STATEMENT REGARDING FEDERALLY SPONSORED RESEARCH OR
DEVELOPMENT
This invention was made with Government support under DTFAWA-I O-C00046
awarded by United States Department of Transportation Federal Aviation
Administration. The Government has certain rights in this invention.
BACKGROUND
Embodiments presented herein relate generally to aerodynamic surfaces
configured for noise reduction, and more specifically to configuration of a tip portion
on an aerodynamic surface, such as an airfoil, for noise reduction.
At least some known machines including aerodynamic surfaces such as, but
not limited to, wind turbines, aircraft airframes, aircraft engines, gas turbine engines
and steam turbine engines, include a plurality of stationary and/or rotating airfoils
which are subject to impinging wakes and vortices generated from an upstream
object, such as an upstream bladerow or an input unsteady airflow. The upstream
generated wakes and vortices are channeled downstream where they may impinge on
the leading edge of downstream airfoils. In one instance, the wake flow impingement
from upstream object on the downstream airfoils moving relative to each other is a
dominant source of aerodynamic noise and aeromechanical loading generated in
turbomachinery applications.
Of particular interest are unducted, contra-rotating engines which have been
developed such as the GE 36 engine, frequently referred to as an unducted fan (UDF)
or propfan engine. In some known unducted contra-rotating engines, noise may be
generated by an upstream rotating airfoil's wake impinging on a leading edge of a
contra-rotating airfoil located downstream. In other known instances, noise may be
generated by an upstream stator component's wake impinging on a leading edge of a
rotating airfoil downstream from the component.
2
Noise generated by aircraft engines may be constrained by international and
local regulations, thereby creating a need to balance fuel efficiency and emissions
with noise pollution. A dominant source of aerodynamic noise and aeromechanical
loading generated in turbomachinery applications is the interaction of wakes from
upstream bladerows on downstream bladerows or vanes moving relative to each other.
As previously indicated, examples include fan wakes and vortices interacting with
downstream contra-rotating fan blades, whereby open rotor noise may be generated
by the forward-aft rotor interaction. In addition, of interest is turbomachinery noise
from stator vane wakes impinging on downstream rotor blades. The impinging wake
flow on the airfoil's leading edge may result in an increase in noise radiated from the
turbomachinery, as well as a potential increase in aeromechanical loading on the
bladerow. At least some known methods of reducing the noise generated by these
unsteady wake flows impinging on airfoils include increasing the distance between
the upstream object or airfoil and the downstream airfoil. This increased distance
mixes the wake flow and thus reduces the amplitude of the wake flow forcing
unsteady motion of the tip vortex of the downstream airfoil. However, increasing the
distance between an upstream object, such as another airfoil, and the downstream
airfoil may increases the size, weight, and cost of the engine, and thereby reduce the
efficiency and performance of the engine.
BRIEF DESCRIPTION
In accordance with one exemplary embodiment, an airfoil is disclosed. The
airfoil comprising: a root portion and a tip portion, wherein the tip portion is
configured extending radially outward from the root portion; a suction side and a
pressure side coupled together at a leading edge and a trailing edge spaced chord-wise
and downstream from the leading edge; a plurality of chord sections having a chord
length and extending in a chord-wise direction between the leading edge and the
trailing edge of the airfoil and spaced apart along a length of the airfoil in a span-wise
direction between the root portion and the tip portion; and a tip profile defining a
reducing slope extending from the leading edge at the tip portion along at least a
portion of the chord length, wherein the tip profile is configured to reduce high
unsteady pressure near the tip portion of the airfoil.
3
In accordance with another exemplary embodiment, a fan assembly is
disclosed. The fan assembly comprising: a disk; and a plurality of fan blades coupled
to the disk, each blade of the plurality of fan blades comprising: an airfoil portion
comprising a suction side and a pressure side coupled together at a leading edge and a
trailing edge spaced chord-wise and downstream from the leading edge; a plurality of
chord sections having a chord length and extending in a chord-wise direction between
the leading edge and the trailing edge of the airfoil and spaced apart along a length of
the airfoil in a span-wise direction between the root portion and the tip portion; and a
tip profile defining a reducing slope extending from the leading edge at the tip portion
along at least a portion of the chord length, wherein the tip profile is configured to
reduce high unsteady pressure near the tip portion of the airfoil.
