U.S. patent application number 15/894206 was filed with the patent office on 2018-08-16 for gas turbine engine fan blade with axial lean.
This patent application is currently assigned to ROLLS-ROYCE plc. The applicant listed for this patent is ROLLS-ROYCE plc. Invention is credited to Marco BARALE, Gabriel GONZALEZ-GUTIERREZ, Kashmir S. JOHAL, Benedict PHELPS, Nigel HS SMITH, Mark J. WILSON.
Application Number | 20180231019 15/894206 |
Document ID | / |
Family ID | 58462133 |
Filed Date | 2018-08-16 |
United States Patent
Application |
20180231019 |
Kind Code |
A1 |
BARALE; Marco ; et
al. |
August 16, 2018 |
GAS TURBINE ENGINE FAN BLADE WITH AXIAL LEAN
Abstract
A fan blade for a gas turbine engine is provided with forward
axial lean. The fan blade may have a substantially straight leading
edge. The geometry of the fan blade results in a lower
susceptibility to flutter, thereby allowing a gas turbine engine
comprising such a fan blade to operate over a wider range of
operating conditions.
Inventors: |
BARALE; Marco; (Derby,
GB) ; GONZALEZ-GUTIERREZ; Gabriel; (Derby, GB)
; WILSON; Mark J.; (Nottingham, GB) ; PHELPS;
Benedict; (Derby, GB) ; JOHAL; Kashmir S.;
(Derby, GB) ; SMITH; Nigel HS; (Derby,
GB) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
ROLLS-ROYCE plc |
London |
|
GB |
|
|
Assignee: |
ROLLS-ROYCE plc
London
GB
|
Family ID: |
58462133 |
Appl. No.: |
15/894206 |
Filed: |
February 12, 2018 |
Current U.S.
Class: |
1/1 |
Current CPC
Class: |
Y02T 50/672 20130101;
F05D 2240/303 20130101; F05D 2230/239 20130101; Y02T 50/673
20130101; F05D 2220/36 20130101; F04D 29/386 20130101; F04D 29/325
20130101; Y02T 50/60 20130101; F01D 5/141 20130101 |
International
Class: |
F04D 29/38 20060101
F04D029/38; F04D 29/32 20060101 F04D029/32 |
Foreign Application Data
Date |
Code |
Application Number |
Feb 14, 2017 |
GB |
1702380.5 |
Claims
1. A fan blade for a gas turbine engine, the gas turbine engine
defining axial, radial and circumferential directions, the fan
blade comprising: an aerofoil portion having a leading edge
extending from a root to a tip, the radial distance between the
leading edge at the root and the leading edge at the tip defining a
blade span wherein, when viewed along a circumferential direction:
the angle (.alpha.) formed between the radial direction and a
straight line (AC) drawn between the leading edge at the root and
at the tip is in the range of from -6.degree. and -0.2.degree.; and
the angle (.alpha. (P1, P2)) formed between the radial direction
and a line drawn between any two points (P1, P2) on the leading
edge that have a difference in radius of at least 5% of the blade
span is in the range of from -6.degree. and 0.degree., where a
negative angle indicates that the respective line has an axial
component that is in the same direction as the axial component of
the direction from a trailing edge to the leading edge of the
blade.
2. A fan blade for a gas turbine engine according to claim 1,
wherein: when viewed along a circumferential direction, the maximum
perpendicular distance (e) between any point on the leading edge
and a straight line drawn between the leading edge at the root and
at the tip is 2% of the blade span.
3. A fan blade for a gas turbine engine according to claim 1,
wherein the fan blade comprises: a platform; and a root portion,
wherein the root portion extends between the platform and the root
of the aerofoil portion.
4. A fan blade for a gas turbine engine according to claim 3,
wherein the radial extent of the root portion is no more than 7% of
the span of the aerofoil portion.
5. A fan blade for a gas turbine engine according to claim 1,
wherein the fan blade comprises a tip portion that extends at least
radially away from the tip of the aerofoil portion.
6. A fan blade for a gas turbine engine according to claim 5,
wherein the radial extent of the tip portion is no more than 7% of
the span of the aerofoil portion.
7. A fan blade for a gas turbine engine according to claim 1,
wherein: for two points on the leading edge that are radially
closer to the tip than to the root and have a difference in radius
of at least 5% of the blade span, the axial position of the
radially outer point is forward of the axial position of the
radially inner point.
