINE
Abstract
A gas turbine engine (10) comprising: a high pressure turbine
(17); a low pressure turbine (19); a high pressure compressor (15)
coupled to the high pressure turbine (17) by a high pressure shaft
(27); a propulsor (23) and a low pressure compressor (14) coupled
to the low pressure turbine (19) via a low pressure shaft (26) and
a reduction gearbox (30); wherein the low pressure compressor (14)
consists of four or five compressor stages (14); the high pressure
compressor (15) consists of eight or nine compressor stages; the
low pressure turbine (19) comprises four or more stages; and the
high pressure compressor (15) and low pressure compressor (14)
together define a core overall pressure ratio of greater than
36:1.
Inventors: |
BOUSFIELD; Ian J.;
(Nottingham, GB) ; HALES; Michael O.; (Bristol,
GB) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
ROLLS-ROYCE plc |
London |
|
GB |
|
|
Assignee: |
ROLLS-ROYCE plc
London
GB
|
Family ID: |
1000005782232 |
Appl. No.: |
17/196345 |
Filed: |
March 9, 2021 |
Current U.S.
Class: |
1/1 |
Current CPC
Class: |
F02C 3/06 20130101; F05D
2220/323 20130101 |
International
Class: |
F02C 3/06 20060101
F02C003/06 |
Foreign Application Data
Date |
Code |
Application Number |
Mar 26, 2020 |
GB |
2004382.4 |
Mar 26, 2020 |
GB |
2004384.0 |
Mar 26, 2020 |
GB |
2004385.7 |
Mar 26, 2020 |
GB |
2004386.5 |
Mar 26, 2020 |
GB |
2004387.3 |
Claims
1. A gas turbine engine comprising: a high pressure turbine; a low
pressure turbine; a high pressure compressor coupled to the high
pressure turbine by a high pressure shaft; a propulsor and a low
pressure compressor coupled to the low pressure turbine via a low
pressure shaft and a reduction gearbox; wherein the low pressure
compressor consists of four or five compressor stages; the high
pressure compressor consists of eight or nine compressor stages;
the low pressure turbine comprises four or more stages; and the
high pressure compressor and low pressure compressor together
define a core overall pressure ratio of greater than 36:1.
2. A gas turbine engine according to claim 1, wherein the core
overall pressure ratio is between 36:1 and 60:1.
3. A gas turbine engine according to claim 1, wherein the low
pressure compressor defines a cruise average stage pressure ratio
of between 1.24:1 and 1.35:1.
4. A gas turbine engine according to claim 1, wherein the low
pressure compressor defines a cruise pressure ratio of between 2.3
and 4.5.
5. A gas turbine engine according to claim 1, wherein the high
pressure compressor defines a cruise pressure ratio of between 8:1
and 18:1.
6. A gas turbine engine according to claim 1, wherein the high
pressure compressor defines a cruise average pressure ratio of
between 1.3 and 1.42.
7. A gas turbine engine according to claim 1, wherein the high
pressure turbine consists of two or fewer stages.
8. A gas turbine engine according to claim 1, wherein the low
pressure turbine consists of four stages.
9. A gas turbine engine according to claim 1, wherein the reduction
gearbox comprises a star gearbox.
10. A gas turbine engine according to claim 1, wherein the
reduction gearbox defines a reduction ratio of approximately 3.
11. A gas turbine engine according to claim 1, wherein the low
pressure compressor is positioned axially upstream of the high
pressure compressor.
12. A gas turbine engine according to claim 1, wherein the
propulsor is in the form of an open rotor, or a ducted fan.
13. A gas turbine engine according to claim 1, wherein each
compressor and/or turbine stage comprises a row of rotor blades and
a row of stator vanes, which may be variable stator vanes.
14. A method of operating the gas turbine engine of claim 1,
comprising, at cruise conditions, operating the low and high
pressure compressors to provide a pressure ratio of greater than
36:1.
Description
[0001] The present disclosure relates to gas turbine engine for an
aircraft
[0002] Existing gas turbine engines are known, in which a reduction
gearbox is provided between a turbine and a propulsive fan. Such
engines are known as "geared turbofans".
