U.S. patent application number 15/875656 was filed with the patent office on 2018-06-07 for method for setting a gear ratio of a fan drive gear system of a gas turbine engine.
The applicant listed for this patent is United Technologies Corporation. Invention is credited to Karl L. Hasel, William G. Sheridan.
Application Number | 20180156135 15/875656 |
Document ID | / |
Family ID | 50337525 |
Filed Date | 2018-06-07 |
United States Patent
Application |
20180156135 |
Kind Code |
A1 |
Sheridan; William G. ; et
al. |
June 7, 2018 |
METHOD FOR SETTING A GEAR RATIO OF A FAN DRIVE GEAR SYSTEM OF A GAS
TURBINE ENGINE
Abstract
A gas turbine engine includes a fan with a plurality of fan
blades rotatable about an engine centerline longitudinal axis. The
fan has a low corrected fan tip speed less than 1400 ft/sec. A
bypass ratio is greater than 13 and less than 20. A fan pressure
ratio less than 1.48. A speed reduction device comprises a gear
system with a gear ratio of at least 2.6 and less than or equal to
4.1. A low and high pressure turbine is in communication with a
first and second shaft, respectively. The low pressure turbine
includes at least three stages and no more than four stages. The
high pressure turbine includes two stages. The gear ratio is
configured such that in operation the fan blade does not exceed a
fan tip speed boundary condition or a second stress level. A low
pressure turbine rotor does not exceed a first stress level.
Inventors: |
Sheridan; William G.;
(Southington, CT) ; Hasel; Karl L.; (Manchester,
CT) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
United Technologies Corporation |
Farmington |
CT |
US |
|
|
Family ID: |
50337525 |
Appl. No.: |
15/875656 |
Filed: |
January 19, 2018 |
Related U.S. Patent Documents
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Application
Number |
Filing Date |
Patent Number |
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14705577 |
May 6, 2015 |
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15875656 |
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PCT/US2013/061115 |
Sep 23, 2013 |
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14705577 |
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13758075 |
Feb 4, 2013 |
8753065 |
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PCT/US2013/061115 |
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Current U.S.
Class: |
1/1 |
Current CPC
Class: |
F16H 1/28 20130101; F16H
1/36 20130101; F02C 7/36 20130101; F05D 2220/36 20130101; F02C 3/04
20130101; F02C 3/06 20130101; F05D 2260/40311 20130101; F02C 7/00
20130101; F05D 2220/32 20130101; F02C 7/32 20130101; F02C 3/107
20130101; F02C 3/113 20130101; F05D 2240/40 20130101; F05D 2210/12
20130101 |
International
Class: |
F02C 7/36 20060101
F02C007/36; F02C 7/32 20060101 F02C007/32; F02C 3/06 20060101
F02C003/06; F16H 1/36 20060101 F16H001/36 |
Claims
1. A gas turbine engine comprising: an engine centerline
longitudinal axis; a fan section including a fan with a plurality
of fan blades and rotatable about the engine centerline
longitudinal axis, wherein the fan has a low corrected fan tip
speed less than 1400 ft/sec, wherein the low corrected fan tip
speed is an actual fan tip speed at an ambient temperature divided
by [(Tram .degree. R)/(518.7.degree. R)].sup.0.5, where T
represents the ambient temperature in degrees Rankine; a bypass
ratio greater than 13 and less than 20; a fan pressure ratio less
than 1.48, wherein the fan pressure ratio is measured across a fan
blade alone; a speed reduction device comprising a gear system with
a gear ratio of at least 2.6 and less than or equal to 4.1; a
plurality of bearing systems; a low pressure turbine in
communication with a first shaft; and a high pressure turbine in
communication a second shaft; wherein the first shaft and second
shaft are concentric and mounted via at least one of the plurality
of bearing systems for rotation about the engine centerline
longitudinal axis, and the first shaft is in communication with the
fan through the speed reduction device; wherein the low pressure
turbine includes at least three stages and no more than four
stages; wherein the high pressure turbine includes two stages; and
wherein the low pressure turbine includes at least one rotor
constrained by a first stress level, at least one of the plurality
of fan blades of the fan constrained by a second stress level and
having a fan tip speed boundary condition, and the gear ratio is
configured such that in operation the fan blade does not exceed the
fan tip speed boundary condition or the second stress level, and
the low pressure turbine rotor does not exceed the first stress
level.
