U.S. patent application number 15/872111 was filed with the patent office on 2018-06-07 for gear system architecture for gas turbine engine.
The applicant listed for this patent is United Technologies Corporation. Invention is credited to William G. Sheridan.
Application Number | 20180156134 15/872111 |
Document ID | / |
Family ID | 49487663 |
Filed Date | 2018-06-07 |
United States Patent
Application |
20180156134 |
Kind Code |
A1 |
Sheridan; William G. |
June 7, 2018 |
GEAR SYSTEM ARCHITECTURE FOR GAS TURBINE ENGINE
Abstract
A disclosed gas turbine engine includes a fan including a
plurality of fan blades rotatable about an axis, a bypass duct, a
compressor section and a bypass ratio greater than 10. A combustor
in fluid communication with the compressor section and a fan drive
turbine in communication with the combustor. A gear system provides
a speed reduction between the fan drive turbine and the fan and
transfer power input from the fan drive turbine to the fan at an
efficiency greater than 98%.
Inventors: |
Sheridan; William G.;
(Southington, CT) |
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Applicant: |
Name |
City |
State |
Country |
Type |
United Technologies Corporation |
Farmington |
CT |
US |
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Family ID: |
49487663 |
Appl. No.: |
15/872111 |
Filed: |
January 16, 2018 |
Related U.S. Patent Documents
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Application
Number |
Filing Date |
Patent Number |
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14190159 |
Feb 26, 2014 |
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15872111 |
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PCT/US2013/041761 |
May 20, 2013 |
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14190159 |
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13557614 |
Jul 25, 2012 |
8572943 |
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PCT/US2013/041761 |
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61653731 |
May 31, 2012 |
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Current U.S.
Class: |
1/1 |
Current CPC
Class: |
Y02T 50/60 20130101;
F02C 7/06 20130101; F01D 25/18 20130101; F02C 3/107 20130101; Y02T
50/671 20130101; F02K 3/06 20130101; F05D 2260/4031 20130101; F02C
7/36 20130101 |
International
Class: |
F02C 7/36 20060101
F02C007/36; F02C 7/06 20060101 F02C007/06; F02C 3/107 20060101
F02C003/107; F01D 25/18 20060101 F01D025/18; F02K 3/06 20060101
F02K003/06 |
Claims
1. A gas turbine engine comprising: a fan including a plurality of
fan blades rotatable about an axis; a bypass duct; a compressor
section; a bypass ratio greater than 10, the bypass ratio being
defined as the portion of air delivered into the bypass duct
divided by the amount of air delivered into the compressor section;
a combustor in fluid communication with the compressor section; a
fan drive turbine in communication with the combustor; a gear
system providing a speed reduction between the fan drive turbine
and the fan and transferring power input from the fan drive turbine
to the fan at an efficiency greater than 98%; a mount flexibly
supporting the gear system extending from a static structure of the
engine to accommodate radial movement between the gear system and
the static structure; and a lubrication system providing lubricant
flow to a plurality of rotating engine components including at
least an engine bearing assembly and the gear system, wherein the
lubrication system has a maximum capacity for removing thermal
energy from the lubricant equal to no more than 2% of power input
into the gear system by the fan drive turbine during operation of
the engine.
2. The gas turbine engine recited in claim 1, further comprising a
second turbine in communication with the combustor, wherein the
gear system comprises a sun gear, a rotatable carrier, a plurality
of planet gears supported on the carrier and driven by the sun
gear, and a ring gear circumscribing the plurality of planet gears,
and wherein the gear system comprises a gear reduction ratio of
greater than 2.3.
3. The gas turbine engine recited in claim 2, the fan drive turbine
including a plurality of fan drive turbine rotors, wherein the
plurality of fan blades is less than twenty (20) and the plurality
of fan drive turbine rotors is less than six (6), and a ratio
between the number of fan blades and the number of fan drive
turbine rotors is between 3.3 and 8.6.