In accordance with another exemplary embodiment, an unducted contrarotating
fan engine is disclosed. The unducted contra-rotating fan engine comprising:
an unducted fan section including a first fan blade row and a second fan blade row
axially spaced aftward from the first fan blade row, the second fan blade row
including a plurality o~ airfoils, each airfoil comprising: a root portion and a tip
portion, wherein the tip portion is configured extending radially outward from the root
portion; a suction side and a pressure side coupled together at a leading edge and a
trailing edge spaced chord-wise and downstream from the leading edge; a plurality of
chord sections having a chord length and extending in a chord-wise direction between
the leading edge and the trailing edge of the airfoil and spaced apart along a length of
the airfoil in a span-wise direction between the root portion and the tip portion; and a
tip profile defining a reducing slope extending from the leading edge at the tip portion
along at least a portion of the chord length, wherein the tip profile is configured to
reduce high unsteady pressure near the tip portion of the airfoil.
In accordance with another exemplary embodiment, a method of fabricating
an airfoil is disclosed. The method of fabricating an airfoil comprising: fabricating at
least one airfoil including a suction side and a pressure side coupled together at a
leading edge and a trailing edge spaced chord-wise and downstream from the leading
edge; wherein the airfoil includes a plurality of chord sections having a chord length
and extending in a chord-wise direction between the leading edge and the trailing
4
edge of the airfoil and spaced apart along a length of the airfoil in a span-wise
direction between the root portion and the tip portion, said tip portion comprises: a tip
profile defining a reducing slope with no slope discontinuity extending from the
leading edge at the tip portion along at least a portion of the chord length, wherein the
tip profile is configured to reduce high unsteady pressure near the tip portion of the
airfoil.
DRAWINGS
The above and other aspects, features, and advantages of the present
disclosure will become more apparent in light of the subsequent detailed description
when taken in conjunction with the accompanying drawings in which:
FIG. I is a perspective view showing an aircraft supporting an engine
including airfoils having a tip profile in accordance with an embodiment;
FIG. 2 is an enlarged side view of the engine shown in FIG. I;
FIG. 3 is a schematic cross-section taken through line 3-3 of FIG. 2,
illustrating an unducted contra-rotating fan engine including airfoils having a tip
profile in accordance with an embodiment;
FIG. 4 is a perspective view of a prior art airfoil showing a standard tip
profile;
FIG. S is an enlarged perspective view of a portion of the prior art airfoil of
FIG. 4, showing the standard tip profile;
FIG. 6 is a perspective view of an exemplary airfoil of FIG. 2 showing a tip
profile according to an embodiment;
FIG. 7 is an enlarged perspective view of the exemplary airfoil of FIG. 6
showing a tip profile according to an embodiment; and
FIG. 8 is an enlarged perspective view of a portion of the airfoil of FIG. 7,
showing the tip profile according to an embodiment.
S
DETAILED DESCRIPTION
Generally provided are exemplary apparatus and methods for fabricating an
airfoil such as, but not limited to, for use in a device incorporating aerodynamic
surfaces, and more particularly for use in a rotary device, such as, but not limited to,
an open rotor propulsion system. The embodiments described herein are not limiting,
but rather are exemplary only. It should be understood that the exemplary apparatus
and methods for fabricating an airfoil disclosed herein may apply to any type of airfoil
or aerodynamic surface, such as, but not limited to, fan blades, rotor blades, ducted
fan blades, unducted fan blades, turbine engine, wind turbines, aircraft wing high-lift
systems and/or aircraft structures. More specifically, the exemplary apparatus and
methods for fabricating an airfoil disclosed herein may apply to any airfoil, or
aerodynamic surface, that is subject to impinging wakes and vortices generated
upstream of the airfoil.
Although the embodiments described herein are described in connection with
an open rotor propulsion system, also referred to herein as an unducted contra-rotating
front fan high bypass ratio engine, or UDF, it should be apparent to those skilled in
the art that, with appropriate modification, the apparatus and methods can be suitable
for any device including airfoils that are subject to impinging wakes and vortices
generated upstream of the airfoil and for which tip vortex noise related to self- and
gust-interaction is of interest.
