U.S. patent application number 15/452175 was filed with the patent office on 2017-06-22 for flexible support structure for a geared architecture gas turbine engine.
The applicant listed for this patent is United Technologies Corporation. Invention is credited to Jason Husband, Michael E. McCune.
Application Number | 20170175582 15/452175 |
Document ID | / |
Family ID | 54701170 |
Filed Date | 2017-06-22 |
United States Patent
Application |
20170175582 |
Kind Code |
A1 |
McCune; Michael E. ; et
al. |
June 22, 2017 |
FLEXIBLE SUPPORT STRUCTURE FOR A GEARED ARCHITECTURE GAS TURBINE
ENGINE
Abstract
A gas turbine engine according to an example of the present
disclosure includes, among other things, a fan shaft driving a fan
having fan blades, a fan shaft support that supports the fan shaft.
The fan shaft support defines a fan shaft support transverse
stiffness. A gear system connected to the fan shaft includes a ring
gear defining a ring gear transverse stiffness, a gear mesh
defining a gear mesh transverse stiffness, and a reduction ratio
greater than 2.3. The ring gear transverse stiffness is less than
20% of the gear mesh transverse stiffness. A flexible support
supports said gear system and defines a flexible support transverse
stiffness. The flexible support transverse stiffness is less than
20% of the fan shaft support transverse stiffness.
Inventors: |
McCune; Michael E.;
(Colchester, CT) ; Husband; Jason; (South
Glstonbury, CT) |
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Applicant: |
Name |
City |
State |
Country |
Type |
United Technologies Corporation |
Farmington |
CT |
US |
|
|
Family ID: |
54701170 |
Appl. No.: |
15/452175 |
Filed: |
March 7, 2017 |
Related U.S. Patent Documents
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Application
Number |
Filing Date |
Patent Number |
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14859381 |
Sep 21, 2015 |
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15452175 |
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14604811 |
Jan 26, 2015 |
9239012 |
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14859381 |
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13623309 |
Sep 20, 2012 |
9133729 |
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14604811 |
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13342508 |
Jan 3, 2012 |
8297916 |
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13623309 |
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61494453 |
Jun 8, 2011 |
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Current U.S.
Class: |
1/1 |
Current CPC
Class: |
F01D 15/12 20130101;
F02C 7/36 20130101; F04D 29/056 20130101; F04D 29/321 20130101;
F04D 19/02 20130101; F04D 25/045 20130101; F01D 25/164 20130101;
Y02T 50/60 20130101; F01D 9/02 20130101; F01D 25/28 20130101; F01D
5/06 20130101; F02C 7/32 20130101; Y10T 29/49321 20150115; F05D
2220/32 20130101; F01D 25/16 20130101; F04D 29/053 20130101; F04D
29/325 20130101; F05D 2260/40311 20130101; F05D 2240/60 20130101;
F05D 2260/96 20130101; F02K 3/06 20130101 |
International
Class: |
F01D 25/16 20060101
F01D025/16; F04D 29/053 20060101 F04D029/053; F04D 25/04 20060101
F04D025/04; F02C 7/36 20060101 F02C007/36; F01D 5/06 20060101
F01D005/06; F01D 9/02 20060101 F01D009/02; F02K 3/06 20060101
F02K003/06; F04D 29/32 20060101 F04D029/32; F01D 15/12 20060101
F01D015/12 |
Claims
1. A gas turbine engine, comprising: a fan shaft driving a fan
having fan blades; a fan shaft support that supports said fan
shaft, said fan shaft support defining a fan shaft support
transverse stiffness; a gear system connected to said fan shaft,
said gear system includes a ring gear defining a ring gear
transverse stiffness, a gear mesh defining a gear mesh transverse
stiffness, and a reduction ratio greater than 2.3, wherein said
ring gear transverse stiffness is less than 20% of said gear mesh
transverse stiffness; and a flexible support supporting said gear
system defining a flexible support transverse stiffness, wherein
said flexible support transverse stiffness is less than 20% of said
fan shaft support transverse stiffness.