[0003] It desirable to increase the fuel efficiency of such
engines. There are essentially two methods to increase fuel
efficiency--increased propulsive efficiency, and increased thermal
efficiency. It is an objective of the present invention to provide
a gas turbine engine architecture that provides improved fuel
efficiency
[0004] According to a first aspect there is provided a gas turbine
engine comprising: [0005] a high pressure turbine; [0006] a low
pressure turbine; [0007] a high pressure compressor coupled to the
high pressure turbine by a high pressure shaft; [0008] a propulsor
and a low pressure compressor coupled to the low pressure turbine
via a low pressure shaft and a reduction gearbox; wherein [0009]
the low pressure compressor consists of four or five compressor
stages, [0010] the high pressure compressor consists of eight or
nine compressor stages; [0011] the low pressure turbine comprises
four or more stages; and [0012] the high pressure compressor and
low pressure compressor together define a core overall pressure
ratio at cruise of greater than 36:1.
[0013] The inventors have found that the above defined
characteristics can provide a gas turbine engine having a high
overall pressure ratio (and so high thermal efficiency), with
relatively few stages, and with a relatively low pressure ratio
high pressure compressor. Such a relatively low pressure ratio high
pressure compressor can provide for numerous advantages, such as a
reduction in variable stator stages and bleed valves, which can in
turn result in reduced weight and cost.
[0014] The core overall pressure ratio may be between 36:1 and
60:1. The overall pressure ratio may be any of 36:1, 38:1, 40:1,
42:1, 44:1, 46:1, 48:1, 50:1, 52:1, 54:1, 56:1 or 58:1.
[0015] The low pressure compressor may define a cruise average
stage pressure ratio of between 1.24:1 and 1.35:1.
[0016] The low pressure compressor may define a cruise pressure
ratio of between 2.3 and 4.5.
[0017] The high pressure compressor may define a cruise pressure
ratio of between 8:1 and 18:1.
[0018] The high pressure compressor may define a cruise average
pressure ratio of between 1.3 and 1.42.
[0019] The high pressure turbine may consist of two or fewer
stages.
[0020] The low pressure turbine may consist of four stages.
[0021] The low pressure compressor may be positioned axially
upstream of the high pressure compressor. The high pressure
compressor may be arranged to receive (for example directly
receive, for example via a generally annular duct) flow from the
low pressure compressor.
[0022] The gearbox may be arranged to be driven by the shaft that
is configured to rotate (for example in use) at the lowest
rotational speed (for example the low pressure shaft in the example
above).
[0023] The engine may comprise a core casing and a nacelle, wherein
at least one of the core casing the nacelle comprise carbon
composite material.
[0024] Any type of reduction 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, or 3.2
to 3.8, 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. Purely by way of example, the gearbox may be a "star"
gearbox having a ratio in the range of from 3.1 or 3.2 to 3.8. In
some arrangements, the gear ratio may be outside these ranges.
[0025] In any gas turbine engine as described and/or claimed
herein, a combustor may be provided axially downstream of the
propulsor and compressors. For example, the combustor may be
directly downstream of (for example at the exit of) the high
pressure compressor. By way of further example, the flow at the
exit to the combustor may be provided to the inlet of the high
pressure turbine.
[0026] The propulsor may be in the form of an open rotor, or a
ducted fan.
[0027] Each compressor and/or turbine 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.
[0028] 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), for example in the range of from 0.28 to 0.32. 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. 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, 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), for example in the range of from 240
cm to 280 cm or 330 cm to 380 cm.
[0029] 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 240 cm to 280 cm or 250 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 330 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 1800 rpm.
[0030] 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 13 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, 0.30, 0.31, 0.32, 0.33, 0.34, 0.35, 0.36, 0.37, 0.38, 0.39 or
0.4 (all values being dimensionless). 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), for
example in the range of from 0.28 to 0.31, or 0.29 to 0.3.
[0031] 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), for example in the range of form 12 to 16, 13 to 15, or 13
to 14. 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.
[0032] The overall core pressure ratio of a gas turbine engine as
described and/or claimed herein may be defined as the ratio of the
stagnation pressure downstream 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: 36, 40, 45, 50, 55. The overall core 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).
[0033] 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), for example in the range of
from 80 Nkg.sup.-1s to 100 Nkg.sup.-1s, or 85 Nkg.sup.-1s to 95
Nkg.sup.-1s. Such engines may be particularly efficient in
comparison with conventional gas turbine engines.