2. The gas turbine engine of claim 1, wherein the low pressure
turbine includes four stages.
3. The gas turbine engine of claim 2, wherein the gear system is a
star gear system with a ring gear, and a sun gear, wherein the gear
ratio is determined by measuring a diameter of the ring gear and
dividing that diameter by the diameter of the sun gear.
4. The gas turbine engine of claim 2, wherein the fan pressure
ratio is less than 1.38.
5. The gas turbine engine of claim 3, wherein the fan pressure
ratio is less than 1.38.
6. A gas turbine engine comprising: an engine centerline
longitudinal axis; a fan section including a fan with a plurality
of fan blades and rotatable about the engine centerline
longitudinal axis, wherein the fan has a low corrected fan tip
speed less than 1400 ft/sec, wherein the low corrected fan tip
speed is an actual fan tip speed at an ambient temperature divided
by [(Tram .degree. R)/(518.7.degree. R)].sup.0.5, where T
represents the ambient temperature in degrees Rankine; a bypass
ratio greater than 11.0 and less than 22.0; a fan pressure ratio
less than 1.38, wherein the fan pressure ratio is measured across a
fan blade alone; a speed reduction device comprising a gear system;
a plurality of bearing systems; a low pressure turbine in
communication with a first shaft; and a high pressure turbine in
communication with a second shaft; wherein the first shaft and
second shaft are concentric and mounted via at least one of the
plurality of bearing systems for rotation about the engine
centerline longitudinal axis, and the first shaft is in
communication with the fan through the speed reduction device;
wherein the high pressure turbine includes two stages; and wherein
the low pressure turbine includes at least one rotor constrained by
a first stress level, at least one of the plurality of fan blades
of the fan constrained by a second stress level and having a fan
tip speed boundary condition, and the gear system is configured
such that in operation the fan blade does not exceed the fan tip
speed boundary condition or the second stress level, and the low
pressure turbine rotor does not exceed the first stress level.
7. The gas turbine engine of claim 6, wherein the low pressure
turbine includes five stages.
8. The gas turbine engine of claim 6, further comprising a low
pressure compressor including three stages, wherein the low
pressure turbine drives the low pressure compressor.
9. The gas turbine engine of claim 6, wherein the low pressure
turbine includes at least three stages and no more than four
stages.
10. The gas turbine engine of claim 6, wherein the speed reduction
device includes a gear ratio less than or equal to 4.1.
11. The gas turbine engine of claim 10, wherein the gear ratio is
greater than or equal to 2.6.
12. The gas turbine engine of claim 9, wherein the low pressure
turbine includes four stages.
13. The gas turbine engine of claim 12, further comprising a
mid-turbine frame arranged between the high pressure turbine and
the low pressure turbine, wherein the mid-turbine frame supports at
least one bearing system.
14. The gas turbine engine of claim 13, wherein the mid-turbine
frame includes one or more airfoils that extend in a flow path.
15. The gas turbine engine of claim 12, further comprising a low
pressure compressor including three stages, wherein the low
pressure turbine drives the low pressure compressor.
16. The gas turbine engine of claim 15, further comprising a high
pressure compressor including eight stages, wherein the high
pressure turbine drives the high pressure compressor.
17. The gas turbine engine of claim 12, the star gear system
further comprising five intermediate gears.
18. The gas turbine engine of claim 17, wherein the gear system is
a star gear system with a ring gear, a sun gear, and a star gear
ratio, and the star gear ratio is determined by measuring a
diameter of the ring gear and dividing that diameter by the
diameter of the sun gear.
19. The gas turbine engine of claim 17, further comprising a
mid-turbine frame arranged between the high pressure turbine and
the low pressure turbine, wherein the mid-turbine frame supports at
least one bearing system and includes one or more airfoils.