2. The gas turbine engine of claim 1, wherein said fan shaft
support defines a fan shaft support lateral stiffness and said
flexible support defines a flexible support lateral stiffness and
said flexible support lateral stiffness is less than 11% of said
fan shaft support lateral stiffness.
3. The gas turbine engine of claim 2, further comprising a fan
section having a low fan pressure ratio of less than about 1.45,
said low fan pressure ratio measured across the fan blades
alone.
4. The gas turbine engine of claim 3, further comprising a low
corrected fan tip speed less than about 1150 ft/second, wherein
said low corrected fan tip speed is an actual fan tip speed at an
ambient temperature divided by [(T.sub.ram.degree.
R)/(518.7.degree. R)].sup.0.5, where T represents said ambient
temperature in degrees Rankine.
5. The gas turbine engine of claim 4, further comprising a
mid-turbine frame including at least one airfoil extending into a
flow path, and wherein said fan shaft support is a K-frame bearing
support.
6. The gas turbine engine of claim 3, wherein said flexible support
transverse stiffness is less than 11% of said fan shaft support
transverse stiffness.
7. The gas turbine engine of claim 6, wherein said gear mesh
defines a gear mesh lateral stiffness and said flexible support
lateral stiffness is less than 8% of said gear mesh lateral
stiffness.
8. The gas turbine engine of claim 7, further comprising a high
speed spool including a two stage high pressure turbine, and a low
speed spool including a low pressure turbine with an inlet, an
outlet, and a low pressure turbine pressure ratio greater than 5:1,
wherein said low pressure turbine pressure ratio is a ratio of a
pressure measured prior to said inlet as related to a pressure at
said outlet prior to any exhaust nozzle.
9. The gas turbine engine of claim 7, wherein said flexible support
transverse stiffness is less than 8% of said gear mesh transverse
stiffness.
10. The gas turbine engine of claim 9, further comprising a two
stage high pressure turbine, and a low corrected fan tip speed less
than about 1150 ft/second, wherein said low corrected fan tip speed
is an actual fan tip speed at an ambient temperature divided by
[(Tram.degree. R)/(518.7.degree. R)]0.5, where T represents said
ambient temperature in degrees Rankine.
11. The gas turbine engine of claim 7, further comprising a gear
system input defining a gear system input transverse stiffness,
wherein said gear system input transverse stiffness is less than
20% of said fan shaft support transverse stiffness.
12. The gas turbine engine of claim 11, wherein said gear system
input transverse stiffness is less than 11% of said fan shaft
support transverse stiffness.
13. The gas turbine engine of claim 12, wherein said gear system
input defines a gear system input lateral stiffness and said gear
system input lateral stiffness is less than 11% of said fan shaft
support lateral stiffness.
14. The gas turbine engine of claim 11, wherein said gear system
input defines a gear system input lateral stiffness and said gear
system input lateral stiffness is less than 11% of said fan shaft
support lateral stiffness.
15. The gas turbine engine of claim 3, wherein said ring gear
transverse stiffness is less than 12% of said gear mesh transverse
stiffness.
16. The gas turbine engine of claim 15, further comprising a low
corrected fan tip speed less than about 1150 ft/second, wherein
said low corrected fan tip speed is an actual fan tip speed at an
ambient temperature divided by [(T.sub.ram.degree.
R)/(518.7.degree. R)].sup.0.5, where T represents said ambient
temperature in degrees Rankine, and further comprising at least one
bearing system, a high speed spool including an outer shaft and a
low speed spool including an inner shaft, wherein said gear system
includes a planet carrier and said fan shaft is mounted to said
planet carrier and said inner shaft and outer shaft are concentric
and rotate via said at least one bearing system about a
longitudinal axis of said engine.