8. A fan blade for a gas turbine engine according to claim 7,
wherein: for two points on the leading edge that are radially
closer to the tip than to the root and have a difference in radius
of at least 2% of the blade span, the axial position of the
radially outer point is forward of the axial position of the
radially inner point.
9. A fan blade for a gas turbine engine according to claim 1,
wherein the aerofoil portion has a trailing edge extending from the
root the tip, wherein, when viewed along a circumferential
direction the angle (.beta.) formed between the radial direction
and a straight line (BD) drawn between the trailing edge at the
root and at the tip is in the range of from -20.degree. and
-5.degree., where a negative angle indicates that the respective
line has an axial component that is in the same direction as the
axial component of the direction from a trailing edge to the
leading edge of the blade.
10. A fan stage for a gas turbine engine comprising: a hub; and a
plurality of fan blades according to claim 1, wherein: the fan
blades extend radially form the hub.
11. A fan stage for a gas turbine engine according to claim 10,
wherein: the ratio of the radius of the position where the leading
edge of one of the fan blades meets the hub to the outermost radial
extent of the leading edge of the fan blade is less than 0.33.
12. A gas turbine engine comprising a fan blade according to claim
1.
13. A gas turbine engine comprising a fan stage according to claim
10.
14. A gas turbine engine comprising: a fan having a plurality of
fan blades according to claim 1; a turbine; and a gearbox, wherein:
the fan is driven from the turbine via the gearbox, in order to
reduce the rotational speed of the fan stage compared with the
driving turbine stage.
15. A gas turbine engine according to claim 12 with a specific
thrust of less than 100 N/Kg/s.
16. A gas turbine engine according to claim 13 with a specific
thrust of less than 100 N/Kg/s.
17. A gas turbine engine according to claim 14 with a specific
thrust of less than 100 N/Kg/s.
18. A method of manufacturing a fan stage for a gas turbine engine
comprising: providing a fan hub; and attaching a plurality of fan
blades the fan hub using linear friction welding, the fan blades
comprising an aerofoil portion having a leading edge extending from
a root to a tip, the radial distance between the leading edge at
the root and the leading edge at the tip defining a blade span
wherein, when viewed along a circumferential direction,the angle
(.alpha.) formed between the radial direction and a straight line
(AC) drawn between the leading edge at the root and at the tip is
in the range of from -6.degree. and -0.2.degree.; and the angle
(.alpha. (P1, P2)) formed between the radial direction and a line
drawn between any two points (P1, P2) on the leading edge that have
a difference in radius of at least 5% of the blade span is in the
range of from -6.degree. and 0.degree., where a negative angle
indicates that the respective line has an axial component that is
in the same direction as the axial component of the direction from
a trailing edge to the leading edge of the blade.
Description
CROSS-REFERENCE TO RELATED APPLICATIONS
[0001] This specification is based upon and claims the benefit of
priority from UK Patent Application Number 1702380.5 filed on 14
Feb. 2017, the entire contents of which are incorporated herein by
reference.
BACKGROUND
1. Field of the Disclosure
[0002] This disclosure relates to a fan blade for a gas turbine
engine, a fan stage comprising at least one such fan blade, and a
gas turbine engine comprising such a fan stage.
2. Description of the Related Art
[0003] Modern gas turbine aero-engines typically comprise a fan,
which compresses the incoming air and directs at least a portion of
that air along a bypass duct, with the remainder of the air flowing
through the engine core. The fan blades of such gas turbine engines
may be susceptible to a known phenomenon called flutter. Flutter
may occur at certain engine operating conditions, for example
certain rotational speeds and/or thrusts and/or combinations
thereof.
[0004] Flutter may be characterized as a self-excited vibration.
When the aerofoils in a blade row (such as the fan blades in a gas
turbine engine fan) vibrate, they generate unsteady aerodynamic
forces on the blade row itself. Under most conditions, these
unsteady aerodynamic forces cause the blade row to do work on the
surrounding air, and the vibrations decay in amplitude. However, at
certain operational conditions, the surrounding air can do work on
the fan itself. If the work done by the air exceeds the work
dissipated by mechanical damping, then the vibrations will grow.
This instability is known as flutter.