4. The gas turbine engine recited in claim 3, wherein the second
turbine is a two stage turbine.
5. The gas turbine engine recited in claim 4, wherein the fan drive
turbine is a three stage turbine.
6. The gas turbine engine recited in claim 4, wherein the
compressor section has a first compressor and a second compressor,
the first compressor is a three stage compressor, and wherein the
second turbine drives the second compressor through an outer shaft
and an inner shaft connects the fan and the first compressor
section to the fan drive turbine.
7. The gas turbine engine recited in claim 4, wherein the
compressor section includes a first compressor and a second
compressor, the second compressor is an eight stage compressor.
8. The gas turbine engine recited in claim 4, further comprising a
low corrected fan tip speed less than about 1150 ft/second, wherein
the 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 the ambient temperature in
degrees Rankine.
9. The gas turbine engine recited in claim 8, further comprising an
input shaft, the fan drive turbine driving the input shaft in
operation, and a flexible support structure supporting at least one
gear of the gear system relative to the input shaft.
10. The gas turbine engine recited in claim 9, wherein the flexible
support structure includes a spring rate that allows a defined
amount of deflection and misalignment of the at least one gear.
11. The gas turbine engine recited in claim 10, further comprising
a load limiter constraining movement of the gear system in the
event of an unbalanced load condition.
12. The gas turbine engine recited in claim 11, wherein the load
limiter constrains radial loads and torsional loads within defined
limits.
13. The gas turbine engine recited in claim 12, wherein the load
limiter includes a stop.
14. The gas turbine engine recited in claim 4, wherein the
lubrication system comprises a first lubrication system providing
lubricant flow to the rotating engine components, and a second
lubrication system that provides lubricant flow to the gear system
in response to an interruption in lubricant flow from the first
lubrication system.
15. The gas turbine engine recited in claim 1, further comprising a
second turbine in communication with the combustor, wherein the
gear system comprises a sun gear driven by the fan drive turbine, a
non-rotatable carrier, a plurality of intermediate gears supported
on the carrier and driven by the sun gear, and a ring gear
circumscribing the plurality of intermediate gears, and wherein the
gear system comprises a gear reduction ratio of greater than
2.3.
16. The gas turbine engine recited in claim 15, the fan drive
turbine including a plurality of fan drive turbine rotors, wherein
the plurality of fan blades is less than twenty (20) and the
plurality of fan drive turbine rotors is less than six (6) fan
drive turbine rotors, and a ratio between the number of fan blades
and the number of fan drive turbine rotors is between 3.3 and
8.6.
17. The gas turbine engine recited in claim 16, wherein the second
turbine is a two stage turbine.
18. The gas turbine engine recited in claim 17, further comprising
a low corrected fan tip speed less than about 1150 ft/second,
wherein the 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 the ambient temperature in
degrees Rankine.
19. The gas turbine engine recited in claim 18, wherein the mount
further comprises a spring rate defined to accommodate deflections
occurring during normal operation of the gear system.
20. The gas turbine engine recited in claim 19, wherein the fan
drive turbine drives the gear system through a support structure
including a spring rate that allows a defined amount of deflection
and misalignment among at least some of the plurality of gears of
the gear system.
21. The gas turbine engine recited in claim 20, further comprising
a load limiter constraining movement of the gear system during an
unbalanced load condition.
22. The gas turbine engine recited in claim 16, wherein the
lubrication system comprises a first lubrication system providing
lubricant flow to the rotating engine components, and a second
lubrication system that provides lubricant flow to the gear system
in response to an interruption in lubricant flow from the first
lubrication system.
23. The gas turbine engine recited in claim 1, wherein the
lubrication system comprises a first lubrication system providing
lubricant flow to the rotating engine components, and a second
lubrication system supplementing operation of the first lubrication
system.
24. The gas turbine engine recited in claim 23, wherein the second
lubrication system supplies lubricant flow to the gear system in
response to an interruption in lubricant flow from the first
lubrication system.
