U.S. patent application number 14/079688 was filed with the patent office on 2015-02-12 for geared turbofan with fan blades designed to achieve laminar flow.
This patent application is currently assigned to United Technologies Corporation. The applicant listed for this patent is United Technologies Corporation. Invention is credited to Frederick M. Schwarz, Thomas G. Tillman, Edwin M. Worth.
Application Number | 20150044052 14/079688 |
Document ID | / |
Family ID | 50731819 |
Filed Date | 2015-02-12 |
United States Patent
Application |
20150044052 |
Kind Code |
A1 |
Worth; Edwin M. ; et
al. |
February 12, 2015 |
Geared Turbofan With Fan Blades Designed To Achieve Laminar
Flow
Abstract
A fan blade comprises a main body having an airfoil extending
between a leading edge and a trailing edge. The fan blade has at
least one of a channel closed by a cover, and an end cap covering
at least one of the leading and trailing edges. At least one of a
cover and an end cap has a pair of opposed ends. A step is defined
extending from at least one of a suction wall and a pressure wall
of the airfoil, to an outer surface of the one of a cover and an
end cap at one of the opposed ends, and the step being less than or
equal to about 0.010 inch (0.0254 centimeter) in dimension.
Inventors: |
Worth; Edwin M.;
(Northville, MI) ; Tillman; Thomas G.; (West
Hartford, CT) ; Schwarz; Frederick M.; (Glastonbury,
CT) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
United Technologies Corporation |
Hartford |
CT |
US |
|
|
Assignee: |
United Technologies
Corporation
Harford
CT
|
Family ID: |
50731819 |
Appl. No.: |
14/079688 |
Filed: |
November 14, 2013 |
Related U.S. Patent Documents
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Application
Number |
Filing Date |
Patent Number |
|
|
61727786 |
Nov 19, 2012 |
|
|
|
61884295 |
Sep 30, 2013 |
|
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Current U.S.
Class: |
416/212R ;
29/889.7; 416/232 |
Current CPC
Class: |
F05D 2220/36 20130101;
Y02T 50/60 20130101; F04D 29/023 20130101; F05D 2250/181 20130101;
F05D 2240/31 20130101; F05D 2300/174 20130101; F01D 5/18 20130101;
F05D 2300/173 20130101; F05D 2250/311 20130101; F01D 5/225
20130101; F01D 5/141 20130101; F05D 2250/321 20130101; F05D
2250/322 20130101; F05D 2250/62 20130101; F05D 2300/516 20130101;
F01D 5/145 20130101; F02C 7/36 20130101; F01D 5/28 20130101; F04D
29/324 20130101; Y10T 29/49336 20150115; F01D 5/147 20130101 |
Class at
Publication: |
416/212.R ;
416/232; 29/889.7 |
International
Class: |
F01D 5/18 20060101
F01D005/18; F01D 5/22 20060101 F01D005/22 |
Claims
1. A fan blade comprising: a main body having an airfoil extending
between a leading edge and a trailing edge and the fan blade having
at least one of: (a) a channel closed by a cover, and (b) an end
cap covering at least one of the leading and trailing edges; and
the at least one of the cover and the end cap having a pair of
opposed ends, and a step defined by at least one of a suction wall
and a pressure wall of the airfoil, to an outer surface of said one
of the cover and the end cap at one of said opposed ends, and said
step being less than or equal to about 0.010 inch (0.0254
centimeter) in dimension.
2. The fan blade as set forth in claim 1, wherein said main body
including both said cover and said end cap, the end cap at said
leading edge, wherein there are a plurality of said step defined at
each of said opposed ends of said cover on one of said suction wall
and said pressure wall, and wherein said steps are defined at each
of said opposed ends of said end cap on both said suction and
pressure walls, and wherein all of said step dimensions are less
than or equal to about 0.010 inch (0.0254 centimeter).
3. The fan blade as set forth in claim 2, wherein an outer surface
of said fan blade having a surface roughness and said surface
roughness having a root means square value of less than about
60.times.10.sup.-6 inch on at least a portion of a radial length of
the main body.