[0005] Modern large gas turbine engines are being designed to have
lower specific thrust and higher fan tip loading than their
predecessors. This may be achieved by driving the fan via a gearbox
in order to reduce the rotational speed of the fan. Lower specific
thrust and/or lower rotational speed and/or higher tip loading may
be beneficial from an efficiency perspective, but may present
significant operability challenges.
[0006] For example, as the cruise and sea level working lines
separate at lower pressure ratios, the challenge to have sufficient
stall and flutter margins relative to the sea level static (SLS)
working line, and acceptable cruise working line efficiency becomes
more difficult.
[0007] Accordingly, the design of modern turbofan gas turbine
engines tends to increase the susceptibility of fan blades to
experience flutter. Flutter is undesirable because it can generate
large stresses in an engine.
[0008] Accordingly, it would be desirable to be able to reduce the
susceptibility of the fan blades in an engine to flutter.
SUMMARY
[0009] According to an aspect, there is provided a fan blade for a
gas turbine engine, the gas turbine engine defining axial, radial
and circumferential directions, the fan blade comprising: [0010] an
aerofoil portion having a leading edge extending from a root to a
tip, the radial distance between the leading edge at the root and
the leading edge at the tip defining a blade span wherein, when
viewed along a circumferential direction: [0011] the angle formed
between the radial direction and a straight line drawn between the
leading edge at the root and at the tip is in the range of from
-6.degree. and -0.2.degree.; and [0012] the angle formed between
the radial direction and a line drawn between any two points on the
leading edge that have a difference in radius of at least 5% of the
blade span is in the range of from -6.degree. and 0.degree., [0013]
where a negative angle indicates that the respective line has an
axial component that is in the same direction as the axial
component of the direction from a trailing edge to the leading edge
of the blade.
[0014] Such a blade may be described as having forward axial
lean.
[0015] Where reference is made to the axial, radial and
circumferential directions, the skilled person will readily
understand this to mean the conventional directions when the fan
blade is assembled as part of a fan stage or is provided in a gas
turbine engine. Viewing the blade along a circumferential direction
may mean viewing the blade in side profile and/or in the meridional
plane and/or projected onto a plane defined by the axial and radial
directions.
[0016] As explained in greater detail elsewhere herein, the present
inventors have understood that a fan blade as described and/or
claimed herein may have reduced susceptibility to flutter compared
with conventional blades, for example because the first (or lowest)
natural frequency mode of the fan blade may have a reduced
torsional content compared to conventional blades. The present
inventors have understood, inter alia, the way in which this
torsional content of the first natural frequency mode may have an
impact on the susceptibility of a fan blade to flutter, and have
developed the blades described and/or claimed herein in order to
reduce this susceptibility.
[0017] The angle formed between the radial direction and a straight
line drawn between the leading edge at the root and at the tip may
be in the range of from -5.degree. and -0.25.degree., for example
-4.degree. and -0.5.degree., for example -3.degree. and
-0.75.degree., for example -2.degree. and -1.degree..
[0018] The angle formed between the radial direction and a line
drawn between any two points on the leading edge that have a
difference in radius of at least 5% of the blade span may be in the
range of from -5.degree. and -0.25.degree., for example -4.degree.
and -0.5.degree., for example -3.degree. and -0.75.degree., for
example -2.degree. and -1.degree..
[0019] The maximum perpendicular distance between any point on the
leading edge and a straight line drawn between the leading edge at
the root and at the tip may be 5% of blade span, for example 4%,
3%, 2%, 1%, 0.5% or 0.1% of the blade span.
[0020] The fan blade may comprise a platform. The fan blade may
comprise a root portion. The root portion may extend between the
platform and the root of the aerofoil portion. Alternatively, the
aerofoil portion may extend directly from the platform, with no
intermediate root portion, such that the root of the aerofoil foil
portion is the root of the fan blade.
[0021] Where the fan blade comprises a root portion, the radial
extent of the root portion may be no more than 15%, for example no
more than 10%, 7%, 5%, 3%, 2% or 1%, of the span of the aerofoil
portion, for example.
[0022] The fan blade may comprise a tip portion that extends at
least radially away from the tip of the aerofoil portion.
Alternatively, the fan blade may comprise no tip portion, such that
the tip of the aerofoil portion is also the tip of the fan
blade.