Referring now to FIG. 1, there is shown an aircraft 10 supporting an engine
assembly 12 in accordance with one embodiment. The aircraft 10 is shown having a
pair of swept back wings 14 and 16. Mounted on wing 14 is the engine assembly 12,
and more particularly in an embodiment, an unducted contra-rotating front fan high
bypass ration engine assembly, also referred to herein as an open-rotor propulsion
system. The pylon configuration shown is not intended to be limiting and that
additional pylon configurations are anticipated (e.g. pusher configurations and puller
configurations) and that the disclosed tip profile is not limited by engine architecture.
6
FIG. 2 shows an enlarged side view of the engine assembly 12 of FIG. 1. FIG.
3 illustrates a sectional view taken through line 3-3 of the engine assembly 12 of FIG.
2 according to an embodiment wherein like parts are identically referenced. Engine
assembly 12 includes a longitudinal center line axis 18 that extends through the
engine assembly 12 from front to back (from left to right on FIGs. 2 and 3). Flow
through the illustrated exemplary engine is generally from front to back. The
direction parallel to the center line axis 18 toward the front of the engine and away
from the back of the engine will be referred to herein as the "upstream" direction 20,
while the opposite direction parallel to the center line axis 18 will be referred to
herein as the "downstream" direction 22.
The engine assembly 12 has an outer shell, or an outer casing 24 disposed coaxially
about center line axis 18. Outer casing is conventionally referred to as a
nacelle.
Engine assembly 12 also includes a gas generator referred to as core engine
26. Such core engine includes a compressor 28, a combustor 30 and a high pressure
turbine 32, either singular or multiple stages.
At the forward part of the engine 12, there is provided a front fan section 44.
Fan section 44 includes a first fan blade row 46 connected to a forward end of an
inner contra-rotating shaft 48 which extends between the power turbine 38 and the fan
section 44. Front fan section 44 includes a second fan blade row 50 connected to the
forward end of an outer drive shaft 52 also connected between the power turbine 38
and the fan section 44. Each of the first and second fan blade rows 46 and 50
comprises a plurality of circumferentially spaced airfoils 54, or fan blades. Fan blade
rows 46 and 50 are contra-rotating which provides a higher propulsive efficiency. It
should be appreciated that the contra-rotating fan blade row 50 serves to remove the
swirl on the circumferential component of air imparted by the contra-rotating fan
blade row 46.
An important feature of the engine design is the positioning and design of the
fan blade rows 46 and 50. Initially, in order to reduce the noise resulting from the fan
7
blade rows 46 and 50, sufficient spacing must be provided between the fan blade
rows. As described below, the airfoils 54 in blade row 50 are further configured to
include a tip profile as described herein, to minimize tip vortex noise related to selfand
gust-interaction. The airfoils 54 in blade row 46 may also be configured to
include a tip profile as described herein, and it is understood that descriptions
henceforth for the novel tip profile described in this disclosure applied to the
downstream blade row are potentially equally applicable to the upstream blade row.
FIG. 4 is a perspective view of one embodiment of prior art fan blade 60,
generally similar to a fan blade that may be used in an engine assembly, generally
similar to the engine assembly 12 of FIG. 1-3. FIG. 5 is an enlarged view ofa portion
of the prior art fan blade 60, as indicated. In the illustrated embodiment, the fan blade
60 includes an airfoil portion 62, a tip portion 64, and a root portion 66. Alternatively,
the airfoil portion 62 may be used with, but not limited to, rotor blades, and/or turbine
blades. As illustrated, tip portion 64 of fan blade 60 is configured as a substantially
straight, constant sloped line 68 defined by the circumferentially-averaged streamline
contraction angle at cruise or max-climb operating condition (i.e. high flight velocity,
Mach no. -0.7-0.8). At takeoff and approach, the contraction angle is much higher,
causing a tip vortex to significantly influence both the steady and unsteady blade
surface pressure on a suction-side 70 of the airfoil portion 62. This creates a strong
localized sound source that adversely affects community noise. This unsteady
interaction noise source contributing to community noise may be dominated by the
open rotor tip vortices, their sensitivity to flow unsteadiness and their proximity to
nearby blade surfaces.