25. The gas turbine engine recited in claim 24, wherein the gear
system comprises a gear reduction ratio of greater than 2.3, the
plurality of fan blades is less than twenty (20) and the fan drive
turbine including less than six (6) fan drive turbine rotors, and a
ratio between the number of fan blades and the number of fan drive
turbine rotors is between 3.3 and 8.6.
26. The gas turbine engine recited in claim 25, wherein the
lubrication system has a maximum capacity for removing thermal
energy from the lubricant equal to no more than about 1% of power
input into the gear system by the fan drive turbine during
operation of the engine.
27. The gas turbine engine recited in claim 1, further comprising a
bypass ratio greater than ten (10), the bypass ratio being defined
as the portion of air delivered into a bypass duct divided by the
amount of air delivered into the compressor section.
28. The gas turbine engine recited in claim 27, further comprising
an input shaft arranged in a driving relationship to the fan drive
turbine, a flexible support structure supporting at least one gear
of the gear system relative to the input shaft, including a spring
rate that allows a defined amount of deflection and misalignment of
the at least one gear.
29. The gas turbine engine recited in claim 28, further comprising
a load limiter constraining movement of the gear system beyond a
defined amount during an unbalanced load condition.
30. The gas turbine engine recited in claim 29, further comprising
a low corrected fan tip speed less than about 1150 ft/second,
wherein the 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 the ambient temperature in
degrees Rankine, and wherein the gear system comprises a gear
reduction ratio of greater than 2.3, the plurality of fan blades is
less than twenty (20) and the plurality of fan drive turbine rotors
is less than six (6) fan drive turbine rotors, and a ratio between
the number of fan blades and the number of fan drive turbine rotors
is between 3.3 and 8.6.
Description
CROSS-REFERENCE TO RELATED APPLICATION
[0001] This application is a continuation of U.S. patent
application Ser. No. 14/190,159, filed Feb. 26, 2014, which is a
continuation of International Application No. PCT/US2013/041761
filed May 20, 2013 that claims priority to U.S. Provisional
Application No. 61/653,731 filed May 31, 2012 and U.S. patent
application Ser. No. 13/557,614 filed Jul. 25, 2012, now U.S. Pat.
No. 8,572,943 granted Nov. 5, 2013.
BACKGROUND
[0002] A gas turbine engine typically includes a fan section, a
compressor section, a combustor section and a turbine section. Air
entering the compressor section is compressed and delivered into
the combustion section where it is mixed with fuel and ignited to
generate a high-speed exhaust gas flow. The high-speed exhaust gas
flow expands through the turbine section to drive the compressor
and the fan section. The compressor section typically includes low
and high pressure compressors, and the turbine section includes low
and high pressure turbines.
[0003] The high pressure turbine drives the high pressure
compressor through an outer shaft to form a high spool, and the low
pressure turbine drives the low pressure compressor through an
inner shaft to form a low spool. A speed reduction device such as
an epicyclical gear assembly may be utilized to drive the fan
section such that the fan section may rotate at a speed different
than the turbine section so as to increase the overall propulsive
efficiency of the engine. The efficiency at which the gear assembly
transfers power is a consideration in the development of a gear
driven fan. Power or energy not transferred through the gearbox
typically results in the generation of heat that is removed with a
lubrication system. The more heat generated, the larger and heavier
the lubrication system.
[0004] Although geared architectures can provide improved
propulsive efficiency, other factors including heat removal and
lubrication can detract from the improved propulsive efficiency.
Accordingly, turbine engine manufacturers continue to seek further
improvements to engine performance including improvements to
thermal, transfer and propulsive efficiencies.
SUMMARY
[0005] A fan drive gear system for a gas turbine engine according
to an exemplary embodiment of this disclosure, among other possible
things includes a gear system that provides a speed reduction
between a fan drive turbine and a fan, a mount flexibly supporting
portions of the gear system, and a lubrication system providing
lubricant to the gear system and removing thermal energy produced
by the gear system, wherein the lubrication system includes a
capacity for removing thermal energy equal to less than about 2% of
power input into the gear system.