[0023] Where the fan blade comprises a tip portion, the radial
extent of the tip portion may be no more than 15%, for example no
more than 10%, 7%, 5%, 3%, 2% or 1%, of the span of the aerofoil
portion, for example.
[0024] For any two points on the leading edge of a fan blade as
described and/or claimed herein that are radially closer to the tip
than to the root and have a difference in radius of at least 1%,
for example at least 2%, for example at least 3%, for example at
least 4%, for example at least 5% of the blade span, the axial
position of the radially outer point may be forward of the axial
position of the radially inner point.
[0025] A stacking axis of the aerofoil portion may be defined by a
line joining the centroids of all of the aerofoil segments that are
stacked to form the aerofoil portion. When viewed along a
circumferential direction, the stacking axis may have a forward
lean. For example, the angle formed between the radial direction
and a straight line drawn between the stacking axis at the root and
at the tip may be in the range of from -40.degree. and 0.degree.,
for example -30.degree. and -1.degree., for example -25.degree. and
-2.degree., for example -20.degree. and -3.degree., for example
-15.degree. and -5.degree., for example -10.degree. and -6.degree.,
where a negative angle indicates that the respective line has an
axial component that is in the same direction as the axial
component of the direction from a trailing edge to the leading edge
of the blade. By way of further example, optionally the stacking
axis may have a forward (negative) lean in the radially outer half
of the aerofoil portion, for example only in the radially outer
half of the aerofoil portion.
[0026] The aerofoil portion may have a trailing edge extending from
a root to a tip. When viewed along a circumferential direction, the
trailing edge may have a forward lean. For example, the angle
formed between the radial direction and a straight line drawn
between the trailing edge at the root and at the tip may be in the
range of from -40.degree. and 0.degree., for example -30.degree.
and -1.degree., for example -25.degree. and -2.5.degree., for
example -20.degree. and -5.degree., for example -15.degree. and
-7.5.degree., for example around -10.degree., where a negative
angle indicates that the respective line has an axial component
that is in the same direction as the axial component of the
direction from a trailing edge to the leading edge of the
blade.
[0027] The trailing edge may be shaped such that the forward
(negative) lean angle over a radially outer half of the trailing
edge is greater, for example significantly greater, than the
forward (negative) lean angle over a radially inner half of the
trailing edge. For example, the forward (negative) lean angle over
a radially outer half of the trailing may be at least 1.5 times,
for example at least twice, for example at least 3, 4, 5, 6, 7, 8,
9 or 10 times the forward (negative) lean angle over a radially
inner half of the trailing edge. In some arrangements, the trailing
edge may be radial (including substantially radial) over a radially
inner portion of the blade (or aerofoil portion), for example over
a radially inner 10%, 20%, 30%, 40% or around 50%.
[0028] In any aspect or example of the present disclosure, the
magnitude of the angle formed between the radial direction and a
straight line drawn between the trailing edge at the root and at
the tip may be greater than the magnitude of the angle formed
between the radial direction and a straight line drawn between the
leading edge at the root and at the tip. For example, the angle
formed between the radial direction and a straight line drawn
between the trailing edge at the root and at the tip may have a
higher negative value than that of the of the angle formed between
the radial direction and a straight line drawn between the leading
edge at the root and at the tip. The angle formed between the
radial direction and a straight line drawn between the trailing
edge at the root and at the tip may be negative. The angle formed
between the radial direction and a straight line drawn between the
leading edge at the root and at the tip may be negative.
[0029] Any fan blade and/or aerofoil portion 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, and/or 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.
[0030] According to an aspect, there is provided a fan stage for a
gas turbine engine comprising a plurality of fan blades as
described and/or claimed herein. The fan stage may comprise a hub,
from which the fan blades may extend, for example in a radial
direction.
[0031] Such a hub may be, or may comprise, a fan disc and/or may be
driven by a shaft. The shaft itself may be driven by a turbine of a
gas turbine engine.
[0032] The fan blades may be attached to the hub 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.
[0033] By way of further example, the fan blades maybe formed
integrally with a hub. Such an arrangement may be referred to as a
blisk or a bling. Any suitable method may be used to manufacture
such a blisk or bling. 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.
[0034] By way of further example, the fan blades may be attached to
a hub in a manner that allows their pitch to be varied.