Turning now to FIGs. 6-8 illustrated is an exemplary fan blade for reduced
community noise according to an embodiment. In particular, FIG. 6 is a perspective
view of an embodiment of an aerodynamic surface, and more particularly the fan
blade embodying an airfoil including the tip profile as disclosed herein. FIG. 7 is an
enlarged perspective view of the airfoil of FIG. 6 wherein like parts are identically
referenced. FIG. 8 is an enlarged view of a portion of the airfoil, as indicated wherein
like parts are identically referenced. More particularly, illustrated is a fan blade 70,
generally similar to the fan blade 50 of FIGs. 2 and 3 that may be used in an engine
8
assembly, generally similar to the engine assembly 12 of FIGs. 1-3. In a preferred
embodiment, fan blade 70 may reside in an aft positioned bladerow, a forward
positioned bladerow, or both forward and aft positioned bladerows, similar to
bladerows 46 and 50 of FIGs. 2 and 3. In an embodiment, the fan blade 70 includes
an airfoil 72, a platform 74, and a root portion 76. Alternatively, the airfoil 72 may be
used with, but not limited to, rotor blades, and/or turbine blades. In an embodiment,
the root portion 76 includes an integral dovetail 78 that enables the airfoil 72 to be
mounted to a disk, such as a fan rotor disk. The airfoil 72 includes a first contoured
sidewall 80 and a second contoured sidewall 82. Specifically, in an embodiment, the
first contoured sidewall 80 defines a suction side 81 of the airfoil 72, and the second
contoured sidewall 82 defines a pressure side 83 of the airfoil 72. The sidewalls 80
and 82 are coupled together at a leading edge 84 and at an axially-spaced trailing edge
86. The trailing edge 86 is spaced chord-wise and downstream from the leading edge
84. The airfoil 72 includes a thickness measured between the pressure side 83 and the
suction side 8I extending from the leading edge 84 to the trailing edge 86, whereby
the airfoil thickness varies in a span-wise direction. The pressure side 83 and the
suction side 8I, and more. particularly first contoured sidewall 80 and second
contoured sidewall 82, respectively, each extend longitudinally, or radially outward,
from the root portion 76 to a tip portion 88. Alternatively, the airfoil 72 may have any
conventional form, with or without the dovetail 78 or platform portion 74. For
example, the airfoil 72 may be formed integrally with a rotor disk in a blisk-type
configuration that does not include the dovetail 78 and the platform portion 74.
In an embodiment, the airfoil 72 includes a tip 98 defining a tip profile 100 at
a tip portion 88. The tip profile 100 is defined by an increased radial angle 94 in a
front portion of the airfoil 72, near the leading edge 84. The increased radial angle 94
alters the shear layer development feeding into a tip vortex created at the tip portion
88 and reduces the magnitude of unsteady pressure on a surface of the suction side 81
near the tip portion 88.
Known aft rotor tip profiles may be defined relative to a streamline contraction
angle at cruise or max climb, i.e. high flight velocity. At takeoff, with a low flight
Mach number, the streamline contraction angle is higher. This causes the tip vortex to
9
influence the surface pressure (steady /unsteady) at the suction side of the tip portion
significantly and creates a very localized and strong noise source. The novel tip
profile 100 disclosed herein for an airfoil, such as airfoil 72, enables a substantial
reduction in noise associated with aft tip vortex/gust interaction while limiting the
aerodynamic impact to be effectively neutral in fan aerodynamic efficiency.