[0006] In a further embodiment of the foregoing fan drive gear
system, wherein the gear system transfers power input from the fan
drive turbine to the fan at an efficiency greater than about
98%.
[0007] In a further embodiment of any of the foregoing fan drive
gear systems, wherein the lubrication system includes a capacity
for removing thermal energy equal to less than about 1% of power
input into the gear system.
[0008] In a further embodiment of any of the foregoing fan drive
gear systems, wherein the lubrication system comprises a main
lubrication system providing lubricant flow to the gear system and
an auxiliary lubrication system that provides lubricant to the gear
system responsive to an interruption of lubricant flow from the
main lubrication system.
[0009] In a further embodiment of any of the foregoing fan drive
gear systems, wherein the mount includes a load limiter for
limiting movement of the gear system responsive to an unbalanced
condition.
[0010] In a further embodiment of any of the foregoing fan drive
gear systems, wherein the gear system comprises a sun gear driven
by the fan drive turbine, a non-rotatable carrier, a plurality of
star gears supported on the carrier and driven by the sun gear and
a ring gear circumscribing the plurality of star gears.
[0011] In a further embodiment of any of the foregoing fan drive
gear systems, wherein the mount includes a first flexible coupling
between an input shaft driven by the fan drive turbine and the sun
gear, and a second flexible coupling between a fixed structure and
the carrier.
[0012] In a further embodiment of any of the foregoing fan drive
gear systems, wherein the gear system comprises a sun gear driven
by the fan drive turbine, a rotatable carrier, a plurality of
planet gears supported on the carrier and driven by the sun gear,
and a ring gear circumscribing the plurality of planet gears.
[0013] In a further embodiment of any of the foregoing fan drive
gear systems, wherein the mount includes a first flexible coupling
between an input shaft driven by the fan drive turbine and the sun
gear, and a second flexible coupling between a fixed structure and
the ring gear.
[0014] A gas turbine engine according to an exemplary embodiment of
this disclosure, among other possible things includes a fan
including a plurality of fan blades rotatable about an axis, a
compressor section, a combustor in fluid communication with the
compressor section, a fan drive turbine in communication with the
combustor, a gear system that provides a speed reduction between
the fan drive turbine and the fan, the gear system transfers power
input from the fan drive turbine to the fan at an efficiency
greater than about 98%, a mount flexibly supporting portions of the
gear system, and a lubrication system providing lubricant to the
gear system and removing thermal energy from the gear system
produced by the gear system.
[0015] In a further embodiment of the foregoing gas turbine engine,
wherein the lubrication system includes a capacity for removing
thermal energy equal to less than about 2% of power input into the
gear system.
[0016] In a further embodiment of any of the foregoing gas turbine
engines, wherein the lubrication system includes a capacity for
removing thermal energy equal to less than about 1% of power input
into the gear system.
[0017] In a further embodiment of any of the foregoing gas turbine
engines, wherein the lubrication system comprises a main
lubrication system providing lubricant flow to the gear system and
an auxiliary lubrication system that provides lubricant to the gear
system responsive to an interruption of lubricant flow from the
main lubrication system.
[0018] In a further embodiment of any of the foregoing gas turbine
engines, wherein the gear system comprises a sun gear driven by the
fan drive turbine, a non-rotatable carrier, a plurality of star
gears supported on the carrier and driven by the sun gear and a
ring gear circumscribing the plurality of star gears and the mount
includes a first flexible coupling between an input shaft driven by
the fan drive turbine and the sun gear, and a second flexible
coupling between a fixed structure and the carrier.
[0019] In a further embodiment of any of the foregoing gas turbine
engines, wherein the gear system comprises a sun gear driven by the
fan drive turbine, a rotatable carrier, a plurality of planet gears
supported on the carrier and driven by the sun gear, and a ring
gear circumscribing the plurality of planet gears and the mount
includes a first flexible coupling between an input shaft driven by
the fan drive turbine and the sun gear, and a second flexible
coupling between a fixed structure and the ring gear.