4. The fan blade as set forth in claim 2, wherein a filler material
is provided between each of said opposed ends of said end caps and
cover and said main body, with said filler material reducing the
size of said steps, and said filler material being part of said
cover and said end cap for purposes of measuring said dimensions of
said steps.
5. The fan blade as set forth in claim 1, wherein an outer surface
of said fan blade having a surface roughness and said surface
roughness having a root means square value of less than about
60.times.10.sup.-6 inch on at least a portion of a radial length of
the main body.
6. The fan blade as set forth in claim 1, wherein a filler material
is provided between at least one of said ends of said one of the
end cap and the cover and said main body, with said filler material
reducing the size of said step, and said filler material being part
of said cover and said end cap for purposes of measuring said
dimensions of said steps.
7. The fan blade as set forth in claim 1, wherein said fan blade
having a chord length, and a ratio of said step dimension to said
chord length being less than or equal to about 0.001.
8. The fan blade as set forth in claim 1, wherein said step
occurring over at least from 20% of a blade span, measured from a
platform to a radially outer tip of said airfoil.
9. The fan blade as set forth in claim 1, wherein said fan blade is
designed to rotate with a fan tip corrected speed below 1225
ft/second (368 meters/second) at bucket cruise.
10. The fan blade as set forth in claim 1, comprising a shroud for
connecting the fan blade to an adjacent fan blade.
11. The fan blade as set forth in claim 1, wherein at least one of
the main body, cover, and cap is formed of aluminum or an aluminum
alloy.
12. The fan blade as set forth in claim 1, wherein at least one of
the main body, cover, and cap is formed of titanium or a titanium
alloy.
13. The fan blade as set forth in claim 1, wherein at least one of
the main body, cover, and cap is formed of composite.
14. The fan blade as set forth in claim 1, wherein at least one of
the main body is formed of composite, a metal, or an alloy, and
wherein the cover or the cap is formed of titanium or a titanium
alloy.
15. A gas turbine engine comprising: a fan drive turbine driving a
fan rotor having a plurality of blades through a gear reduction;
each of said plurality of blades including a main body having an
airfoil extending between a leading edge, and a trailing edge and
each of said plurality of blades having a chord length, and each of
said plurality of blades having at least one of: (a) a channel
closed by a cover, and (b) an end cap covering at least one of the
leading and trailing edges; and the at least one of the cover and
the end cap having a pair of opposed ends, and a step defined by at
least one of a suction wall and a pressure wall of the airfoil, to
an outer surface of said one of the cover and the end cap at one of
said opposed ends, and a ratio of said step dimension to said chord
length being less than or equal to about 0.001.
16. The gas turbine engine as set forth in claim 15, wherein said
main body including both said cover and said end cap, the end cap
at said leading edge, wherein there are a plurality of said step
defined at each of said opposed ends of said cover on one of said
suction wall and said pressure wall, and wherein said steps are
defined at each of said opposed ends of said end cap on both said
suction and pressure walls, and wherein all of said step dimensions
having said ratio being less than or equal to about 0.001.
17. The gas turbine engine as set forth in claim 16, wherein a
filler material is provided between each of said opposed ends of
said end caps and cover and said main body, with said filler
material reducing the size of said step dimensions, and said filler
material being part of said cover and said end cap.
18. The gas turbine engine as set forth in claim 15, wherein an
outer surface of said fan blade having a surface roughness and said
surface roughness having a root means square value of less than
about 60.times.10.sup.-6 inch on at least a portion of a radial
length of the main body.
19. The gas turbine engine as set forth in claim 15, wherein a
filler material is provided between at least one of said ends of
said one of the end cap and the cover and said main body, with said
filler material reducing the size of said step dimensions, and said
filler material being part of said one of said cover and said end
cap.
20. The gas turbine engine as set forth in claim 15, wherein said
step occurring over at least from 20% of a blade span, measured
from the platform to a radially outer tip of said airfoil.
21. The gas turbine engine as set forth in claim 15, wherein said
fan blade is designed to rotate with a fan tip corrected speed
below 1225 ft/second (368 meters/second) at bucket cruise.