As illustrated in FIG. 6, tip profile 100 is defined by the tip 98 wherein a curve
having a reducing slope defines the tip profile. More particularly, a first portion 102
of the tip profile 100 is located near the leading edge extending generally chord-wise
along at least a portion of a chord length 94 and defining a first slope 104. A second
portion 106 of the tip profile 100 is located adjacent the first portion 102 extending
generally chord-wise between the first portion 102 to the trailing edge 86 ofthe airfoil
72 and defining a second slope 108. The tip profile 100 is configured wherein the
first slope 104 is greater than the second slope 108, thereby defining a reducing slope
tip profile 100. In an embodiment, the first portion 102 and the second portion 106
are defined having no slope discontinuity to form a smooth curve profile. In an
embodiment, the first portion 102 of the tip profile having the first slope 104 extends
generally chord-wise from the leading edge 84 to approximately 25% of the chord
length 94. Thus, the second portion 106 of th~ tip profile having the second slope 108
extends generally chord-wise extending from the first portion 102 to the trailing edge
86, thus approximately 75% of the remaining chord length 94. In an alternate
embodiment, the first portion 102 of the tip profile having the first slope 104 may
extend less than 25% of the chord length 94 from the leading edge 84, and thus the
second portion 106 of the tip profile having the second slope 108 may extend greater
than 75% of the chord length 94 from the first portion 102 to the trailing edge 86. In
yet another alternate embodiment, the first portion 102 of the tip profile having the
first slope 104 may extend more than 25% of the chord length 94 from the leading
edge 84, and thus the second portion 106 of the tip profile having the second slope
108 may extend less than 75% of the chord length 94 from the first portion 102 to the
trailing edge 86. The slope configurations shown are not intended to be limiting and
additional slope configurations wherein a plurality of slopes, with no slope
discontinuity define a reducing slope are anticipated by this disclosure.
10
Detennination and optimization of the change in slope is dependent on the engine
application and fan design, and is affected by differences between design (e.g., cruise)
and takeoff flight conditions, in particular, the changes in thrust, flight velocity and
fan rotation rate, i.e., fan advance ratio. The chordwise location of the change in
slope is affected by the blade design (e.g., sweep), mean aerodynamic loading, etc.,
and the effects these have on the strength and distribution of the tip vorticity. Detailed
implementation and optimization of this novel tip profile to reduce noise while
simultaneously minimizing aerodynamic perfonnance penalty is accomplished using
detailed computational simulations of the aerodynamic flow and blade unsteady
surface pressure resulting from its unsteady interaction with an upstream unsteady
disturbance.
The tip profile 100 reduces the open rotor noise and aeromechanicalloading of
impinging wakes and vortices upon an aft positioned fan blade airfoil, such as airfoil
72. More specifically, the tip profile 100 provides for a reduction in the blade
unsteady response to its own vortex pulsating and oscillating under the action of
incident flow disturbances from upstream. As previously stated, of particular interests
is a reduction in fan tone noise emanating from unducted fan (or open rotor)
propulsion systems. The novel tip profile enables a reduction in open rotor noise and
may provide an effective alternative to other noise designs/technologies that require
undesirable perfonnance compromise.
Further disclosed is a method of fabricating an airfoil. The method includes
fabricating at least one airfoil including a root portion, a tip portion, a suction side and
a pressure side coupled together at a leading edge and at a trailing edge spaced chordwise
and downstream from the leading edge. The airfoil includes a plurality of chord
sections having a chord length and extending in a chord-wise direction between the
leading edge and the trailing edge of the airfoil and spaced apart along a length of the
airfoil in a span-wise direction between the root portion and the tip portion. The tip
portion comprises: a tip profile defining a reducing slope with no slope discontinuity
extending from the leading edge at the tip portion along at least a portion of the chord
11
length, wherein the tip profile is configured to reduce high unsteady pressure near the
tip portion of the airfoil.
An airfoil tip portion configured in this manner addresses the unsteady
ae'rodynamic and aeroacoustic response of a blade to a relative unsteady incoming
flow disturbance. More specifically, the airfoil tip portion configured as described
herein facilitates a reduction in unsteady airfoil response of the wake flow impinging
on the tip of the airfoil such that the noise and aeromechanical loading are facilitated
to be reduced. The reduction in noise resulting from a tip vortex oscillating in
response to an upstream gust and thereby generating high unsteady pressure
fluctuations at the airfoil tip portion may facilitate engine system performance
improvements such as reducing the axial distance necessary between the airfoils and
upstream components. As a result, engine efficiency and performance are facilitated
to be improved in comparison to engines using standard airfoils without a tip profile
defined on a tip portion at least one airfoil. In addition, the reduction in radiated noise
and aeromechanicalloading are achieved without an increase in blade or vane weight,
without substantially decreasing aerodynamic performance, and without any
otherwise impact on the overall engine system (length, weight, structure, etc.). In an
embodiment, the tip profile design disclosed herein may allow for a change in engine
design that would otherwise in some manner increase noise (e.g., reduced fan-fan
axial separation distance, reduced fan diameter, increased fan tip speed, etc.) but
allow for maintenance of target noise levels while gaining overall system
performance.