[0034] 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). Purely by way of example, a gas turbine as
described and/or claimed herein may be capable of producing a
maximum thrust in the range of from 330 kN to 420 kN, for example
350 kN to 400 kN. The thrust referred to above may be the maximum
net thrust at standard atmospheric conditions at sea level plus 15
degrees C. (ambient pressure 101.3 kPa, temperature 30 degrees C.),
with the engine static.
[0035] 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: 1400 K, 1450 K, 1500 K, 1550 K, 1600 K or 1650 K. 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: 1700
K, 1750 K, 1800 K, 1850 K, 1900 K, 1950 K or 2000 K. 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), for example in the range of from 1800 K to 1950 K. The
maximum TET may occur, for example, at a high thrust condition, for
example at a maximum take-off (MTO) condition.
[0036] 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.
[0037] 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.
[0038] 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. 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.
[0039] As used herein, cruise conditions have the conventional
meaning and would be readily understood by the skilled person.
Thus, for a given gas turbine engine for an aircraft, the skilled
person would immediately recognise cruise conditions to mean the
operating point of the engine at mid-cruise of a given mission
(which may be referred to in the industry as the "economic
mission") of an aircraft to which the gas turbine engine is
designed to be attached. In this regard, mid-cruise is the point in
an aircraft flight cycle at which 50% of the total fuel that is
burned between top of climb and start of descent has been burned
(which may be approximated by the midpoint--in terms of time and/or
distance--between top of climb and start of descent. Cruise
conditions thus define an operating point of, the gas turbine
engine that provides a thrust that would ensure steady state
operation (i.e. maintaining a constant altitude and constant Mach
Number) at mid-cruise of an aircraft to which it is designed to be
attached, taking into account the number of engines provided to
that aircraft. For example where an engine is designed to be
attached to an aircraft that has two engines of the same type, at
cruise conditions the engine provides half of the total thrust that
would be required for steady state operation of that aircraft at
mid-cruise.
[0040] In other words, for a given gas turbine engine for an
aircraft, cruise conditions are defined as the operating point of
the engine that provides a specified thrust (required to
provide--in combination with any other engines on the
aircraft--steady state operation of the aircraft to which it is
designed to be attached at a given mid-cruise Mach Number) at the
mid-cruise atmospheric conditions (defined by the International
Standard Atmosphere according to ISO 2533 at the mid-cruise
altitude). For any given gas turbine engine for an aircraft, the
mid-cruise thrust, atmospheric conditions and Mach Number are
known, and thus the operating point of the engine at cruise
conditions is clearly defined.
[0041] 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 part of 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.
[0042] Purely by way of example, the cruise conditions may
correspond to standard atmospheric conditions (according to the
International Standard Atmosphere, ISA) 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.
[0043] Purely by way of example, the cruise conditions may
correspond to an operating point of the engine that provides a
known required thrust level (for example a value in the range of
from 30 kN to 35 kN) at a forward Mach number of 0.8 and standard
atmospheric conditions (according to the International Standard
Atmosphere) at an altitude of 38000 ft (11582 m). Purely by way of
further example, the cruise conditions may correspond to an
operating point of the engine that provides a known required thrust
level (for example a value in the range of from 50 kN to 65 kN) at
a forward Mach number of 0.85 and standard atmospheric conditions
(according to the International Standard Atmosphere) at an altitude
of 35000 ft (10668 m).
[0044] 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.
[0045] According to an aspect, there is provided an aircraft
comprising a gas turbine engine as described and/or claimed herein.
The aircraft according to this aspect is the aircraft for which the
gas turbine engine has been designed to be attached. Accordingly,
the cruise conditions according to this aspect correspond to the
mid-cruise of the aircraft, as defined elsewhere herein.
[0046] According to an aspect, there is provided a method of
operating a gas turbine engine as described and/or claimed herein.
The operation may be at the cruise conditions as defined elsewhere
herein (for example in terms of the thrust, atmospheric conditions
and Mach Number).
[0047] According to an aspect, there is provided a method of
operating an aircraft comprising a gas turbine engine as described
and/or claimed herein. The operation according to this aspect may
include (or may be) operation at the mid-cruise of the aircraft, as
defined elsewhere herein.