[0020] In a further embodiment of any of the foregoing gas turbine
engines, wherein the mount includes a load limiter for limiting
movement of the gear system responsive to an unbalanced
condition.
[0021] In a further embodiment of any of the foregoing gas turbine
engines, wherein the gear system comprises a gear reduction having
a gear ratio greater than about 2.3.
[0022] In a further embodiment of any of the foregoing gas turbine
engines, wherein said fan delivers a portion of air into a bypass
duct, and a bypass ratio being defined as the portion of air
delivered into the bypass duct divided by the amount of air
delivered into the compressor section, with the bypass ratio being
greater than about 6.0.
[0023] In a further embodiment of any of the foregoing gas turbine
engines, wherein a fan pressure ratio across the fan is less than
about 1.5.
[0024] In a further embodiment of any of the foregoing gas turbine
engines, wherein said fan has 26 or fewer blades.
[0025] Although the different examples have the specific components
shown in the illustrations, embodiments of this invention are not
limited to those particular combinations. It is possible to use
some of the components or features from one of the examples in
combination with features or components from another one of the
examples.
[0026] These and other features disclosed herein can be best
understood from the following specification and drawings, the
following of which is a brief description.
BRIEF DESCRIPTION OF THE DRAWINGS
[0027] FIG. 1 is a schematic view of an example gas turbine
engine.
[0028] FIG. 2 is a schematic view of an example fan drive gear
system including star epicyclical geared architecture.
[0029] FIG. 3 is a schematic view of an example fan drive gear
system including planetary epicyclical geared architecture.
DETAILED DESCRIPTION
[0030] FIG. 1 schematically illustrates an example gas turbine
engine 20 that includes a fan section 22, a compressor section 24,
a combustor section 26 and a turbine section 28. Alternative
engines might include an augmenter section (not shown) among other
systems or features. The fan section 22 drives air along a bypass
flow path B while the compressor section 24 draws air in along a
core flow path C where air is compressed and communicated to a
combustor section 26. In the combustor section 26, air is mixed
with fuel and ignited to generate a high pressure exhaust gas
stream that expands through the turbine section 28 where energy is
extracted and utilized to drive the fan section 22 and the
compressor section 24.
[0031] Although the disclosed non-limiting embodiment depicts a
turbofan gas turbine engine, it should be understood that the
concepts described herein are not limited to use with turbofans as
the teachings may be applied to other types of turbine engines; for
example a turbine engine including a three-spool architecture in
which three spools concentrically rotate about a common axis and
where a low spool enables a low pressure turbine to drive a fan via
a gearbox, an intermediate spool that enables an intermediate
pressure turbine to drive a first compressor of the compressor
section, and a high spool that enables a high pressure turbine to
drive a high pressure compressor of the compressor section.
[0032] The example 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.
[0033] The low speed spool 30 generally includes an inner shaft 40
that connects a fan 42 and a low pressure (or first) compressor
section 44 to a low pressure (or first) turbine section 46. The
inner shaft 40 drives the fan 42 through a speed change device,
such 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 (or second)
compressor section 52 and a high pressure (or second) turbine
section 54. The inner shaft 40 and the outer shaft 50 are
concentric and rotate via the bearing systems 38 about the engine
central longitudinal axis A.
[0034] A combustor 56 is arranged between the high pressure
compressor 52 and the high pressure turbine 54. In one example, the
high pressure turbine 54 includes at least two stages to provide a
double stage high pressure turbine 54. In another example, the high
pressure turbine 54 includes only a single stage. As used herein, a
"high pressure" compressor or turbine experiences a higher pressure
than a corresponding "low pressure" compressor or turbine.
[0035] The example low pressure turbine 46 has a pressure ratio
that is greater than about 5. The pressure ratio of the example low
pressure turbine 46 is measured prior to an inlet of the low
pressure turbine 46 as related to the pressure measured at the
outlet of the low pressure turbine 46 prior to an exhaust
nozzle.