22. The gas turbine engine as set forth in claim 15, comprising a
shroud connecting adjacent ones of said blades.
23. A method of manufacturing a fan blade comprising the steps of:
providing a main body extending between a leading edge and a
trailing edge, and having a suction wall and a pressure wall, with
said main body having at least one of a channel enclosed by a
cover, and an end cap covering at least one of said leading and
trailing edges, with said at least one of the cover and the end cap
being assembled to said main body, and defining a step by at least
one of the suction and pressure walls and an end of the at least
one of the cover and the end cap, and the step being made to be
less than or equal to about 0.010 inch (0.0254 centimeter) in
dimension.
24. The method as set forth in claim 23, wherein the size of the
step is reduced by adding a filler material which is considered
part of the at least one of the end cap and the cover for purposes
of measuring said step dimension.
25. The method as set forth in claim 23, wherein the main body
includes both the cover and the end cap and there are at least four
of said steps associated with ends of the cover having one spaced
toward the leading edge and one spaced towards the trailing edges
and ends of the end cap on each of the pressure and suction walls,
and all of said dimensions of said steps being made to be less than
or equal to about 0.010 inch (0.0254 centimeter).
26. The method as set forth in claim 23, wherein a surface
roughness of an outer surface of the main body, and the at least
one of the cover and the end cap is made to be less than about
60.times.10.sup.-6 inch over at least a portion of a radial length
of the main body.
27. The method as set forth in claim 23, wherein a machining step
is utilized to reduce the surface roughness.
28. The method as set forth in claim 23, wherein the fan blade
defining a chord length, and a ratio of said step dimension to said
chord length being less than or equal to about 0.001.
29. A method of designing a fan blade comprising: providing a main
body having an airfoil extending between a leading edge and a
trailing edge, and the fan blade having a chord length, the airfoil
extending radially outwardly from a platform, and the fan blade
having at least one of: (a) a channel closed by a cover; and (b) an
end cap covering at least one of the leading and trailing edges;
and the at least one of the cover and the end cap having a pair of
opposed ends, and a step defined by at least one of a suction wall
and a pressure wall of the airfoil to an outer surface of said one
of the cover and the end cap at one of said opposed ends, and a
ratio of said step dimension to said chord length being less than
or equal to about 0.001.
30. The method as set forth in claim 29, wherein an outer surface
of said fan blade having a surface roughness and said surface
roughness having a root means square value of less than about
60.times.10.sup.-6 inch on at least a portion of a radial length of
the main body.
Description
CROSS-REFERENCE TO RELATED APPLICATIONS
[0001] This application claims priority to U.S. Provisional
Application 61/727,786 filed Nov. 19, 2012, and U.S. Provisional
Application 61/884,295, filed Sep. 30, 2013.
BACKGROUND
[0002] Gas turbine engines may be provided with a fan for
delivering air to a compressor section and into a bypass section.
From the compressor section, the air is compressed and delivered
into a combustion section. The combustion section mixes fuel with
the air and combusts the combination. Products of the combustion
pass downstream over turbine rotors which are driven to rotate and,
in turn, drive the compressor and the fan.
[0003] Historically, a single turbine rotor may have driven a lower
pressure compressor and a fan at the same speed. More recently, a
gear reduction has been proposed such as intermediate the lower
pressure compressor and the fan, such that the fan can rotate at
lower speeds relative to the lower pressure compressor. With this
change, the diameter of the fan has increased dramatically and its
speed has decreased.
[0004] As the fan blade diameter increases, its weight be expected
to increase. To address this increase, hollow fan blades have been
developed. One type of hollow fan blade has at least one channel
and an outer cover attached over a main fan blade body to contain
the channel. In addition, an end cap may be placed on the fan
body.
[0005] The interface of the ends of the end cap and the cover skin,
relative to the main fan body, provides an interface that may be in
the form of a step.
SUMMARY
[0006] In a featured embodiment, a fan blade comprises a main body
having an airfoil extending between a leading edge and a trailing
edge. The fan blade has at least one of a channel closed by a
cover, and an end cap covering at least one of the leading and
trailing edges. At least one of a cover and an end cap has a pair
of opposed ends. A step is defined extending from at least one of a
suction wall and a pressure wall of the airfoil, to an outer
surface of the one of a cover and an end cap at one of the opposed
ends, and the step being less than or equal to about 0.010 inch
(0.0254 centimeter) in dimension.