[0048] The method may comprise, at cruise conditions, operating the
low pressure compressor (14) to provide a pressure ratio of between
2.4:1 and 3.3:1, operating the high pressure compressor (15) to
provide a pressure ratio of less than 18:1, and operating the low
and high pressure compressors (14, 15) to provide a pressure ratio
of less than 36:1.
[0049] The skilled person will appreciate that except where
mutually exclusive, a feature or parameter described in relation to
any one of the above 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.
[0050] Embodiments will now be described by way of example only,
with reference to the Figures, in which:
[0051] FIG. 1 is a sectional side view of a gas turbine engine;
[0052] FIG. 2 is a close up sectional side view of an upstream
portion of the gas turbine engine of FIG. 1;
[0053] FIG. 3 is a close up sectional side view of a turbine
section of the gas turbine engine of FIG. 1;
[0054] FIG. 4 is a sectional front view of a reduction gearbox of
the gas turbine engine of FIG. 1;
[0055] FIG. 5 is a graph illustrating a design space for the
compressor section of FIG. 2;
[0056] FIG. 6 is a graph illustrating a design space for high
pressure compressors of gas turbine engines in accordance with the
present disclosure; and
[0057] FIG. 7 is a graph illustrating a design space for low
pressure compressors of gas turbine engines in accordance with the
present disclosure.
[0058] 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. In some cases, the low pressure compressor 14
may be referred to as an intermediate pressure compressor (IPC).
Similarly, the low pressure turbine may be referred to as an
intermediate pressure turbine (IPT). 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.
[0059] The fan 23 is attached to and driven by the low pressure
turbine 19 via a shaft 26 and an epicyclic gearbox 30.
[0060] The engine core 11 is surrounded by a core casing 37, which
contains the compressor 14, 15, combustor 16 and turbines 17, 19.
The core casing 37 comprises one or more handling bleeds comprising
one or more valves 38 configured to communicate between the core
compressor flow path A (e.g. at the downstream end of the high
pressure compressor 15) and the fan flowpath B. The core casing 37
comprises a carbon composite material such as carbon fibre
reinforce plastic (CFRP).
[0061] Similarly, the engine nacelle 21 comprises a carbon
composite material, such as CFRP. For example, a Thrust Reverser
Unit (TRU) 39 provided at a rear of the nacelle 21 may comprise
carbon composite material.
[0062] 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.
[0063] An exemplary arrangement for a geared fan gas turbine engine
10 is shown in FIG. 4. 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 is statically mounted, which constrains the
planet gears 32 whilst enabling each planet gear 32 to rotate about
its own axis. Radially outwardly of the planet gears 32 and
intermeshing therewith is an annulus or ring gear 38 that that is
coupled via linkages 36 to the fan 23 in order to drive its
rotation about the engine axis 9. Such an arrangement is typically
referred to as a "star" gearbox.
[0064] 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.
[0065] The epicyclic gearbox 30 is shown by way of example in
greater detail in FIG. 4. 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. 4. 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.
[0066] The epicyclic gearbox 30 illustrated by way of example in
FIG. 4 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.
[0067] It will be appreciated that the arrangement shown in FIG. 4
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. 1 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.
[0068] 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.
[0069] Other gas turbine engines to which the present disclosure
may be applied may have alternative configurations. 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.
[0070] 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.
[0071] Referring now to FIG. 2, the low pressure and high pressure
compressor 14, 15 are shown in more detail. As can be seen, each of
the compressors comprises a multi-stage, axial flow compressor.
[0072] The low pressure compressor consists of four or five stages
(i.e. no more than five stages, and no fewer than four stages)
41a-d. Each stage 41a-d comprises at least one respective
compressor rotor 43, and may comprise a respective stator 44. The
respective rotor 43 and stator 44 are generally axially spaced. In
the present case, the first stator 44 is downstream in core flow of
the first rotor 43. One or more further stators such as an inlet
stator (not shown) may be provided--however, since no additional
rotor is associated with the inlet stator, this does not constitute
an additional stage, since no pressure rise is provided by the
inlet stator alone. As will be appreciated by the person skilled in
the art, the rotors 43 are coupled to the respective shaft (i.e.
the low pressure shaft 26 in the case of the low pressure
compressor 14) by corresponding discs 46a-d, and so turn with the
shaft 26. On the other hand, the stators 44 are held stationary. In
some cases, the stators 44 may pivot about their long axes, to
adjust the angle of attack and inlet and outlet area for the
respective compressor stage. Such stators are known as "variable
stator vanes" or VSVs.