[0036] A mid-turbine frame 58 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 58 further supports
bearing systems 38 in the turbine section 28 as well as setting
airflow entering the low pressure turbine 46.
[0037] The core airflow C is compressed by the low pressure
compressor 44 then by the high pressure compressor 52 mixed with
fuel and ignited in the combustor 56 to produce high speed exhaust
gases that are then expanded through the high pressure turbine 54
and low pressure turbine 46. The mid-turbine frame 58 includes
vanes 60, which are in the core airflow path and function as an
inlet guide vane for the low pressure turbine 46. Utilizing the
vane 60 of the mid-turbine frame 58 as the inlet guide vane for low
pressure turbine 46 decreases the length of the low pressure
turbine 46 without increasing the axial length of the mid-turbine
frame 58. Reducing or eliminating the number of vanes in the low
pressure turbine 46 shortens the axial length of the turbine
section 28. Thus, the compactness of the gas turbine engine 20 is
increased and a higher power density may be achieved.
[0038] The disclosed gas turbine engine 20 in one example is a
high-bypass geared aircraft engine. In a further example, the gas
turbine engine 20 includes a bypass ratio greater than about six
(6), with an example embodiment being greater than about ten (10).
The example geared architecture 48 is an epicyclical gear train,
such as a planetary gear system, star gear system or other known
gear system, with a gear reduction ratio of greater than about
2.3.
[0039] In one disclosed embodiment, the gas turbine engine 20
includes a bypass ratio greater than about ten (10:1) and the fan
diameter is significantly larger than an outer diameter of the low
pressure compressor 44. It should be understood, however, that the
above parameters are only exemplary of one embodiment of a gas
turbine engine including a geared architecture and that the present
disclosure is applicable to other gas turbine engines.
[0040] 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
pound-mass (lbm) of fuel per hour being burned divided by
pound-force (lbf) of thrust the engine produces at that minimum
point.
[0041] "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 1.50. In another
non-limiting embodiment the low fan pressure ratio is less than
about 1.45.
[0042] "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).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.
[0043] The example gas turbine engine includes the fan 42 that
comprises in one non-limiting embodiment less than about 26 fan
blades. In another non-limiting embodiment, the fan section 22
includes less than about 20 fan blades. Moreover, in one disclosed
embodiment the low pressure turbine 46 includes no more than about
6 turbine rotors schematically indicated at 34. In another
non-limiting example embodiment the low pressure turbine 46
includes about 3 turbine rotors. A ratio between the number of fan
blades 42 and the number of low pressure turbine rotors is between
about 3.3 and about 8.6. The example low pressure turbine 46
provides the driving power to rotate the fan section 22 and
therefore the relationship between the number of turbine rotors 34
in the low pressure turbine 46 and the number of blades 42 in the
fan section 22 disclose an example gas turbine engine 20 with
increased power transfer efficiency.
[0044] The example gas turbine engine includes a lubrication system
98. The lubrication system 98 provides lubricant flow to the
rotating components of the gas turbine engine including the bearing
assemblies 38 and the geared architecture 48. The lubrication
system 98 further provides for the removal of heat generated in the
various bearing systems and the geared architecture 48.
[0045] The example lubrication system 98 includes a main system 80
that provides lubrication during normal operating conditions of the
gas turbine engine. An auxiliary system 82 is also included to
supplement operation of the main lubrication system 80. The size
and weight of the lubrication system 90 is directly related to its
capacity for removing heat from the geared architecture 48. The
greater the need for removal of heat, the larger and heavier the
lubrication system 98 becomes. The amount of heat generated by the
geared architecture 48 is therefore an important consideration in
the configuration of a fan drive gear system.