[0035] In any arrangement of fan stage, the ratio of the radius of
the position where the leading edge of one of the fan blades meets
the hub to the outermost radial extent of the leading edge of the
fan blade is less than 0.4, for example less than 0.37, for example
less than 0.35, for example less than 0.33, for example less than
0.3, for example less than 0.25. This may be referred to as the
hub-to-tip ratio and/or may be the same as the ratio of the radius
of the root at the leading edge of the aerofoil portion to the
radius of the tip at the leading edge of the aerofoil portion, for
example where the fan blade is not provided with a root portion or
a tip portion. 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.
[0036] According to an aspect, there is provided a gas turbine
engine comprising at least one fan blade as described and/or
claimed herein and/or a fan stage as described and/or claimed
herein.
[0037] Such a gas turbine engine (which may, of course, be a
turbofan gas turbine engine) may have a specific thrust of less
than 15 lbf/lb/s (or approximately 150 N/Kg/s), for example less
than 12 lbf/lb/s (or approximately 120 N/Kg/s), for example less
than 10 lbf/lb/s (or approximately 110 N/Kg/s or 100 N/Kg/s), for
example less than 9 lbf/lb/s (or approximately 90 N/Kg/s), for
example less than 8.5 lbf/lb/s (or approximately 85 N/Kg/s), for
example less than 8 lbf/lb/s (or approximately 80 N/Kg/s).
[0038] Any gas turbine engine described and/or claimed herein may
have a fan tip loading (dH/U.sub.tip.sup.2) at cruise conditions of
greater than 0.3, for example in the range of from 0.3 to 0.37, for
example 0.32 to 0.36, for example on the order of 0.35 (all units
being JKg.sup.-1K.sup.-1/(ms.sup.-1).sup.2), where dH is the
enthalpy rise across the fan (for example the 1-D average enthalpy
rise of the flow across the fan at cruise conditions), and
U.sub.tip is the velocity of the tip, for example at cruise
conditions, which may be calculated as the rotational speed
multiplied by the tip radius at the leading edge. Cruise may be
defined as the phase between the initial ascent and final descent
of an aircraft to which the engine may be attached. As used herein,
cruise may mean, for example, mid-cruise, i.e. mid-point (for
example in terms of time and/or fuel burn) of a flight (or at least
of the cruise phase of a flight).
[0039] 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 250 cm, for example greater than 260 cm, 270 cm, 280
cm, 290 cm, 300 cm, 310 cm, 320 cm, 330 cm, 340 cm or 350cm.
[0040] 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 10, for example greater than 11, for
example greater than 11.5, for example greater than 12, for example
greater than 13, for example greater than 14, for example greater
than 15. 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.
[0041] 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 at least 170 kN of thrust, for example at
least 180 kN, for example at least 190 kN, for example at least 200
kN, for example at least 250 kN, for example at least 300 kN, for
example at least 350 kN, for example at least 400 kN. The thrust
referred to above may be at standard atmospheric conditions.
[0042] Such a gas turbine engine may be of any suitable form. For
example, the gas turbine engine may be an aero gas turbine engine
for use on aircraft. Such an engine may be a geared turbofan gas
turbine engine, in which the fan stage is driven from a turbine via
a gearbox, in order to reduce (or increase) the rotational speed of
the fan stage compared with the driving turbine stage(s).
[0043] The arrangements of the present disclosure may be
particularly effective in addressing any operability issues
presented by the use of such lower speed fans, such as those driven
via a gearbox.
[0044] The input to such a gearbox may be directly from a core
shaft that connects a turbine to a compressor, 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).
[0045] Any number of fan stages may be provided to an engine. For
example, a gas turbine engine may have a single fan stage, such
that the next downstream rotor stage after the fan is a compressor
rotor stage, for example a compressor rotor stage in the core of
the engine. The skilled person will appreciate that except where
mutually exclusive, a feature described in relation to any one of
the above aspects may be applied to any other aspect. Furthermore,
except where mutually exclusive, any feature described herein may
be applied to any aspect and/or combined with any other feature
described herein.
BRIEF DESCRIPTION OF THE DRAWINGS
[0046] Embodiments will now be described by way of example only,
with reference to the Figures, in which:
[0047] FIG. 1 is a sectional side view of a gas turbine engine on
accordance with the present disclosure;
[0048] FIG. 2 is a radial view of a fan blade according to an
example of the present disclosure;
[0049] FIG. 3 is a side view of a fan blade according to an example
of the present disclosure;
[0050] FIG. 4 is another side view of a fan blade according to an
example of the present disclosure;
[0051] FIG. 5 is a close-up view of a leading edge portion of a fan
blade according to an example of the present disclosure; and
[0052] FIG. 6 is a side view of a fan blade according to an example
of the present disclosure.