17. A gas turbine engine, comprising: a fan shaft driving a fan
having fan blades; a fan shaft support that supports said fan
shaft, said fan shaft support defining a fan shaft support
transverse stiffness; a gear system connected to said fan shaft,
said gear system includes a ring gear defining a ring gear
transverse stiffness, a gear mesh defining a gear mesh transverse
stiffness and a gear mesh lateral stiffness, and a reduction ratio
greater than 2.3, wherein said ring gear transverse stiffness is
less than 20% of said gear mesh transverse stiffness; and a
flexible support supporting said gear system defining a flexible
support transverse stiffness and a flexible support lateral
stiffness, wherein said flexible support transverse stiffness is
less than 20% of said fan shaft support transverse stiffness, and
said flexible support lateral stiffness is less than 8% of said
gear mesh lateral stiffness.
18. The gas turbine engine of claim 17, further comprising a fan
section having a low fan pressure ratio of less than about 1.45,
said low fan pressure ratio measured across the fan blades
alone.
19. The gas turbine engine of claim 18, further comprising a low
corrected fan tip speed less than about 1150 ft/second, wherein
said low corrected fan tip speed is an actual fan tip speed at an
ambient temperature divided by [(T.sub.ram.degree.
R)/(518.7.degree. R)].sup.0.5, where T represents said ambient
temperature in degrees Rankine.
20. The gas turbine engine of claim 19, further comprising a
mid-turbine frame including at least one airfoil extending into a
flow path.
21. The gas turbine engine of claim 18, wherein said flexible
support transverse stiffness is less than 8% of said gear mesh
transverse stiffness.
22. The gas turbine engine of claim 21, further comprising a low
pressure turbine driving said gear system having an inlet, an
outlet, and a low pressure turbine pressure ratio greater than
about 5:1, wherein said low pressure turbine pressure ratio is a
ratio of a pressure measured prior to said inlet as related to a
pressure at said outlet prior to an exhaust nozzle.
23. The gas turbine engine of claim 22, further comprising a low
corrected fan tip speed less than about 1150 ft/second, wherein
said low corrected fan tip speed is an actual fan tip speed at an
ambient temperature divided by [(T.sub.ram.degree.
R)/(518.7.degree. R)].sup.0.5, where T represents said ambient
temperature in degrees Rankine.
24. The gas turbine engine of claim 23, wherein said gear system
drives a compressor rotor at a common speed with said fan
shaft.
25. The gas turbine engine of claim 24, further comprising a three
stage low pressure compressor, and a two stage high pressure
turbine.
26. The gas turbine engine of claim 18, further comprising a gear
system input defining a gear system input transverse stiffness,
wherein said gear system input transverse stiffness is less than
20% of said fan shaft support transverse stiffness.
27. The gas turbine engine of claim 26, further comprising said
gear system input defining a gear system input lateral stiffness
and said fan shaft support defining a fan shaft support lateral
stiffness, wherein said gear system input lateral stiffness is less
than 11% of said fan shaft support lateral stiffness.
28. The gas turbine engine of claim 27, wherein said gear system
input transverse stiffness is less than 11% of said fan shaft
support transverse stiffness.
29. The gas turbine engine of claim 18, further comprising a gear
system input defining a gear system input transverse stiffness,
wherein said gear system input transverse stiffness is less than
11% of said fan shaft support transverse stiffness.
30. The gas turbine engine of claim 29, wherein said gear system
input both transfers torque and facilitates the segregation of
vibrations.
Description
CROSS REFERENCE TO RELATED APPLICATIONS
[0001] The present disclosure is a continuation of U.S. patent
application Ser. No. 14/859,381, filed Sep. 21, 2015, which is a
continuation of U.S. patent application Ser. No. 14/604,811, filed
Jan. 26, 2015, now U.S. Pat. No. 9,239,012, issued Jan. 19, 2016,
which is a continuation-in-part of U.S. patent application Ser. No.
13/623,309, filed Sep. 20, 2012, now U.S. Pat. No. 9,133,729,
issued Sep. 15, 2015, which is a continuation-in-part of U.S.
application Ser. No. 13/342,508, filed Jan. 3, 2012, now U.S. Pat.