[0007] In another embodiment according to the previous embodiment,
the main body includes both the cover and the end cap, the end cap
at the leading edge, wherein the steps are defined at each of the
opposed ends of the cover on one of the suction wall and the
pressure wall, and wherein the steps are defined at each of the
opposed ends of the end cap on both the suction and pressure walls,
and wherein all of the step dimensions are less than or equal to
about 0.010 inch (0.0254 centimeter).
[0008] In another embodiment according to any of the previous
embodiments, an outer surface of the fan blade has a surface
roughness. The surface roughness has a root means square value of
less than about 60.times.10.sup.-6 inch on at least a portion of a
radial length of the main body.
[0009] In another embodiment according to any of the previous
embodiments, a filler material is provided between each of the
opposed ends of the end caps and cover and the main body, with the
filler material reducing the size of the steps, and the filler
material being part of the cover and the end cap for purposes of
measuring the dimensions of the steps.
[0010] In another embodiment according to any of the previous
embodiments, an outer surface of the fan blade has a surface
roughness. The surface roughness has a root means square value of
less than about 60.times.10.sup.-6 inch on at least a portion of a
radial length of the main body.
[0011] In another embodiment according to any of the previous
embodiments, a filler material is provided between the end of the
one of an end cap and a cover and the main body, with the filler
material reducing the size of the steps, and the filler material
being part of the cover and the end cap for purposes of measuring
the dimensions of the steps.
[0012] In another embodiment according to any of the previous
embodiments, the fan blade has a chord length. A ratio of the step
dimension to the chord length is less than or equal to about
0.001.
[0013] In another embodiment according to any of the previous
embodiments, the step occurs over at least from 20% of a blade
span, measured from a platform to a radially outer tip of the
airfoil.
[0014] In another embodiment according to any of the previous
embodiments, the fan blade is designed to rotate with a fan tip
corrected speed below 1225 ft/second (368 meters/second) at bucket
cruise.
[0015] In another embodiment according to any of the previous
embodiments, a shroud connects the fan blade to an adjacent fan
blade.
[0016] In another embodiment according to any of the previous
embodiments, at least one of the main body, cover, and cap is
formed of aluminum or an aluminum alloy.
[0017] In another embodiment according to any of the previous
embodiments, at least one of the main body, cover, and cap is
formed of titanium or a titanium alloy.
[0018] In another embodiment according to any of the previous
embodiments, at least one of the main body, cover, and cap is
formed of composite.
[0019] In another embodiment according to any of the previous
embodiments, at least one of the main body is formed of composite,
a metal, or an alloy, and wherein the cover or the cap is formed of
titanium or a titanium alloy.
[0020] In another featured embodiment, a gas turbine engine
comprises a fan drive turbine driving a fan rotor having a
plurality of blades through a gear reduction. The blades include a
main body having an airfoil extending between a leading edge, and a
trailing edge and the blades having a chord length. The fan blade
has at least one of a channel closed by a cover and an end cap
covering at least one of the leading and trailing edges. At least
one of a cover and an end cap has a pair of opposed ends. A step is
defined extending from at least one of a suction wall and a
pressure wall of the airfoil, to an outer surface of the one of a
cover and an end cap at one of the opposed ends. A ratio of the
step dimension to the chord length is less than or equal to about
0.001.
[0021] In another embodiment according to the previous embodiment,
the main body includes both the cover and the end cap, the end cap
at the leading edge, wherein the steps are defined at each of the
opposed ends of the cover on one of the suction wall and the
pressure wall. The steps are defined at each of the opposed ends of
the end cap on both the suction and pressure walls. All of the step
dimensions have a ratio of less than or equal to about 0.001.
[0022] In another embodiment according to any of the previous
embodiments, a filler material is provided between each of the
opposed ends of the end caps and cover and the main body. The
filler material reduces the size of the step dimensions, and is
part of the cover and the end cap.