Exemplary embodiments of airfoils including fan blades are described above
in detail. The airfoils are not limited to the specific embodiments described herein, but
rather, may be applied to any type of airfoil that are subjected to impinging wakes,
vortices, and turbulence from an upstream object, such as a fan blade, stator, airframe,
or an unsteady fluid flow. The airfoils described herein may be used in combination
with other blade system components with other engines.
While the disclosure has been illustrated and described in typical
embodiments, it is not intended to be limited to the details shown, since various
12
modifications and substitutions can be made without departing in any way from the
spirit of the present disclosure. As such, further modifications and equivalents of the
disclosure herein disclosed may occur to persons skilled in the art using no more than
routine experimentation, and all such modifications and equivalents are believed to be
within the spirit and scope of the disclosure as defined by the subsequent claims.
13
10 aircraft
12 engine assembly
14 wing
16 wing
18 longitudinal center line axis
20 upstream direction
22 downstream direction
24 outer shell/casing
26 core engine
28 compressor
30 combustor
32 high pressure turbine
34 34A/34B drive shafts
36 diffuser section
38 power turbine
40 first turbine blade rows
42 second turbine blade rows
44 front fan section
46 first fan blade row
48 inner counterrotating shaft
50 second fan blade row
52 outer drive shaft
54 airfoils
56 pylon
58
60 prior art fan blade
62 airfoil portion
64 tip portion
66 root portion
68 sloped line
14
70 fan blade
72 airfoil portion
74 platform
76 root portion
78 integral dovetail
80 first contoured sidewall
81 pressure side
82 second contoured sidewall
83 suction side
84 leading edge
86 trailing edge
88 tip portion
90 a camber and thickness variation initiation line
92 chord-wise distance
94 chord length
96 chord sections
98 tip cut
100 tip cut profile
102 first portion
104 first slope
106 second portion
108 second slope
110
112
114
116
118
120
122
124
126
15
128
130
132
134
136
138
140
142
16

WE CLAIM:
1. An airfoil (62) comprising:
a root portion (76) and a tip portion (88), wherein the tip portion (88) is
configured extending radially outward from the root portion (76);
a suction side (83) and a pressure side (81) coupled together at a leading edge
(84) and a trailing edge (86) spaced chord-wise and downstream from the
leading edge (84);
a plurality of chord sections (96) having a chord length (94) and extending in a
chord-wise direction between the leading edge (84) and the trailing edge (86)
of the airfoil (62) and spaced apart along a length of the airfoil (62) in a spanwise
direction between the root portion (76) and the tip portion (88); and
a tip profile (100) defining a reducing slope extending from the leading edge
(84) at the tip portion (88) along at least a portion of the chord length (94),
wherein the tip profile (100) is configured to reduce high unsteady pressure
near the tip portion (100) of the airfoil (62).
2. An airfoil in accordance with Claim 1, further including a tip profile
(100) defining a first portion (102) having a first reducing slope (104)
extending from the leading edge (84) at the tip portion (88) along at least a
portion of the chord length (94) and at least one additional portion (106)
having a reducing slope (108) extending from the first portion (102) toward
the trailing edge (86) along at least a portion of the chord length (94), wherein
the first reducing slope (104) is greater than the reducing slope (108) of the at
least one additional portion (106).
3. An airfoil in accordance with Claim 2, wherein the first portion (102)
having a first reducing slope (104) extends from the tip portion (88) at the
leading edge (84) in a chord-wise direction at least 25% of the chord length
(94) of the airfoil (62).