DETAILED DESCRIPTION OF THE DISCLOSURE
[0053] With reference to FIG. 1, a gas turbine engine is generally
indicated at 10, having a principal and rotational axis 11. The
engine 10 comprises, in axial flow series, an air intake 12, a
propulsive fan 13, an intermediate pressure compressor 14, a
high-pressure compressor 15, combustion equipment 16, a
high-pressure turbine 17, an intermediate pressure turbine 18, a
low-pressure turbine 19 and an exhaust nozzle 20. A nacelle 21
generally surrounds the engine 10 and defines both the intake 12
and the exhaust nozzle 20.
[0054] The gas turbine engine 10 works in the conventional manner
so that air entering the intake 12 is accelerated by the fan 13 to
produce two air flows: a first air flow into the intermediate
pressure compressor 14 and a second air flow which passes through a
bypass duct 22 to provide propulsive thrust. The intermediate
pressure compressor 14 compresses the air flow directed into it
before delivering that air to the high pressure compressor 15 where
further compression takes place.
[0055] 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 combusted. The resultant hot
combustion products then expand through, and thereby drive the
high, intermediate and low-pressure turbines 17, 18, 19 before
being exhausted through the nozzle 20 to provide additional
propulsive thrust. The high 17, intermediate 18 and low 19 pressure
turbines drive respectively the high pressure compressor 15,
intermediate pressure compressor 14 and fan 13, each by suitable
interconnecting shaft.
[0056] The gas turbine engine 10 and/or the fan stage 13 and/or the
fan blades 100 of the fan stage 13 shown in FIG. 1 may be in
accordance with examples of the present disclosure, aspects of
which are described by way of example only in relation to FIGS. 2
to 6.
[0057] Any gas turbine engine in accordance with the present
disclosure (such as the gas turbine engine 10 of FIG. 1) may, for
example, have a specific thrust in the ranges described herein (for
example less than 10) and/or a fan blade hub to tip ratio in the
ranges described herein and/or a fan tip loading in the ranges
described herein.
[0058] The present disclosure may relate to any suitable gas
turbine engine. For example, other gas turbine engines to which the
present disclosure may be applied may have related or alternative
configurations. By way of example such engines may have an
alternative number of interconnecting shafts (e.g. two) and/or an
alternative number of compressors and/or turbines. Further the
engine may comprise a gearbox provided in the drive train from a
turbine to a compressor and/or fan. The gas turbine engine shown in
FIG. 1 has a mixed flow nozzle 20, meaning that the flow through
the bypass duct 22 and the flow through the core 15, 16, 17, 18, 19
are mixed, or combined, before (or upstream of) the nozzle 20).
However, this is not limiting, and any aspect of the present
disclosure may also, for example, relate to engines 10 having a
split flow nozzle, which may mean that the flow through the bypass
duct 22 has its own nozzle that is separate to and may be radially
outside a core engine 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.
[0059] The geometry of the gas turbine engine 10, and components
thereof, is defined by a conventional axis system, comprising an
axial direction 30 (which is aligned with the rotational axis 11),
a radial direction 40, and a circumferential direction 50 (shown
perpendicular to the page in the FIG. 1 view). The axial, radial
and circumferential directions 30, 40, 50 are mutually
perpendicular.
[0060] The fan stage 13 comprises a plurality of fan blades 100
extending from a hub 200. The fan blades 100 may be defined with
respect to the axial direction 30, radial direction 40, and
circumferential direction 50 shown in FIG. 1 in relation to the gas
turbine engine 10.
[0061] FIG. 2 is a view in a radially inward direction of a fan
blade 100. FIG. 3 is a side view (that is, a view in the
axial-radial plane) of the fan blade 100. The fan blade 100 has an
aerofoil portion 110. The aerofoil portion 110 has a leading edge
120 and a trailing edge 130. The aerofoil portion 110 extends from
a root 140 to a tip 150 in a substantially radial spanwise
direction. The leading edge 110 may be defined as the line defined
by the axially forwardmost points of the aerofoil portion 110 from
its root 140 to its tip 150.