[0023] In another embodiment according to any of the previous
embodiments, an outer surface of the fan blade has a surface
roughness. The surface roughness has a root means square value of
less than about 60.times.10.sup.-6 inch on at least a portion of a
radial length of the main body.
[0024] In another embodiment according to any of the previous
embodiments, a filler material is provided between the end of the
one of an end cap and a cover and the main body. The filler
material reduces the size of the step dimensions, and is part of
the one of the cover and the end cap.
[0025] In another embodiment according to any of the previous
embodiments, the step occurs over at least from 20% of a blade
span, measured from the platform to a radially outer tip of the
airfoil.
[0026] In another embodiment according to any of the previous
embodiments, the fan blade is designed to rotate with a fan tip
corrected speed below 1225 ft/second (368 meters/second) at bucket
cruise.
[0027] In another embodiment according to any of the previous
embodiments, a shroud connects adjacent ones of the blades.
[0028] In another featured embodiment, a method of manufacturing a
fan blade comprising the steps of providing a main body extending
between a leading edge and a trailing edge, and having a suction
wall and a pressure wall. The main body has at least one of a
channel enclosed by a cover, and an end cap covering at least one
of the leading and trailing edges. At least one of a cover and an
end cap is assembled to the main body, and defines a step between
at least one of the suction and pressure walls and an end of the at
least one of a cover and an end cap. The step is made to be less
than or equal to about 0.010 inch (0.0254 centimeter) in
dimension.
[0029] In another embodiment according to the previous embodiment,
the size of the step is reduced by adding a filler material which
is considered part of the at least one of an end cap and a cover
for purposes of measuring the step dimension.
[0030] In another embodiment according to any of the previous
embodiments, the main body includes both a cover and an end cap.
There are at least four steps associated with ends of the cover
spaced toward both the leading and trailing edges and ends of the
end cap on each of the pressure and suction walls. All of the
dimensions of the steps are made to be less than or equal to about
0.010 inch (0.0254 centimeter).
[0031] In another embodiment according to any of the previous
embodiments, a surface roughness of the outer surface of the main
body, and the at least one of the cover and the end cap is made to
be less than about 60.times.10.sup.-6 inch over at least a portion
of a radial length of the main body.
[0032] In another embodiment according to any of the previous
embodiments, a machining step is utilized to reduce the surface
roughness.
[0033] In another embodiment according to any of the previous
embodiments, the fan blade defines a chord length. A ratio of the
step dimension to the chord length is less than or equal to about
0.001.
[0034] In another featured embodiment, a method of designing a fan
blade comprising providing a main body having an airfoil extending
between a leading edge and a trailing edge, and the fan blade
having a chord length. The airfoil extends radially outwardly from
a platform. The fan blade has at least one of a channel closed by a
cover, and an end cap covering at least one of the leading and
trailing edges. At least one of a cover and an end cap has a pair
of opposed ends. A step is defined extending from at least one of a
suction wall and a pressure wall of the airfoil to an outer surface
of the one of a cover and an end cap at one of the opposed ends. A
ratio of the step dimension to the chord length is less than or
equal to about 0.001.
[0035] In another embodiment according to the previous embodiment,
an outer surface of the fan blade has a surface roughness which has
a root means square value of less than about 60.times.10.sup.-6
inch on at least a portion of a radial length of the main body.
[0036] These and other features may be best understood from the
following drawings and specification.
BRIEF DESCRIPTION OF THE DRAWINGS
[0037] FIG. 1A schematically shows a gas turbine engine.
[0038] FIG. 1B is a side view of a removed fan blade.
[0039] FIG. 1C shows the fan blade of FIG. 1B in an installed
condition.
[0040] FIG. 2 is a cross-sectional view of the fan blade of FIGS.
1B, 1C.
[0041] FIG. 3A is a detail of a location identified by A in FIG.
2.
[0042] FIG. 3BA is a first possibility at an area identified by B
in FIG. 2.
[0043] FIG. 3BB shows a second possibility.
[0044] FIG. 3C shows a possibility at an area identified by C in
FIG. 2.