No. 8,297,916, issued Oct. 30, 2012, which claimed priority to
United States Provisional Application No. 61/494,453, filed Jun. 8,
2011.
BACKGROUND
[0002] The present disclosure relates to a gas turbine engine, and
more particularly to a flexible support structure for a geared
architecture therefor.
[0003] Epicyclic gearboxes with planetary or star gear trains may
be used in gas turbine engines for their compact designs and
efficient high gear reduction capabilities. Planetary and star gear
trains generally include three gear train elements: a central sun
gear, an outer ring gear with internal gear teeth, and a plurality
of planet gears supported by a planet carrier between and in meshed
engagement with both the sun gear and the ring gear. The gear train
elements share a common longitudinal central axis, about which at
least two rotate. An advantage of epicyclic gear trains is that a
rotary input can be connected to any one of the three elements. One
of the other two elements is then held stationary with respect to
the other two to permit the third to serve as an output.
[0004] In gas turbine engine applications, where a speed reduction
transmission is required, the central sun gear generally receives
rotary input from the powerplant, the outer ring gear is generally
held stationary and the planet gear carrier rotates in the same
direction as the sun gear to provide torque output at a reduced
rotational speed. In star gear trains, the planet carrier is held
stationary and the output shaft is driven by the ring gear in a
direction opposite that of the sun gear.
[0005] During flight, light weight structural cases deflect with
aero and maneuver loads causing significant amounts of transverse
deflection commonly known as backbone bending of the engine. This
deflection may cause the individual sun or planet gear's axis of
rotation to lose parallelism with the central axis. This deflection
may result in some misalignment at gear train journal bearings and
at the gear teeth mesh, which may lead to efficiency losses from
the misalignment and potential reduced life from increases in the
concentrated stresses.
SUMMARY
[0006] A gas turbine engine according to an example of the present
disclosure includes a fan shaft configured to drive a fan, a
support configured to support at least a portion of the fan shaft,
the support defining a support transverse stiffness and a support
lateral stiffness, a gear system coupled to the fan shaft, and a
flexible support configured to at least partially support the gear
system. The flexible support defines a flexible support transverse
stiffness with respect to the support transverse stiffness and a
flexible support lateral stiffness with respect to the support
lateral stiffness. The input defines an input transverse stiffness
with respect to the support transverse stiffness and an input
lateral stiffness with respect to the support lateral
stiffness.
[0007] In a further embodiment of any of the forgoing embodiments,
the support and the flexible support are mounted to a static
structure.
[0008] In a further embodiment of any of the forgoing embodiments,
the static structure is a front center body of the gas turbine
engine.
[0009] In a further embodiment of any of the forgoing embodiments,
the flexible support is mounted to a planet carrier of the gear
system, and the input is mounted to a sun gear of the gear
system.
[0010] In a further embodiment of any of the forgoing embodiments,
the fan shaft is mounted to a ring gear of the gear system.
[0011] In a further embodiment of any of the forgoing embodiments,
the gear system is a star system.
[0012] In a further embodiment of any of the forgoing embodiments,
the flexible support is mounted to a ring gear of the gear system,
and the input is mounted to a sun gear of the gear system.
[0013] In a further embodiment of any of the forgoing embodiments,
the fan shaft is mounted to a planet carrier of the gear
system.
[0014] In a further embodiment of any of the forgoing embodiments,
the flexible support transverse stiffness and the input transverse
stiffness are both less than the support transverse stiffness.
[0015] In a further embodiment of any of the forgoing embodiments,
the flexible support transverse stiffness and the input transverse
stiffness are each less than about 20% of the support transverse
stiffness.
[0016] In a further embodiment of any of the forgoing embodiments,
the flexible support transverse stiffness and the input transverse
stiffness are each less than about 11% of the support transverse
stiffness.
[0017] In a further embodiment of any of the forgoing embodiments,
the input to the gear system is coupled to a turbine section, and
the gear system is configured to drive a compressor rotor at a
common speed with the fan shaft.