20. A gas turbine engine comprising: an engine centerline
longitudinal axis; a fan section including a fan with a plurality
of fan blades and rotatable about the engine centerline
longitudinal axis; a low corrected fan tip speed less than 1400
ft/sec, wherein the low corrected fan tip speed is an actual fan
tip speed at an ambient temperature divided by [(Tram .degree.
R)/(518.7.degree. R)].sup.0.5, where T represents the ambient
temperature in degrees Rankine; a bypass ratio of greater than
11.0; a speed reduction device comprising a gear system; a
plurality of bearing systems; a low pressure turbine in
communication with a first shaft and includes a pressure ratio
greater than about 5:1, the low pressure turbine includes an inlet
having an inlet pressure, and an outlet having an outlet pressure,
and the pressure ratio of the low pressure turbine is a ratio of
the inlet pressure to the outlet pressure; and a high pressure
turbine in communication with a second shaft; wherein the first
shaft and second shaft are concentric and mounted via at least one
of the plurality of bearing systems for rotation about the engine
centerline longitudinal axis, and the first shaft is in
communication with the fan through the speed reduction device;
wherein the low pressure turbine includes at least three stages and
no more than four stages; and wherein the low pressure turbine
includes at least one rotor constrained by a first stress level, at
least one of the plurality of fan blades of the fan constrained by
a second stress level and having a fan tip speed boundary
condition, and the gear system is configured such that in operation
the fan blade does not exceed the fan tip speed boundary condition
or the second stress level, and the low pressure turbine rotor does
not exceed the first stress level.
21. The gas turbine engine of claim 20, wherein the high pressure
turbine includes two stages.
22. The gas turbine engine of claim 21, wherein the low pressure
turbine includes three stages.
23. The gas turbine engine of claim 21, wherein the gear system has
a gear ratio of less than or equal to 4.1.
24. The gas turbine engine of claim 23, wherein the gear reduction
ratio is at least 2.6.
25. The gas turbine engine of claim 24, wherein the gear system is
a star gear system with a ring gear, and a sun gear, wherein the
gear ratio is determined by measuring a diameter of the ring gear
and dividing that diameter by the diameter of the sun gear.
26. The gas turbine engine of claim 24, wherein the low pressure
turbine includes four stages.
27. A gas turbine engine comprising: an engine centerline
longitudinal axis; a fan section including a fan with a plurality
of fan blades and rotatable about the engine centerline
longitudinal axis; a low corrected fan tip speed less than 1400
ft/sec, wherein the low corrected fan tip speed is an actual fan
tip speed at an ambient temperature divided by [(Tram .degree.
R)/(518.7.degree. R)].sup.0.5, where T represents the ambient
temperature in degrees Rankine; a bypass ratio greater than 11 and
less than 22; a speed reduction device comprising a gear system
with a gear ratio; a plurality of bearing systems; a low pressure
turbine in communication with a first shaft; and a high pressure
turbine in communication with a second shaft; wherein the first
shaft and second shaft are concentric and mounted via at least one
of the plurality of bearing systems for rotation about the engine
centerline longitudinal axis, and the first shaft is in
communication with the fan through the speed reduction device;
wherein the low pressure turbine includes four stages; and wherein
the low pressure turbine includes at least one rotor constrained by
a first stress level, at least one of the plurality of fan blades
of the fan constrained by a second stress level and having a fan
tip speed boundary condition, and the gear ratio is configured such
that in operation the fan blade does not exceed the fan tip speed
boundary condition or the second stress level, and the low pressure
turbine rotor does not exceed the first stress level.
28. The gas turbine engine of claim 27, wherein the high pressure
turbine includes two stages.
29. The gas turbine engine of claim 27, wherein the gear reduction
ratio is at least 2.6 and less than or equal to 4.1.
30. The gas turbine engine of claim 29, wherein the low pressure
turbine has a pressure ratio greater than 5:1, the low pressure
turbine includes an inlet having an inlet pressure, and an outlet
having an outlet pressure, and the pressure ratio of the low
pressure turbine is a ratio of the inlet pressure to the outlet
pressure.
Description
CROSS-REFERENCE TO RELATED APPLICATIONS
[0001] This disclosure is a continuation of U.S. application Ser.