[0046] Referring to FIG. 2 with continued reference to FIG. 1, the
example geared architecture 48 is part of a fan drive gear system
70. The example geared architecture 48 comprises a gear assembly 65
that includes a sun gear 62 driven by a fan drive turbine 46. In
this example, the fan drive turbine is the low pressure turbine 46.
The sun gear 62 in turn drives intermediate gears 64 mounted on a
carrier 74 by journal bearings. The carrier 74 is grounded to the
static engine structure 36 and therefore the intermediate gears 64
do not orbit about the sun gear 62. The intermediate gears 64
intermesh and drive a ring gear 66 coupled to a fan shaft 68 to
drive the fan 42.
[0047] The gear assembly 65 is flexibly mounted such that it may be
isolated from vibrational and transient movement that could disturb
alignment between the gears 62, 64 and 66. In this example,
flexible mounts 76 support the carrier 74 and accommodate relative
movement between the gear assembly 65 and the static structure 36.
The example flexible mount 76 includes a spring rate that
accommodates deflections that occur during normal operation of the
fan drive gear system 70.
[0048] Power input through the inner shaft 40 of the fan drive
turbine 46 is transmitted through a flexible coupling 72. The
flexible coupling 72 also includes a spring rate that allows a
defined amount of deflection and misalignment such that components
of the gear assembly 65 are not driven out of alignment.
[0049] Although some relative movement is compensated by the
flexible coupling 72 and the flexible mounts 76, movement beyond a
desired limitation can detrimentally affect meshing engagement
between the gears and therefore a load limiting device 78 is
provided as part of the gear box mounting structure. The load
limiter 78 constrains movement of the gear box 65. The limiter 78
further provides a stop that reacts to unbalanced loads on the gear
box 65. Accordingly, the limiter prevents radial unbalanced loads
and/or torsional overloads from damaging the gas turbine engine
20.
[0050] The example fan drive gear system 70 is supported by a
lubrication system 98. The lubrication system 98 provides for
lubrication and cooling of the gears 62, 64 and 66 along with
bearings supporting rotation of the gears. It is desirable to
circulate lubricant as quickly as possible to maintain a desired
temperature. Power transmission efficiency through the gear box 65
is detrimentally affected by elevated temperatures.
[0051] In this example, the lubricant system 98 includes a main
system 80 that provides the desired lubricant flow through a
plurality of conduits schematically illustrated by the line 88 to
and from the gear box 65. The main oil system 80 also transmits
heat, schematically by arrows 92, away from the gear box 65 to
maintain a desired temperature.
[0052] The lubrication system 98 also includes the auxiliary oil
system 82 that supplies oil flow to the gear box 65 in response to
a temporary interruption in lubricant flow from the main oil system
80.
[0053] The efficiency of the example gear box 65 and overall geared
architecture 48 is a function of the power input, schematically
indicated by arrow 94, through the shaft 40 relative to power
output, schematically indicated by arrows 96, to the fan shaft 68.
Power input 94 compared to the amount of power output 96 is a
measure of gear box efficiency. The example gear box 65 operates at
an efficiency of greater than about 98%. In another disclosed
example the example gear box 65 operates at an efficiency greater
than about 99%.
[0054] The disclosed efficiency is a measure of the amount of power
94 that is specifically transferred to the fan shaft 68 to rotate
the fan 42. Power that is not transmitted through the gear box 65
is lost as heat and reduces the overall efficiency of the fan drive
gear system 70. Any deficit between the input power 94 and output
power 96 results in the generation of heat. Accordingly, in this
example, the deficit of between 1-2% between the input power 94 and
output power 96 generates heat. In other words, between 1% and 2%
of the input power 94 is converted to heat energy that must be
accommodated by the lubrication system 98 to maintain a working
lubricant temperature within operational limits.
[0055] The example lubricant system 98 provides for the removal of
thermal energy equal to or less than about 2% of the input power 94
from the low pressure turbine 46. In another non-limiting
embodiment of the example fan drive gear system 70, the efficiency
of the gear box 65 is greater than about 99% such that only 1% of
power input from the low pressure turbine 46 is transferred into
heat energy that must be handled by the lubricant system 98.