[0062] As mentioned above, the susceptibility of a fan blade 100 to
flutter is at least part dependent on the torsional content of the
lowest natural frequency mode shape (also referred to as the 1F
mode shape)
[0063] This torsional content can be defined at the tip 150 of the
blade 100 by a parameter l/Ch, where l is the distance (for example
the shortest distance) from the leading edge 120 at the tip 150 to
the centre of twist 310 in the mode shape, and Ch is the chord
length at the tip 150 (see FIG. 2). Therefore a relatively higher
value of l/Ch represents a relatively lower torsional content in
the mode shape. The parameter l/Ch may be thought of as describing
the relative motion (for example in a substantially circumferential
direction) between leading edge 120 and trailing edge 130 at the
tip 150 (the motion of the leading edge 120 at the tip 150 usually
being greater than the motion of the trailing edge 130 at the tip
150). In FIG. 2, the reference label 120' indicates the leading
edge 120 in the deformed (or displaced) 1F mode shape, and the
reference label 150' indicates the tip 150 in that deformed (or
displaced) 1F mode shape.
[0064] FIG. 3 shows the nodal point 320 on the leading edge 120 and
the nodal point 330 on the trailing edge 130 in the 1F mode shape,
the nodal points 320, 330 being points on the blade 100 (for
example towards the root 140) which are stationary in the 1F mode
shape. To a first approximation the displacement of the tip 150 at
the leading edge 120 depends on (or at least is affected by) the
distance `a` between the leading edge 120 at the tip 150 and the 1F
nodal point 320 on the leading edge 120. Similarly, to a first
approximation the displacement of the tip 150 at the trailing edge
130 depends on (or at least is affected by) the distance `b`
between the trailing edge 130 at the tip 150 and the 1F nodal point
330 on the trailing edge 130.
[0065] Typically, the distance `a` is greater than the distance
`b`. This may be at least in part because typically the 1F nodal
point 320 on the leading edge 120 is radially inside the 1F nodal
point 330 on the trailing edge 130.
[0066] The blade 100 shown in the Figures by way of example of the
present disclosure may be said to be leant axially forwards, for
example by at least having the line `a` pointing axially forwards,
i.e. to the left in FIG. 3. As shown in the FIG. 3 example, the
line `b` may also point axially forwards, for example at a greater
angle than the line `a`.
[0067] As explained elsewhere herein, the blade geometry described
and/or claimed herein may reduce the susceptibility of the blades
to flutter. For example, and without being limited or bound to a
particular theory, the ratio between lengths `a` and `b` may be
decreased compared with conventional blades, which may result in a
decrease in the ratio between the tip displacement at the leading
edge 120 and the tip displacement at the trailing edge 130 in the
1F mode. The fan blades 100 (and/or aerofoil portions 110) may have
increased l/Ch compared with conventional fan blades, which may
help to reduce the torsional content of the 1F mode shape, and thus
reduce the susceptibility to flutter.
[0068] Various features of an exemplary fan blade 100 will now be
described with reference to FIGS. 4 and 5. It will be appreciated
that these features may be applied alone or in combination, as
defined in the claims. The variables shown in FIGS. 4 and 5 are
explained in the table below, where the term "LE" refers to the
leading edge 120, and the term "TE" refers to the trailing edge
130:
TABLE-US-00001 Point Name Definition A LE root Leading edge 120 at
the root 140 of the aerofoil 110 B TE root Trailing edge 130 at the
root 140 of the aerofoil 110 C LE tip Leading edge 120 at the tip
150 of the aerofoil 110 D TE tip Trailing edge 130 at the tip 150
of the aerofoil 110 P1, P2 LE points Points on leading edge 120
with radii that are at least 5% of the aerofoil span apart r P 1 -
r P 2 r C - r A .gtoreq. 0.05 ; r P 2 .gtoreq. r P 1 ##EQU00001## E
Point on LE Point on leading edge 120 with maximum perpendicular
distance to the line AC H Point on AC e Distance EH e = ( x H - x E
) 2 + ( r H - r E ) 2 ##EQU00002## Alpha .alpha. LE global slope
.varies. = 180 .pi. atan ( x c - x A r c - r A ) ##EQU00003##
alpha(P1, P2) LE local slope .varies. ( P 1 , P 2 ) = 180 .pi. atan
( x P 2 - x P 1 r P 2 - r P 1 ) ##EQU00004## e % LE "straightness"
e % = e r C - r A 100 ##EQU00005## Span Aerofoil Span Difference in
the radius of the leading edge 120 at the root 140 and at the tip
150: r.sub.C - r.sub.A Note in the above table, that "x" refers to
a position in the axial direction 30 and "r" refers to a position
in the radial direction 40.