No. 14/705,577 filed May 6, 2015, which is a continuation in part
of PCT/US2013/061115 filed on Sep. 23, 2013, which is a
continuation of U.S. application Ser. No. 13/758,075 filed Feb. 4,
2013, which is now U.S. Pat. No. 8,753,065 issued on Jun. 17,
2014.
BACKGROUND
[0002] This disclosure relates to a gas turbine engine, and more
particularly to a method for setting a gear ratio of a fan drive
gear system of a gas turbine engine.
[0003] A gas turbine engine may include a fan section, a compressor
section, a combustor section, and a turbine section. Air entering
the compressor section is compressed and delivered into the
combustor 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. Among other variations, the compressor section
can include low and high pressure compressors, and the turbine
section can include low and high pressure turbines.
[0004] Typically, a high pressure turbine drives a high pressure
compressor through an outer shaft to form a high spool, and a low
pressure turbine drives a low pressure compressor through an inner
shaft to form a low spool. The fan section may also be driven by
the inner shaft. A direct drive gas turbine engine may include a
fan section driven by the low spool such that a low pressure
compressor, low pressure turbine, and fan section rotate at a
common speed in a common direction.
[0005] A speed reduction device, which may be a fan drive gear
system or other mechanism, may be utilized to drive the fan section
such that the fan section may rotate at a speed different than the
turbine section. This allows for an overall increase in propulsive
efficiency of the engine. In such engine architectures, a shaft
driven by one of the turbine sections provides an input to the
speed reduction device that drives the fan section at a reduced
speed such that both the turbine section and the fan section can
rotate at closer to optimal speeds.
[0006] Although gas turbine engines utilizing speed change
mechanisms are generally known to be capable of improved propulsive
efficiency relative to conventional engines, gas turbine engine
manufacturers continue to seek further improvements to engine
performance including improvements to thermal, transfer and
propulsive efficiencies.
SUMMARY
[0007] In one exemplary embodiment, a gas turbine engine includes
an engine centerline longitudinal axis. A fan section includes a
fan with a plurality of fan blades and rotatable about the engine
centerline longitudinal axis. The fan has a low corrected fan tip
speed less than 1400 ft/sec. The 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. T represents the ambient
temperature in degrees Rankine. A bypass ratio is greater than 13
and less than 20. A fan pressure ratio less than 1.48. The fan
pressure ratio is measured across a fan blade alone. A speed
reduction device comprises a gear system with a gear ratio of at
least 2.6 and less than or equal to 4.1. There is a plurality of
bearing systems. A low pressure turbine is in communication with a
first shaft. A high pressure turbine is in communication a second
shaft. The first shaft and second shaft are concentric and mounted
via at least one of the plurality of bearing systems for rotation
about the engine centerline longitudinal axis. The first shaft is
in communication with the fan through the speed reduction device.
The low pressure turbine includes at least three stages and no more
than four stages. The high pressure turbine includes two stages.
The low pressure turbine includes at least one rotor constrained by
a first stress level. At least one of the plurality of fan blades
of the fan is constrained by a second stress level and has a fan
tip speed boundary condition. The gear ratio is configured such
that in operation the fan blade does not exceed the fan tip speed
boundary condition or the second stress level. The low pressure
turbine rotor does not exceed the first stress level.
[0008] In a further embodiment of any of the above, the low
pressure turbine includes four stages.
[0009] In a further embodiment of any of the above, the gear system
is a star gear system with a ring gear and a sun gear. The gear
ratio is determined by measuring a diameter of the ring gear and
dividing that diameter by the diameter of the sun gear.
[0010] In a further embodiment of any of the above, the fan
pressure ratio is less than 1.38.
[0011] In a further embodiment of any of the above, the fan
pressure ratio is less than 1.38.
[0012] In another exemplary embodiment, a gas turbine engine
includes an engine centerline longitudinal axis. A fan section
includes a fan with a plurality of fan blades and rotatable about
the engine centerline longitudinal axis. The fan has a low
corrected fan tip speed less than 1400 ft/sec. The low corrected
fan tip speed is an actual fan tip speed at an ambient temperature
divided by [(Tram .degree. R)/(518.7.degree. R)].sup.0.5. T
represents the ambient temperature in degrees Rankine. A bypass
ratio is greater than 11.0 and less than 22.0. A fan pressure ratio
is less than 1.38. The fan pressure ratio is measured across a fan
blade alone. A speed reduction device comprises a gear system.