[0018] A gas turbine engine according to an example of the present
disclosure includes a fan shaft configured to drive a fan, a
support configured to support at least a portion of the fan shaft,
and a gear system configured to drive the fan shaft. The gear
system includes a gear mesh that defines a gear mesh transverse
stiffness and a gear mesh lateral stiffness. A flexible support is
configured to at least partially support the gear system. The
flexible support defines a flexible support transverse stiffness
with respect to the gear mesh transverse stiffness and a flexible
support lateral stiffness with respect to the gear mesh lateral
stiffness. The input defines an input transverse stiffness with
respect to the gear mesh transverse stiffness and an input lateral
stiffness with respect to the gear mesh lateral stiffness.
[0019] In a further embodiment of any of the forgoing embodiments,
both the flexible support transverse stiffness and the input
transverse stiffness are less than the gear mesh transverse
stiffness.
[0020] In a further embodiment of any of the forgoing embodiments,
the flexible support transverse stiffness is less than about 8% of
the gear mesh transverse stiffness, the input transverse stiffness
is less than about 5% of the gear mesh transverse stiffness, and a
transverse stiffness of a ring gear of the gear system is less than
about 20% of the gear mesh transverse stiffness.
[0021] In a further embodiment of any of the forgoing embodiments,
a transverse stiffness of a planet journal bearing which supports a
planet gear of the gear system is less than or equal to the gear
mesh transverse stiffness.
[0022] In a further embodiment of any of the forgoing embodiments,
the support and the flexible support are mounted to a front center
body of the gas turbine engine.
[0023] A method of designing a gas turbine engine according to an
example of the present disclosure includes providing a fan shaft,
and providing a support configured to support at least a portion of
the fan shaft, the support defining at least one of a support
transverse stiffness and a support lateral stiffness, and providing
a gear system coupled to the fan shaft. The gear system includes a
gear mesh that defines a gear mesh lateral stiffness and a gear
mesh transverse stiffness. The method includes providing a flexible
support configured to at least partially support the gear system,
and providing an input to the gear system. The flexible support
defines a flexible support transverse stiffness with respect to the
gear mesh transverse stiffness and a flexible support lateral
stiffness with respect to the gear mesh lateral stiffness. The
input defines an input transverse stiffness with respect to the
gear mesh transverse stiffness and an input lateral stiffness with
respect to the gear mesh lateral stiffness.
[0024] In a further embodiment of any of the forgoing embodiments,
the flexible support lateral stiffness is less than the gear mesh
lateral stiffness, and the flexible support transverse stiffness is
less than the gear mesh transverse stiffness.
[0025] In a further embodiment of any of the forgoing embodiments,
both the flexible support transverse stiffness and the input
transverse stiffness are less than the gear mesh transverse
stiffness.
[0026] The various features and advantages of this invention will
become apparent to those skilled in the art from the following
detailed description of an embodiment. The drawings that accompany
the detailed description can be briefly described as follows.
BRIEF DESCRIPTION OF THE DRAWINGS
[0027] Various features will become apparent to those skilled in
the art from the following detailed description of the disclosed
non-limiting embodiment. The drawings that accompany the detailed
description can be briefly described as follows:
[0028] FIG. 1 is a schematic cross-section of a gas turbine
engine;
[0029] FIG. 2 is an enlarged cross-section of a section of the gas
turbine engine which illustrates a fan drive gear system
(FDGS);
[0030] FIG. 3 is a schematic view of a flex mount arrangement for
one non-limiting embodiment of the FDGS;
[0031] FIG. 4 is a schematic view of a flex mount arrangement for
another non-limiting embodiment of the FDGS;
[0032] FIG. 5 is a schematic view of a flex mount arrangement for
another non-limiting embodiment of a star system FDGS;
[0033] FIG. 6 is a schematic view of a flex mount arrangement for
another non-limiting embodiment of a planetary system FDGS.