[0073] The high pressure compressor 15 similarly consists of either
eight or nine stages, and in the described embodiment consists of
eight stages. Again, each stage comprises at least a rotor, and may
also comprise a stator.
[0074] The turbine is shown in FIG. 3. To drive the high pressure
compressor 15, a high pressure turbine 17 having two stages 47a,
47b may be necessary. Again, the number of turbine stages can be
determined in a similar manner to the number of compressor stages.
Alternatively, a single turbine stage may be provided. In
particular, it has been found that high pressure compressors having
cruise pressure ratios up to 13:1 can be driven by single stage
turbines.
[0075] Similarly, to drive the low pressure compressor 14 and fan
23, at least four low pressure turbine stages 49a-d are provided.
In some cases, five low pressure turbine stages may be
provided.
[0076] Between them, the high and low pressure compressors 15, 16
define a maximum in use overall core pressure ratio (OPR). The core
OPR is defined as the ratio of the stagnation pressure upstream of
the first stage 44 of the low pressure compressor 15 to the
stagnation pressure at the exit of the highest pressure compressor
16 (before entry into the combustor). The core OPR excludes any
pressure rise generated by the fan 23 where the fan provides air
flow to the core, so a total engine overall pressure ratio (EPR)
may be higher than the core OPR. In the present disclosure, the
overall core OPR is at least 36:1, and may be between 36:1 and
56:1. In the described embodiment, the core OPR is 40, and may take
any value between these upper and lower bounds. For example, the
core OPR may be any of 36, 40, 45, 50, 55 and 60, or even
higher.
[0077] As will be understood, the core OPR will vary according to
atmospheric, flight and engine conditions. However, the cruise OPR
is as defined above.
[0078] As will be understood, a large design space must be
considered when designing a gas turbine engine to determine an
optimal engine with respect to a chosen metric (such as engine
weight, cost, thermal efficiency, propulsive efficiency, or a
balance of these). In many cases, there may be a large number of
feasible solutions for a given set of conditions to achieve a
desired metric.
[0079] One such variable is core OPR. As core OPR increases,
thermal efficiency also tends to increase, and so a high OPR is
desirable. Even once a particular OPR is chosen however, a number
of design variables must be chosen to meet the chosen OPR.
[0080] One such design variable is the amount of pressure rise
provided by the low pressure compressor 15 relative to that
provided by the high pressure compressor 16 (sometimes referred to
as "worksplit"). As will be understood, the total core OPR can be
determined by multiplying the low pressure compressor pressure
ratio (i.e. the ratio between the stagnation pressure at the outlet
of the low pressure compressor to the stagnation pressure at the
inlet of the low pressure compressor 15) by the high pressure
compressor ratio (i.e. the ratio between the stagnation pressure at
the outlet of the high pressure compressor 16 to the stagnation
pressure at the inlet of the high pressure compressor 16).
Consequently, a higher core OPR can be provided by increasing the
high pressure compressor ratio, the low pressure compressor ratio,
or both.
[0081] The inventors have found that a particularly efficient work
split for a gas turbine engine having a core OPR in the above
described range can be provided by providing a low pressure
compressor 14 consisting of four or five stages, and having a
pressure ratio of between 2.4:1 and 4.5:1, and a high pressure
compressor is then provided having a pressure ratio above 8:1, such
that the overall core pressure ratio is above 36:1. It has been
found to be feasible to provide a pressure ratio of 18:1 on a high
pressure compressor provided on a single shaft using current
technology using a reasonable number of compressor stages, without
requiring an excessive number of variable stages, and at a
reasonable rotational speed to give high overall efficiency.
Consequently, to provide the necessary core OPR, a low pressure
compressor ratio of between 2.4:1 and 4.5:1 is required.
[0082] Similarly, there are several ways to increase the compressor
pressure ratio. A first method is to increase the stage loading.