There is a plurality of bearing systems. A low pressure turbine is
in communication with a first shaft. A high pressure turbine is in
communication with a second shaft. The first shaft and second shaft
are concentric and mounted via at least one of the plurality of
bearing systems for rotation about the engine centerline
longitudinal axis. The first shaft is in communication with the fan
through the speed reduction device. The high pressure turbine
includes two stages. The low pressure turbine includes at least one
rotor constrained by a first stress level. At least one of the
plurality of fan blades of the fan is constrained by a second
stress level and has a fan tip speed boundary condition. The gear
system is configured such that in operation the fan blade does not
exceed the fan tip speed boundary condition or the second stress
level. The low pressure turbine rotor does not exceed the first
stress level.
[0013] In a further embodiment of any of the above, the low
pressure turbine includes five stages.
[0014] In a further embodiment of any of the above, a low pressure
compressor including three stages. The low pressure turbine drives
the low pressure compressor.
[0015] In a further embodiment of any of the above, the low
pressure turbine includes at least three stages and no more than
four stages.
[0016] In a further embodiment of any of the above, the speed
reduction device includes a gear ratio less than or equal to
4.1.
[0017] In a further embodiment of any of the above, the gear ratio
is greater than or equal to 2.6.
[0018] In a further embodiment of any of the above, the low
pressure turbine includes four stages.
[0019] In a further embodiment of any of the above, a mid-turbine
frame is arranged between the high pressure turbine and the low
pressure turbine. The mid-turbine frame supports at least one
bearing system.
[0020] In a further embodiment of any of the above, the mid-turbine
frame includes one or more airfoils that extend in a flow path.
[0021] In a further embodiment of any of the above, a low pressure
compressor including three stages. The low pressure turbine drives
the low pressure compressor.
[0022] In a further embodiment of any of the above, a high pressure
compressor includes eight stages. The high pressure turbine drives
the high pressure compressor.
[0023] In a further embodiment of any of the above, the star gear
system further comprising five intermediate gears.
[0024] In a further embodiment of any of the above, the gear system
is a star gear system with a ring gear, a sun gear, and a star gear
ratio. The star gear ratio is determined by measuring a diameter of
the ring gear and dividing that diameter by the diameter of the sun
gear.
[0025] In a further embodiment of any of the above, a mid-turbine
frame is arranged between the high pressure turbine and the low
pressure turbine. The mid-turbine frame supports at least one
bearing system and includes one or more airfoils.
[0026] In another exemplary embodiment, a gas turbine engine
includes an engine centerline longitudinal axis. A fan section
includes a fan with a plurality of fan blades and rotatable about
the engine centerline longitudinal axis and has a low corrected fan
tip speed less than 1400 ft/sec. The low corrected fan tip speed is
an actual fan tip speed at an ambient temperature divided by [(Tram
.degree. R)/(518.7.degree. R)].sup.0.5. T represents the ambient
temperature in degrees Rankine. A bypass ratio of greater than
11.0. A speed reduction device comprising a gear system. There is a
plurality of bearing systems. A low pressure turbine is in
communication with a first shaft and includes a pressure ratio
greater than about 5:1. The low pressure turbine includes an inlet
that has an inlet pressure and an outlet having an outlet pressure.
The pressure ratio of the low pressure turbine is a ratio of the
inlet pressure to the outlet pressure. A high pressure turbine in
communication with a second shaft. The first shaft and second shaft
are concentric and mounted via at least one of the plurality of
bearing systems for rotation about the engine centerline
longitudinal axis. The first shaft is in communication with the fan
through the speed reduction device. The low pressure turbine
includes at least three stages and no more than four stages. The
low pressure turbine includes at least one rotor constrained by a
first stress level. At least one of the plurality of fan blades of
the fan is constrained by a second stress level and has a fan tip
speed boundary condition. The gear system is configured such that
in operation the fan blade does not exceed the fan tip speed
boundary condition or the second stress level. The low pressure
turbine rotor does not exceed the first stress level.