[0045] FIG. 4A shows a corrective method at the location of FIG.
3A.
[0046] FIG. 4BA shows a corrective method at the location of FIG.
3BA.
[0047] FIG. 4BB shows a corrective method at the location of FIG.
3BB.
[0048] FIG. 4C shows a corrective method at the location of FIG.
3C.
[0049] FIG. 5 explains a feature of the fan blade of FIGS.
1B-4C.
[0050] FIG. 6 shows another embodiment.
DETAILED DESCRIPTION
[0051] FIG. 1A schematically illustrates a gas turbine engine 20.
The gas turbine engine 20 is disclosed herein as a two-spool
turbofan that generally incorporates a fan section 22, a compressor
section 24, a combustor section 26 and a turbine section 28.
Alternative engines might include an augmentor section (not shown)
among other systems or features. The fan section 22 drives air
along a bypass flow path B in a bypass duct defined within a
nacelle 15, while the compressor section 24 drives air along a core
flow path C for compression and communication into the combustor
section 26 then expansion through the turbine section 28. Although
depicted as a two-spool turbofan gas turbine engine in the
disclosed non-limiting embodiment, it should be understood that the
concepts described herein are not limited to use with two-spool
turbofans as the teachings may be applied to other types of turbine
engines including three-spool architectures.
[0052] The exemplary engine 20 generally includes a low speed spool
30 and a high speed spool 32 mounted for rotation about an engine
central longitudinal axis A relative to an engine static structure
36 via several bearing systems 38. It should be understood that
various bearing systems 38 at various locations may alternatively
or additionally be provided, and the location of bearing systems 38
may be varied as appropriate to the application.
[0053] The low speed spool 30 generally includes an inner shaft 40
that interconnects a fan 42, a low pressure compressor 44 and a low
pressure turbine 46. The inner shaft 40 is connected to the fan 42
through a speed change mechanism, which in exemplary gas turbine
engine 20 is illustrated as a geared architecture 48 to drive the
fan 42 at a lower speed than the low speed spool 30. The high speed
spool 32 includes an outer shaft 50 that interconnects a high
pressure compressor 52 and high pressure turbine 54. A combustor 56
is arranged in exemplary gas turbine 20 between the high pressure
compressor 52 and the high pressure turbine 54. A mid-turbine frame
57 of the engine static structure 36 is arranged generally between
the high pressure turbine 54 and the low pressure turbine 46. The
mid-turbine frame 57 further supports bearing systems 38 in the
turbine section 28. The inner shaft 40 and the outer shaft 50 are
concentric and rotate via bearing systems 38 about the engine
central longitudinal axis A which is collinear with their
longitudinal axes.
[0054] The core airflow is compressed by the low pressure
compressor 44 then the high pressure compressor 52, mixed and
burned with fuel in the combustor 56, then expanded over the high
pressure turbine 54 and low pressure turbine 46. The mid-turbine
frame 57 includes airfoils 59 which are in the core airflow path C.
The turbines 46, 54 rotationally drive the respective low speed
spool 30 and high speed spool 32 in response to the expansion. It
will be appreciated that each of the positions of the fan section
22, compressor section 24, combustor section 26, turbine section
28, and fan drive gear system 48 may be varied. For example, gear
system 48 may be located aft of combustor section 26 or even aft of
turbine section 28, and fan section 22 may be positioned forward or
aft of the location of gear system 48.
[0055] The engine 20 in one example is a high-bypass geared
aircraft engine. In a further example, the engine 20 bypass ratio
is greater than about six (6), with an example embodiment being
greater than about ten (10), the geared architecture 48 is an
epicyclic gear train, such as a planetary gear system or other gear
system, with a gear reduction ratio of greater than about 2.3 and
the low pressure turbine 46 has a pressure ratio that is greater
than about five. In one disclosed embodiment, the engine 20 bypass
ratio is greater than about ten (10:1), the fan diameter is
significantly larger than that of the low pressure compressor 44,
and the low pressure turbine 46 has a pressure ratio that is
greater than about five 5:1. Low pressure turbine 46 pressure ratio
is pressure measured prior to inlet of low pressure turbine 46 as
related to the pressure at the outlet of the low pressure turbine
46 prior to an exhaust nozzle. The geared architecture 48 may be an
epicycle gear train, such as a planetary gear system or other gear
system, with a gear reduction ratio of greater than about 2.3:1. It
should be understood, however, that the above parameters are only
exemplary of one embodiment of a geared architecture engine and
that the present invention is applicable to other gas turbine
engines including direct drive turbofans.