[0034] FIG. 7 is a schematic view of a flex mount arrangement for
another non-limiting embodiment of a star system FDGS;
[0035] FIG. 8 is a schematic view of a flex mount arrangement for
another non-limiting embodiment of a planetary system FDGS;
[0036] FIG. 9 shows another embodiment; and
[0037] FIG. 10 shows yet another embodiment.
DETAILED DESCRIPTION
[0038] FIG. 1 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.
[0039] 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.
[0040] The low speed spool 30 generally includes an inner shaft 40
that interconnects a fan 42, a first (or low) pressure compressor
44 and a first (or 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 second (or high) pressure compressor 52 and a
second (or 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.
[0041] 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.
[0042] 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.
[0043] 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 1 bm
of fuel being burned divided by 1 bf 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.
[0044] With reference to FIG. 2, the geared architecture 48
generally includes a fan drive gear system (FDGS) 60 driven by the
low speed spool 30 (illustrated schematically) through an input 62.
The input 62, which may be in the form of a coupling, both
transfers torque from the low speed spool 30 to the geared
architecture 48 and facilitates the segregation of vibrations and
other transients therebetween. In the disclosed non-limiting
embodiment, the FDGS 60 may include an epicyclic gear system which
may be, for example, a star system or a planet system.
[0045] The input coupling 62 may include an interface spline 64
joined, by a gear spline 66, to a sun gear 68 of the FDGS 60. The
sun gear 68 is in meshed engagement with multiple planet gears 70,
of which the illustrated planet gear 70 is representative. Each
planet gear 70 is rotatably mounted in a planet carrier 72 by a
respective planet journal bearing 75. Rotary motion of the sun gear
68 urges each planet gear 70 to rotate about a respective
longitudinal axis P.
[0046] Each planet gear 70 is also in meshed engagement with
rotating ring gear 74 that is mechanically connected to a fan shaft
76. Since the planet gears 70 mesh with both the rotating ring gear
74 as well as the rotating sun gear 68, the planet gears 70 rotate
about their own axes to drive the ring gear 74 to rotate about
engine axis A. The rotation of the ring gear 74 is conveyed to the
fan 42 (FIG. 1) through the fan shaft 76 to thereby drive the fan
42 at a lower speed than the low speed spool 30. It should be
understood that the described geared architecture 48 is but a
single non-limiting embodiment and that various other geared
architectures will alternatively benefit herefrom.
[0047] With reference to FIG. 3, a flexible support 78 supports the
planet carrier 72 to at least partially support the FDGS 60A with
respect to the static structure 36 such as a front center body
which facilitates the segregation of vibrations and other
transients therebetween. It should be understood that various gas
turbine engine case structures may alternatively or additionally
provide the static structure and flexible support 78. It should be
understood that lateral as defined herein is generally transverse
to the axis of rotation A and the term "transverse" refers to a
pivotal bending movement with respect to the axis of rotation A
which typically absorbs deflection applied to the FDGS 60. The
static structure 36 may further include a number 1 and 1.5 bearing
support static structure 82 which is commonly referred to as a
"K-frame" which supports the number 1 and number 1.5 bearing
systems 38A, 38B. Notably, the K-frame bearing support defines a
lateral stiffness (represented as Kframe in FIG. 3) and a
transverse stiffness (represented as Kframe.sup.BEND in FIG. 3) as
the referenced factors in this non-limiting embodiment.
[0048] In this disclosed non-limiting embodiment, the lateral
stiffness (KFS; KIC) of both the flexible support 78 and the input
coupling 62 are each less than about 11% of the lateral stiffness
(Kframe). That is, the lateral stiffness of the entire FDGS 60 is
controlled by this lateral stiffness relationship. Alternatively,
or in addition to this relationship, the transverse stiffness of
both the flexible support 78 and the input coupling 62 are each
less than about 11% of the transverse stiffness (Kframe.sup.BEND).
That is, the transverse stiffness of the entire FDGS 60 is
controlled by this transverse stiffness relationship.