[0027] In a further embodiment of any of the above, the high
pressure turbine includes two stages.
[0028] In a further embodiment of any of the above, the low
pressure turbine includes three stages.
[0029] In a further embodiment of any of the above, the gear system
has a gear ratio of less than or equal to 4.1.
[0030] In a further embodiment of any of the above, the gear
reduction ratio is at least 2.6.
[0031] In a further embodiment of any of the above, the gear system
is a star gear system with a ring gear and a sun gear. The gear
ratio is determined by measuring a diameter of the ring gear and
dividing that diameter by the diameter of the sun gear.
[0032] In a further embodiment of any of the above, the low
pressure turbine includes four stages.
[0033] In another exemplary embodiment, a gas turbine engine
includes an engine centerline longitudinal axis. A fan section
includes a fan with a plurality of fan blades and rotatable about
the engine centerline longitudinal axis and has a low corrected fan
tip speed less than 1400 ft/sec. The low corrected fan tip speed is
an actual fan tip speed at an ambient temperature divided by [(Tram
.degree. R)/(518.7.degree. R)].sup.0.5. T represents the ambient
temperature in degrees Rankine. A bypass ratio is greater than 11
and less than 22. A speed reduction device comprises a gear system
with a gear ratio. A plurality of bearing systems. A low pressure
turbine is in communication with a first shaft and a high pressure
turbine in communication with a second shaft. The first shaft and
second shaft are concentric and mounted via at least one of the
plurality of bearing systems for rotation about the engine
centerline longitudinal axis. The first shaft is in communication
with the fan through the speed reduction device. The low pressure
turbine includes four stages. The low pressure turbine includes at
least one rotor constrained by a first stress level. At least one
of the plurality of fan blades of the fan is constrained by a
second stress level and has a fan tip speed boundary condition. The
gear ratio is configured such that in operation the fan blade does
not exceed the fan tip speed boundary condition or the second
stress level. The low pressure turbine rotor does not exceed the
first stress level.
[0034] In a further embodiment of any of the above, the high
pressure turbine includes two stages.
[0035] In a further embodiment of any of the above, the gear
reduction ratio is at least 2.6 and less than or equal to 4.1.
[0036] In a further embodiment of any of the above, the low
pressure turbine has a pressure ratio greater than 5:1. The low
pressure turbine includes an inlet that has an inlet pressure and
an outlet that has an outlet pressure. The pressure ratio of the
low pressure turbine is a ratio of the inlet pressure to the outlet
pressure.
BRIEF DESCRIPTION OF THE DRAWINGS
[0037] FIG. 1 illustrates a schematic, cross-sectional view of an
example gas turbine engine.
[0038] FIG. 2 illustrates a schematic view of one configuration of
a low speed spool that can be incorporated into a gas turbine
engine.
[0039] FIG. 3 illustrates a fan drive gear system that can be
incorporated into a gas turbine engine.
[0040] FIG. 4 shows another embodiment.
[0041] FIG. 5 shows yet another embodiment.
DETAILED DESCRIPTION
[0042] FIG. 1 schematically illustrates a gas turbine engine 20.
The exemplary gas turbine engine 20 is a two-spool turbofan engine
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, while the compressor section 24 drives air along a
core flow path C for compression and communication into the
combustor section 26. The hot combustion gases generated in the
combustor section 26 are expanded 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 two-spool turbofan
engines and these teachings could extend to other types of engines,
including but not limited to, three-spool engine architectures.
[0043] The exemplary gas turbine engine 20 generally includes a low
speed spool 30 and a high speed spool 32 mounted for rotation about
an engine centerline longitudinal axis A. The low speed spool 30
and the high speed spool 32 may be mounted relative to an engine
static structure 33 via several bearing systems 31. It should be
understood that other bearing systems 31 may alternatively or
additionally be provided, and the location of bearing systems 31
may be varied as appropriate to the application.