[0056] As appreciated, the larger the capacity for handling and
removing thermal energy, the larger and heavier the lubricant
system 98. In this example, the main oil system includes a heat
exchanger 90 that accommodates heat 92 that is generated within the
gear box 65. The heat exchanger 90 is an example of one element of
the lubrication system 98 that is scaled to the desired capacity
for removing thermal energy. As appreciated, other elements, such
as for example lubricant pumps, conduit size along with overall
lubricant quantity within the lubrication system 98 would also be
increased in size and weight to provide increased cooling capacity.
Accordingly, it is desirable to increase power transfer efficiency
to reduce required overall heat transfer capacity of lubrication
system 98.
[0057] In this example, the high efficiency of the example gear box
65 enables a relatively small and light lubricant system 98. The
example lubricant system 98 includes features that can accommodate
thermal energy generated by no more than about 2% of the input
power 94. In other words, the lubrication system 98 has an overall
maximum capacity for removing thermal energy equal to no more than
about 2% of the input power provided by the low pressure turbine
46.
[0058] Greater amounts of capacity for removal of thermal energy
results in an overall increase in the size and weight of the
lubrication system 98. Lubrication systems that are required to
remove greater than about 2% of input power 94 require larger
lubricant systems 98 that can detrimentally impact overall engine
efficiency and detract from the propulsion efficiencies provided by
the reduction in fan speed.
[0059] Referring to FIG. 3 with continued reference to FIG. 1,
another example epicyclical gear box 85 is disclosed and comprises
a planetary configuration. In a planetary configuration, planet
gears 84 are supported on a carrier 86 that is rotatable about the
engine axis A. The sun gear 62 remains driven by the inner shaft 40
and the low pressure turbine 46. The ring gear 66 is mounted to a
fixed structure 36 such that it does not rotate about the axis.
Accordingly, rotation of the sun gear 62 drives the planet gears 84
within the ring gear 66. The planet gears 84 are supported on the
rotatable carrier 86 that in turn drives the fan shaft 68. In this
configuration, the fan shaft 68 and the sun gear 62 rotate in a
common direction, while the planet gears 84 individually rotate in
a direction opposite to the sun gear 62 but collectively rotate
about the sun gear 62 in the same direction as the rotation of the
sun gear 62.
[0060] The example planetary gear box illustrated in FIG. 3
includes the ring gear 66 that is supported by flexible mount 76.
The flexible mount 76 allows some movement of the gearbox 85 to
maintain a desired alignment between meshing teeth of the gears 62,
84, 66. The limiter 78 prevents movement of the planetary gear box
85 beyond desired limits to prevent potential damage caused by
radial imbalances and/or torsional loads.
[0061] The example low pressure turbine 46 inputs power 94 to drive
the gear box 85. As in the previous example, the example gear box
85 transmits more than about 98% of the input power 94 to the fan
drive shaft 68 as output power 96. In another example, the gear box
85 transmits more than about 99% of the input power 94 to the fan
drive shaft 68 as output power 96.
[0062] The difference between the input power 94 and the output
power 96 is converted into heat energy that is removed by the
lubrication system 98. In this example, the lubrication system 98
has a capacity of removing no more heat 92 than is generated by
about 2% of the input power 94 from the low pressure turbine 46. In
another example. The lubrication system 98 has a capacity of
removing no more heat 92 than is generated by about 1% of the input
power 94. Accordingly, the efficiency provided by the example gear
box 85 enables the lubrication system 98 to be of size that does
not detract from the propulsive efficiency realized by turning the
fan section 22 and low pressure turbine 46 at separate and nearer
optimal speeds.
[0063] Accordingly the example fan drive gear system provides for
the improvement and realization of propulsive efficiencies by
limiting losses in the form of thermal energy, thereby enabling
utilization of a lower capacity and sized lubrication system.
[0064] Although an example embodiment 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 scope and content of this disclosure.
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