4. An airfoil in accordance with Claim 2, wherein the first portion (102)
having a first reducing slope (104) extends from the tip portion (88) at the
17
leading edge (84) in a chord-wise direction less than 25% of the chord length
(94) of the airfoil (62).
5. An airfoil in accordance with Claim 2, further including a tip profile
(100) defining a first portion (102) having a first reducing slope (104)
extending from the leading edge (84) at the tip portion (88) along at least a
portion of the chord length (94) and a second portion (106) having a second
reducing slope (108) extending from the first portion (102) to the trailing edge
(86) along at least a portion of the chord length (94), wherein the first reducing
slope (104) is greater than the second reducing slope (108).
6. An airfoil in accordance with Claim 1, wherein the airfoil (62) is one
of a ducted or unducted fan blade, rotor blade, wind turbine biade, or an
aircraft aerodynamic surface.
7. A fan assembly comprising:
a disk (44); and
a plurality of fan blades (46, 50) coupled to the disk (44), each blade of the
plurality of fan blades (46, 50) comprising:
an airfoil (72) comprising a root portion (76), a tip portion (88), a
suction side (83) and a pressure side (81) coupled together at a leading edge
(84) and at a trailing edge (86) spaced chord-wise and downstream from the
leading edge; (84)
a plurality of chord sections (96) having a chord length «4) and
extending in a chord-wise direction between the leading edge (84) and
the trailing edge (86) of the airfoil (72) and spaced apart along a length
of the airfoil (72) in a span-wise direction between the root portion
(76) and the tip portion (88); and
a tip profile (100) defining a reducing slope (104) extending from the
leading edge (84) at the tip portion (88) along at least a portion of the
chord length (94),
wherein the tip profile (100) is configured to reduce high unsteady
pressure near the tip portion (99) of the airfoil(72).
18
8. A fan assembly in accordance with Claim 7, wherein the airfoil (72) is
configured to facilitate a reduction in noise associated with gust/tip vortex
interaction.
9. An unducted contra-rotating fan engine comprising:
an unducted fan section (44) including a first fan blade row (46) and a second
fan blade row (50) axially spaced aftward from the first fan blade row (46), the
second fan blade row (50) including a plurality of airfoils (72), each airfoil
comprising:
a root portion (76) and a tip portion (88), wherein the tip portion (88) is
configured extending radially outward from the root portion (76);
a suction side (83) and a pressure side (81) coupled together at a
leading edge (84) and a trailing edge (86) spaced chord-wise and
downstream from the leading edge (84);
a plurality of chord sections (96) having a chord length (94) and
extending in a chord-wise direction between the leading edge (84) and
the trailing edge (86) of the airfoil (72) and spaced apart along a length
of the airfoil (72) in a span-wise direction between the root portion
(76) and the tip portion (88); and
a tip profile (100) defining a reducing slope (104) extending from the
leading edge (84) at the tip portion (100) along at least a portion of the
chord length (94),
wherein the tip profile (100) is configured to reduce high unsteady
pressure near the tip portion (100) ofthe airfoil (72).
10. A method of fabricating an airfoil, said method comprising:
fabricating at least one airfoil (72) including a root portion (76), a tip portion
(88), a suction side (83) and a pressure side (81) coupled together at a leading
edge (84) and a trailing edge (86) spaced chord-wise and downstream from the
leading edge (84);
wherein the airfoil (72) includes a plurality of chord sections (96) having a
chord length (94) and extending in a chord-wise direction between the leading
19
edge (84) and the trailing edge (86) of the airfoil (72) and spaced apart along a
length of the airfoil (72) in a span-wise direction between the root portion (76)
and the tip portion (88), said tip portion (88) comprising:
a tip profile (100) defining a reducing slope (104) extending from the
leading edge (84) at the tip portion (88) along at least a portion of the
chord length (94),
wherein the tip profile (100) is configured to reduce high unsteady
pressure near the tip portion (l00) of the airfoil (72).
~~4-0~
MANISHA SINGH NAIR
Agent for the Applicant [IN/PA-740]
LEX ORBIS
Intellectual Property Practice
709/710, Tolstoy House,
15-17, Tolstoy Marg,
New Delhi-llOOOI