[0049] With reference to FIG. 4, another non-limiting embodiment of
a FDGS 60B includes a flexible support 78' that supports a
rotationally fixed ring gear 74'. The fan shaft 76' is driven by
the planet carrier 72' in the schematically illustrated planet
system which otherwise generally follows the star system
architecture of FIG. 3.
[0050] With reference to FIG. 5, the lateral stiffness relationship
within a FDGS 60C itself (for a star system architecture) is
schematically represented. The lateral stiffness (KIC) of an input
coupling 62, a lateral stiffness (KFS) of a flexible support 78, a
lateral stiffness (KRG) of a ring gear 74 and a lateral stiffness
(KJB) of a planet journal bearing 75 are controlled with respect to
a lateral stiffness (KGM) of a gear mesh within the FDGS 60.
[0051] In the disclosed non-limiting embodiment, the stiffness
(KGM) may be defined by the gear mesh between the sun gear 68 and
the multiple planet gears 70. The lateral stiffness (KGM) within
the FDGS 60 is the referenced factor and the static structure 82'
rigidly supports the fan shaft 76. That is, the fan shaft 76 is
supported upon bearing systems 38A, 38B which are essentially
rigidly supported by the static structure 82'. The lateral
stiffness (KJB) may be mechanically defined by, for example, the
stiffness within the planet journal bearing 75 and the lateral
stiffness (KRG) of the ring gear 74 may be mechanically defined by,
for example, the geometry of the ring gear wings 74L, 74R (FIG.
2).
[0052] In the disclosed non-limiting embodiment, the lateral
stiffness (KRG) of the ring gear 74 is less than about 12% of the
lateral stiffness (KGM) of the gear mesh; the lateral stiffness
(KFS) of the flexible support 78 is less than about 8% of the
lateral stiffness (KGM) of the gear mesh; the lateral stiffness
(KJB) of the planet journal bearing 75 is less than or equal to the
lateral stiffness (KGM) of the gear mesh; and the lateral stiffness
(KIC) of an input coupling 62 is less than about 5% of the lateral
stiffness (KGM) of the gear mesh.
[0053] With reference to FIG. 6, another non-limiting embodiment of
a lateral stiffness relationship within a FDGS 60D itself are
schematically illustrated for a planetary gear system architecture,
which otherwise generally follows the star system architecture of
FIG. 5.
[0054] It should be understood that combinations of the above
lateral stiffness relationships may be utilized as well. The
lateral stiffness of each of structural components may be readily
measured as compared to film stiffness and spline stiffness which
may be relatively difficult to determine.
[0055] By flex mounting to accommodate misalignment of the shafts
under design loads, the FDGS design loads have been reduced by more
than 17% which reduces overall engine weight. The flex mount
facilitates alignment to increase system life and reliability. The
lateral flexibility in the flexible support and input coupling
allows the FDGS to essentially `float` with the fan shaft during
maneuvers. This allows: (a) the torque transmissions in the fan
shaft, the input coupling and the flexible support to remain
constant during maneuvers; (b) maneuver induced lateral loads in
the fan shaft (which may otherwise potentially misalign gears and
damage teeth) to be mainly reacted to through the number 1 and 1.5
bearing support K-frame; and (c) both the flexible support and the
input coupling to transmit small amounts of lateral loads into the
FDGS. The splines, gear tooth stiffness, journal bearings, and ring
gear ligaments are specifically designed to minimize gear tooth
stress variations during maneuvers. The other connections to the
FDGS are flexible mounts (turbine coupling, case flex mount). These
mount spring rates have been determined from analysis and proven in
rig and flight testing to isolate the gears from engine maneuver
loads. In addition, the planet journal bearing spring rate may also
be controlled to support system flexibility.