[0056] A significant amount of thrust is provided by the bypass
flow B due to the high bypass ratio. The fan section 22 of the
engine 20 is designed for a particular flight condition--typically
cruise at about 0.8 Mach and about 35,000 feet. The flight
condition of 0.8 Mach and 35,000 ft, with the engine at its best
fuel consumption--also known as "bucket cruise Thrust Specific Fuel
Consumption (`TSFC`)"--is the industry standard parameter of lbm of
fuel being burned divided by lbf of thrust the engine produces at
that minimum point. "Low fan pressure ratio" is the pressure ratio
across the fan blade alone, without a Fan Exit Guide Vane ("FEGV")
system. The low fan pressure ratio as disclosed herein according to
one non-limiting embodiment is less than about 1.45. "Low corrected
fan tip speed" is the actual fan tip speed in ft/sec divided by an
industry standard temperature correction of [(Tram .degree.
R)/(518.7.degree. R)].sup.0.5. The "Low corrected fan tip speed" as
disclosed herein according to one non-limiting embodiment is less
than about 1150 ft/second.
[0057] A fan blade 120 which may be utilized in the gas turbine
engine 20 is illustrated in FIG. 1B having an airfoil 18 extending
radially outwardly from a platform 124, which may include, as
shown, a dovetail. A leading edge 21 and a trailing edge 122 define
the forward and rear limits of the airfoil. As shown in FIG. 1B,
the airfoil 18 extends from a radially inner end 501 adjacent the
platform 124 to a radially outer end 500.
[0058] As shown in FIG. 1C, a fan rotor 16 will receive the
platform 124 to mount the fan blade with the airfoil 18 extending
radially outwardly. As the rotor 16 is driven to rotate, it carries
the fan blade 120 with it.
[0059] FIG. 2 shows a cross-section of the fan blade 120 at the
airfoil 18. As shown, the leading edge 21 receives a cap 37 secured
to a main body 128. A cover 132 closes off cavities or channels 130
in the main body 128. The main body 128, the cap 37 and the cover
132 may all be formed of aluminum or various aluminum alloys. Other
materials, such as titanium, titanium alloys or other appropriate
metals, may alternatively be utilized for any one or more of the
cap 37, the cover 132, and/or the main body 128. Further, other,
non-metallic materials, such as composites or plastics, may
alternatively and/or additionally be utilized for any one or more
of the cap 37, the cover 132, and/or the main body 128.
[0060] In addition, while the fan blade 120 is shown having one
cover 132 and channels 130 having a closed inner end, it is also
possible that the main body 128 would provide a channel extending
across its entire thickness with covers at each side. As shown, a
plurality of ribs 126 separate the channels 130 in the
cross-section illustrated in FIG. 2. Filler material 100 may be
deposited within the channels 130 and would typically be of a
lighter weight than the main body 128.
[0061] Applicant has discovered that with the increasing diameter
of the fan blade 120 when utilized in geared gas turbine engine,
surface smoothness becomes important. If a laminar flow can be
achieved at the surface of the airfoil, the fuel burn efficiency
and the fan efficiency can be increased dramatically. However, it
is challenging to achieve laminar flow on fan blades 120 and, in
particular, as their diameter increases and their speed
decreases.
[0062] In fact, the fan blades mentioned above having a cover 132,
or an end cap 37, could be defined as assembled fan blades. These
assembled fan blades, applicant has recognized, create steps which
could move the actual flow further from laminar than it might be
with a solid fan blade.
[0063] As shown in FIG. 3A, an area identified by A in FIG. 2 is
enlarged. As shown, the cap 37 has an end 137, which is spaced
above an outer surface 210 of the main body 128. There is a step of
a dimension d.sub.1 between the two.