Stage loading is defined as the stagnation pressure ratio across an
individual stage (rotor and stator) of a compressor. Similarly, an
average stage loading can be defined as the sum of the stage
loadings of each compressor stage of a compressor, divided by the
number of stages. For example, in the present disclosure, the
average stage loading of the low pressure compressor 14 is between
1.24 and 1.35. This can in turn be managed by one or more of
increasing the rotor speed at the maximum compression conditions,
increasing the turning provided by the blades, or increasing the
radius of the tips of the compressor rotors, which in turn
necessitates an increase in the radius of the roots of the
compressor rotors to maintain a given flow area. Each of these
options has associated advantages and disadvantages. For instance,
increasing low pressure compressor rotor speed necessitates either
an increase in the reduction ratio of the gearbox 30, or a
reduction in the fan 23 radius, in order to maintain fan tip speeds
at a desired level for noise and efficiency reasons. On the other
hand, increasing the compressor tip radius necessitates an increase
in weight, in view of the larger compressor discs that are
required. Increased turning of the airflow may result in lower
surge margin, and reduced efficiency. In any case, a higher stage
loading may result in a lower efficiency, since the increased rotor
tip speed or higher turning leads to lower compressor efficiencies,
in view of losses associated with aerodynamic shocks as the tips
significantly exceed the speed of sound.
[0083] A second option is to increase the number of stages in the
respective compressors, thereby maintaining a low stage loading,
low rotational speed, and low disc weight. Again, this can be
achieved by adding a stage to either the low pressure compressor 15
or high pressure compressor 16. However, this will generally result
in a higher weight and cost associated with the additional
stage.
[0084] A further complication is the presence of the gearbox 30.
The gearbox provides additional design freedom, since, as noted
above, the gearbox reduction ratio can be selected to provide a
preferred fan tip speed independently of both fan radius and low
pressure compressor rotor speed. However, the gearbox also presents
constraints in view of its large size. Consequently, the large
radius required radially inward of the fan 23 inherent in a geared
turbofan having an epicyclic gearbox dictates a fan 23 having a
large hub radius, i.e. a large radial distance between the engine
centre 9 and the aerodynamic root of the fan blades 23.
Furthermore, in view of the relatively slow turning fan typical of
geared turbofans, relatively little pressure rise is provided by
the inner radius of the fan 23, and so geared turbofans tend to
have a high hub to tip ratio fan 23.
[0085] The inventors have explored this design space, and found an
optimum range of stage numbers and compressor pressure ratios, that
provides an optimal mix of weight and efficiency.
[0086] In the present disclosure, a relatively large pressure ratio
is provided by the low pressure compressor 14, compared to prior
geared turbofans. This is achieved using a four or five stage
compressor 14, which rotates at relatively low speed. This low
speed allows a relatively low ratio gearbox to be employed
(typically of the order of 3:1), while maintaining a relatively low
fan blade tip speed, with a high diameter fan. The permits a "star"
gearbox, as described above, in conjunction with a high bypass
ratio (typically greater than 10). Such an arrangement has several
advantages over planetary gearboxes (in which the ring gear is held
static, and the planet carrier is used to drive the fan), such as
more convenient oil distribution.
[0087] In view of the relatively high pressure high work provided
by the low pressure compressor 14, additional work is required from
the low pressure turbine 19. In view of the relatively low shaft
speed, the inventors have found that a fourth or even a fifth
turbine stage is necessary, which is unusual on a geared turbofan.
This is because the low rotational speed results in low turbine tip
speeds, which results in relatively low work per revolution.
Consequently, a four or five stage low pressure turbine is provided
in this case.
[0088] The above optimum parameters define a design space for the
compressor, as illustrated in FIGS. 6 and 7.
[0089] One corner of the design space is defined by the maximum low
pressure compressor 14 cruise pressure ratio (4.5:1), and the
minimum high pressure compressor 15 cruise pressure ratio (8:1) to
achieve the minimum require overall core pressure ratio (36:1).
Above this low pressure cruise pressure ratio (4.5:1), it has been
found that compressor stability cannot be assured, without
increasing either rotational speed or diameter (and so compressor
blade tip speed in either case), or increasing low pressure
compressor stage count above five stages. However, where compressor
tip speed is increased, efficiency begins to fall, and so the
advantages of higher loading are lost. The inventors have found
that a low pressure cruise pressure ratio of 4.5:1 can be provided
with no more than five stages. Indeed, the inventors have found
that a cruise low pressure ratio of up to 3.3:1 can be provided
with only four low pressure compressor stages. Similarly, overall
engine efficiency suffers when the overall core engine pressure
ratio falls below 36:1, particularly in view of the relatively
small pressure rise generated by the fan in a geared turbofan.