[0056] FIG. 7 is similar to FIG. 5 but shows the transverse
stiffness relationships within the FDGS 60C (for a star system
architecture). The transverse stiffness (KIC.sup.BEND) of the input
coupling 62, a transverse stiffness (KFS.sup.BEND) of the flexible
support 78, a transverse stiffness (KRG.sup.BEND) of the ring gear
74 and a transverse stiffness (KJB.sup.BEND) of the planet journal
bearing 75 are controlled with respect to a transverse stiffness
(KGM.sup.BEND) of the gear mesh within the FDGS 60.
[0057] In the disclosed non-limiting embodiment, the stiffness
(KGM.sup.BEND) may be defined by the gear mesh between the sun gear
68 and the multiple planet gears 70. The transverse stiffness
(KGM.sup.BEND) within the FDGS 60 is the referenced factor and the
static structure 82' rigidly supports the fan shaft 76. That is,
the fan shaft 76 is supported upon bearing systems 38A, 38B which
are essentially rigidly supported by the static structure 82'. The
transverse stiffness (KJB.sup.BEND) may be mechanically defined by,
for example, the stiffness within the planet journal bearing 75 and
the transverse stiffness (KRG.sup.BEND) of the ring gear 74 may be
mechanically defined by, for example, the geometry of the ring gear
wings 74L, 74R (FIG. 2).
[0058] In the disclosed non-limiting embodiment, the transverse
stiffness (KRG.sup.BEND) of the ring gear 74 is less than about 12%
of the transverse stiffness (KGM.sup.BEND) of the gear mesh; the
transverse stiffness (KFS.sup.BEND) of the flexible support 78 is
less than about 8% of the transverse stiffness (KGM.sup.BEND) of
the gear mesh; the transverse stiffness (KJB.sup.BEND) of the
planet journal bearing 75 is less than or equal to the transverse
stiffness (KGM.sup.BEND) of the gear mesh; and the transverse
stiffness (KIC.sup.BEND) of an input coupling 62 is less than about
5% of the transverse stiffness (KGM.sup.BEND) of the gear mesh.
[0059] FIG. 8 is similar to FIG. 6 but shows the transverse
stiffness relationship within the FDGS 60D for the planetary gear
system architecture.
[0060] FIG. 9 shows an embodiment 200, wherein there is a fan drive
turbine 208 driving a shaft 206 to in turn drive a fan rotor 202. A
gear reduction 204 may be positioned between the fan drive turbine
208 and the fan rotor 202. This gear reduction 204 may be
structured, mounted and operate like the gear reduction disclosed
above. A compressor rotor 210 is driven by an intermediate pressure
turbine 212, and a second stage compressor rotor 214 is driven by a
turbine rotor 216. A combustion section 218 is positioned
intermediate the compressor rotor 214 and the turbine section
216.
[0061] FIG. 10 shows yet another embodiment 300 wherein a fan rotor
302 and a first stage compressor 304 rotate at a common speed. The
gear reduction 306 (which may be structured, mounted and operate as
disclosed above) is intermediate the compressor rotor 304 and a
shaft 308, which is driven by a low pressure turbine section.
[0062] It should be understood that relative positional terms such
as "forward," "aft," "upper," "lower," "above," "below," and the
like are with reference to the normal operational attitude of the
vehicle and should not be considered otherwise limiting.
[0063] It should be understood that like reference numerals
identify corresponding or similar elements throughout the several
drawings. It should also be understood that although a particular
component arrangement is disclosed in the illustrated embodiment,
other arrangements will benefit herefrom.
[0064] Although particular step sequences are shown, described, and
claimed, it should be understood that steps may be performed in any
order, separated or combined unless otherwise indicated and will
still benefit from the present disclosure.
[0065] The foregoing description is exemplary rather than defined
by the limitations within. Various non-limiting embodiments are
disclosed herein, however, one of ordinary skill in the art would
recognize that various modifications and variations in light of the
above teachings will fall within the scope of the appended claims.
It is therefore to be understood that within the scope of the
appended claims, the disclosure may be practiced other than as
specifically described. For that reason the appended claims should
be studied to determine true scope and content.
* * * * *