[0069] The global slope .varies. of the aerofoil portion 110 of fan
blades 100 as described and/or claimed herein, such as that shown
by way of example in FIG. 4, may be in the range of from -6.degree.
and -0.2.degree., for example within any of the ranges defined
elsewhere herein. In this regard, the global slope .varies. may
represent the angle formed between the radial direction and a
straight line AC drawn between the leading edge point A at the root
140 and the leading edge point C at the tip 150.
[0070] The local slope .varies. (P1, P2) of the aerofoil portion
110 of fan blades 100 as described and/or claimed herein, such as
that shown by way of example in FIG. 4, may be in the range of from
-6.degree. and 0.degree., for example within any of the ranges
defined elsewhere herein. In this regard, the local slope .varies.
(P1, P2) may represent the angle formed between the radial
direction and a line drawn between any two points on the leading
edge that have a difference in radius of at least 5% of the blade
span.
[0071] The relationship between `E` and `H` as defined in the table
above is seen most easily in FIG. 5. The distance `e` between the
points `E` and `H` may be said to represent the maximum
perpendicular distance between any point on the leading edge and a
straight line drawn between the leading edge at the root. As a
percentage (`e%`)the distance `e` of the aerofoil portion 110 of
fan blades 100 as described and/or claimed herein may be less than
5%, for example less than 2% (or any other range as described
and/or claimed herein) of the span of the aerofoil portion 110.
[0072] The trailing edge 130 of the aerofoil portion 110 may also
define a global slope .beta.. The global slope .beta. of the
trailing edge 130 of fan blades 100 may be in the range of from
-40.degree. and 0.degree., for example -30.degree. and -1.degree.,
for example -25.degree. and -2.5.degree., for example -20.degree.
and -5.degree., for example -15.degree. and -7.5.degree., for
example around -10.degree.. In this regard, the global slope .beta.
of the trailing edge 130 may represent the angle between the radial
direction and a straight line I drawn between a point B on the
trailing edge 130 at the root 140 and a point D on the trailing
edge 130 at the tip 150.
[0073] The fan blade 100 comprises a platform 160. The aerofoil
portion 110 may extend directly from the platform 160, as in the
FIG. 4 example. Alternatively, as shown by way of example in FIG.
6, a fan blade 100 may have a root portion 170. The root portion
170 may be said to extend between the platform 160 and the root 140
of the aerofoil portion 110. The radial extent of the root portion
170 may be no more than 7%, for example no more than 5%, of the
span of the aerofoil portion 110.
[0074] Also as shown by way of example in FIG. 6, the fan blade 100
may comprise a tip portion 180. The tip portion 180 may be said to
extend from the tip 150 of the aerofoil portion 110. The radial
extent of the tip portion 180 may be no more than 5% of the span of
the aerofoil portion 110.
[0075] The fan blade 100 may be attached to the hub 200 in any
desired manner. For example, the fan blade 100 may comprise a
fixture 190 such as that shown by way of example in FIG. 6 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.
[0076] Alternatively, the fan blade 100 and the hub 200 may be
formed as a unitary part, with no mechanical and/or releasable
connections, so as to form a unitary fan stage 13. Such a unitary
fan stage 13 may be referred to as a "blisk". Such a unitary fan
stage 13 may be manufactured in any suitable manner, for example by
machining and/or by linear friction welding the fan blades 100 to
the hub 200, or at least linear friction welding the aerofoil
portions 110 to a hub 200 that includes radially inner stub
portions of the fan blades 100.
[0077] The hub to tip ratio, which may have a value as indicated
elsewhere herein, may be defined as the radius of the leading edge
120 at the root 140 (which may itself be referred to as a hub) of
the aerofoil 110 (point A) divided by the radius of the leading
edge 120 at the tip 150 of the aerofoil 110 (point B), i.e.
r.sub.A/r.sub.B.
[0078] 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.
* * * * *