[0090] A second corner of the design space is defined by the
maximum low pressure compressor 14 cruise pressure ratio (4.5:1),
and the maximum high pressure compressor 15 cruise pressure ratio
(18:1) that can reasonably be sustained, without requiring
excessive stage numbers, and increased weight. This combination
gives an overall core pressure ratio of 80:1. Above this value,
increases in thermal efficiency begin to be outweighed by increases
in weight, and so the design goals of increased overall propulsion
system efficiency are not achieved. In particular, the inventors
have found that the above parameters can be provided using a high
pressure compressor having eight or nine stages, with relatively
low work per stage. This relatively low work per stage provides for
high compressor efficiency, while the high overall pressure ratio
results in high overall engine efficiency.
[0091] A third corner of the design space is defined by the maximum
high pressure compressor 15 cruise pressure ratio (18:1) that can
reasonably be sustained, and the minimum low pressure compressor
cruise pressure ratio (2.4:1) that requires four compressor stages.
Below this value, only three low pressure compressor stages are
required, and so the engine weight is not optimised. This
combination gives an overall core pressure ratio of 43:1, which
provides good thermal efficiency, with a small number of overall
compressor stages.
[0092] A fourth corner of the design space is defined by the
minimum high pressure compressor cruise pressure ratio required to
achieve the required overall core pressure ratio of 36:1 at the
minimum low pressure compressor 14 stage loading for which four
compressor stages are required (2.4:1). This gives a high pressure
compressor cruise pressure ratio of approximately 15:1.
[0093] A fifth corner of the design space is defined. At this
point, a higher low pressure compressor cruise pressure ratio of
3.0:1 is provided, and a lower high pressure compressor cruise
pressure ratio of 12:1 is provided, while providing the minimum
overall core compressor pressure ratio of 36:1.
[0094] In one example, such as that shown in FIG. 2, the engine
comprises a low pressure compressor consisting of four stages, and
a high pressure compressor consisting of eight stages. FIG. 5
illustrates the design space for this combination. As can be seen,
the high pressure compressor overall pressure ratio can vary
between 15 and 17:1, and the low pressure compressor pressure ratio
can vary between 2.7:1 and 3.3:1. Such a combination is thought to
allow for a minimum total stage count, while providing a high
overall pressure ratio, since the loading of both spools is
maximised without compromising stability.
[0095] The designer is hence taught how to design a compressor
which achieves the desired characteristics of high overall core
cruise pressure ratio (greater than 36:1), while minimising stage
count and maximising compressor efficiency.
[0096] Two example gas turbine engines that have been considered by
the inventors are described below.
[0097] A first example engine has a maximum take-off thrust at sea
level under ISO conditions of approximately 45,000 pounds-force
(lbf). The low pressure compressor has four stages, and is
configured to provide a cruise pressure ratio of approximately
2.8:1. The high pressure compressor is configured to provide a
cruise pressure ratio of approximately 13:1. This gives an overall
core pressure ratio of approximately 36:1. Such an engine is
thought to provide an optimum mix of weight and thermal efficiency
for an engine in this class, since weight is a more important
factor in this class than for higher thrust engines, in view of the
shorter typical mission ranges of aircraft for which engines of
this thrust are designed.
[0098] A second example engine has a maximum take-off thrust at sea
level under ISO conditions of approximately 84,000 pounds-force
(lbf). The low pressure compressor has four stages, and is
configured to provide a cruise pressure ratio of approximately
2.8:1. The high pressure compressor is configured to provide a
cruise pressure ratio of approximately 17:1. This gives an overall
core pressure ratio of approximately 48:1. Such an engine is
thought to provide an optimum mix of weight and thermal efficiency
for an engine in this class, since thermal efficiency is a more
important factor in this class than for lower thrust engines, in
view of the longer typical mission ranges of aircraft for which
engines of this thrust are designed.
[0099] 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.
* * * * *