[0064] Similarly, FIG. 3BA shows one possibility at the location B
in FIG. 2. Here, the cover 132 has its end 139 spaced from the
outer surface 210 by a step of a dimension d.sub.2. This would be a
"negative" step.
[0065] FIG. 3BB shows the opposite wherein the cover 132 extends
above the surface 210 by a dimension d.sub.3. This might be called
a positive step.
[0066] Applicant has discovered that these steps must be minimized
to achieve laminar flow. In particular, the steps should be less
than or equal to about 0.010 inch (0.0254 centimeter). This
requirement can be performed as part of a quality control step and,
if any of the dimensions d.sub.1-d.sub.3 are outside of this
dimension, then corrective steps may be taken. As an example, as
shown in FIGS. 4A, 4BB and 4BA, a putty 301 may be included to take
up the step and reduce the sudden change between the two
surfaces.
[0067] Stated another way, a chord length C for the blade airfoils
18 may be defined as shown in FIG. 5. In one embodiment, rather
than the 0.010 inch (0.0254 centimeter) maximum, the dimensions
d.sub.1-d.sub.3 could be defined as being kept within a maximum
ratio with regard to the chord length C. In one embodiment, the
maximum allowable step was 0.010 inch (0.0254 centimeter), and the
chord length C was 10 inches (25.14 centimeters). In this
embodiment, a ratio of d.sub.1-d.sub.3 to C is less than or equal
to about 0.001. For this embodiment, C is measured at a tip of the
airfoil 18, and between its leading and trailing edges.
[0068] As mentioned above, the reduction of the steps may be
provided on each of the suction side 99 and pressure side 97 (see
FIG. 2) of the airfoil at all positions wherein there is a step. In
addition, the corrective measure may be more important at different
radial locations between the radial ends 500 and 501 of the airfoil
18 (see FIG. 1B).
[0069] For purposes of measuring the step height after the
corrective steps of FIGS. 4A, 4BA, and 4BB, the putty 301 is
considered part of the cover 132 or end cap 37. While putty is
disclosed, other filler materials may be used.
[0070] FIG. 3C shows yet another concern. A surface roughness at
the surface 210 may be identified as surface irregularities 211 and
may have a highest dimension d.sub.4. It would be desirable that
this surface roughness be minimized Applicant has found that
maintaining a surface roughness with a root means square value of
less than about 60.times.10.sup.-6-inch would result in a fan blade
providing more laminar flow.
[0071] As shown in FIG. 4C, this may be achieved by machining such
as applying a polishing or smoothing tool 310 to the irregularities
211.
[0072] Applicant has also discovered that the most important
portion of the fan blade to have the required smoothness are from
about 20% of the blade span radially outwardly, measured along a
length of airfoil 18 to 100% of the airfoil 18, at its tip.
[0073] In addition, applicant has determined that the results
achieved by a fan blade having the disclosed characteristics are
most beneficial when a fan tip corrected speed is below about 1225
ft/second at bucket cruise, and even more beneficial when the fan
speed is below 1150 ft/second. Further, the benefits are more
pronounced when the fan rotor carries 26 or fewer fan blades.
[0074] Now, an assembled fan blade having either the small step
size or the very smooth outer surface will achieve laminar flow
over a greater percentage of its surface area. These treatments can
be applied at any radial location between ends 501 and 500 or over
all of those portions. In addition, they may be provided on only
the suction side 99, only the pressure side 97 or both.
[0075] FIG. 6 shows an alternate embodiment fan rotor 300 wherein
blades 302 and 304 have a shroud 306 extending between them. The
shroud 306 provides additional rigidity to the structure to enhance
laminar flow across the fan blades 302, 304. The shroud of this
embodiment may be used in conjunction with any of the foregoing
surface treatments described with respect to the embodiments of
FIGS. 2-5.
[0076] Although an embodiment of this invention has been disclosed,
a worker of ordinary skill in this art would recognize that certain
modifications would come within the scope of this disclosure. For
that reason, the following claims should be studied to determine
the true scope and content of this disclosure.
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