[0044] The low speed spool 30 generally includes an inner shaft 34
that interconnects a fan 36, a low pressure compressor 38 and a low
pressure turbine 39. The inner shaft 34 can be connected to the fan
36 through a speed change mechanism, which in exemplary gas turbine
engine 20 is illustrated as a geared architecture 45, such as a fan
drive gear system 50 (see FIGS. 2 and 3). The speed change
mechanism drives the fan 36 at a lower speed than the low speed
spool 30. The high speed spool 32 includes an outer shaft 35 that
interconnects a high pressure compressor 37 and a high pressure
turbine 40. In this embodiment, the inner shaft 34 and the outer
shaft 35 are supported at various axial locations by bearing
systems 31 positioned within the engine static structure 33.
[0045] A combustor 42 is arranged in exemplary gas turbine 20
between the high pressure compressor 37 and the high pressure
turbine 40. A mid-turbine frame 44 may be arranged generally
between the high pressure turbine 40 and the low pressure turbine
39. The mid-turbine frame 44 can support one or more bearing
systems 31 of the turbine section 28. The mid-turbine frame 44 may
include one or more airfoils 46 that extend within the core flow
path C. 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 50 may be varied. For
example, gear system 50 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 50.
[0046] The inner shaft 34 and the outer shaft 35 are concentric and
rotate via the bearing systems 31 about the engine centerline
longitudinal axis A, which is co-linear with their longitudinal
axes. The core airflow is compressed by the low pressure compressor
38 and the high pressure compressor 37, is mixed with fuel and
burned in the combustor 42, and is then expanded over the high
pressure turbine 40 and the low pressure turbine 39. The high
pressure turbine 40 and the low pressure turbine 39 rotationally
drive the respective high speed spool 32 and the low speed spool 30
in response to the expansion.
[0047] In a non-limiting embodiment, the gas turbine engine 20 is a
high-bypass geared aircraft engine. In a further example, the gas
turbine engine 20 bypass ratio is greater than about six (6:1). The
geared architecture 45 can include an epicyclic gear train, such as
a planetary gear system, a star gear system, or other gear system.
The geared architecture 45 enables operation of the low speed spool
30 at higher speeds, which can enable an increase in the
operational efficiency of the low pressure compressor 38 and low
pressure turbine 39, and render increased pressure in a fewer
number of stages.
[0048] The pressure ratio of the low pressure turbine 39 can be
pressure measured prior to the inlet of the low pressure turbine 39
as related to the pressure at the outlet of the low pressure
turbine 39 and prior to an exhaust nozzle of the gas turbine engine
20. In one non-limiting embodiment, the bypass ratio of the gas
turbine engine 20 is greater than about ten (10:1), the fan
diameter is significantly larger than that of the low pressure
compressor 38, and the low pressure turbine 39 has a pressure ratio
that is greater than about five (5:1). In another non-limiting
embodiment, the bypass ratio is greater than 11 and less than 22,
or greater than 13 and less than 20. It should be understood,
however, that the above parameters are only exemplary of a geared
architecture engine or other engine using a speed change mechanism,
and that the present disclosure is applicable to other gas turbine
engines, including direct drive turbofans. In one non-limiting
embodiment, the low pressure turbine 39 includes at least one stage
and no more than eight stages, or at least three stages and no more
than six stages. In another non-limiting embodiment, the low
pressure turbine 39 includes at least three stages and no more than
four stages.
[0049] In this embodiment of the exemplary gas turbine engine 20, a
significant amount of thrust is provided by the bypass flow path B
due to the high bypass ratio. The fan section 22 of the gas turbine
engine 20 is designed for a particular flight condition--typically
cruise at about 0.8 Mach and about 35,000 feet. This flight
condition, with the gas turbine engine 20 at its best fuel
consumption, is also known as bucket cruise Thrust Specific Fuel
Consumption (TSFC). TSFC is an industry standard parameter of fuel
consumption per unit of thrust.
[0050] Fan Pressure Ratio is the pressure ratio across a blade of
the fan section 22 without the use of a Fan Exit Guide Vane system.
The low Fan Pressure Ratio according to one non-limiting embodiment
of the example gas turbine engine 20 is less than 1.45. In another
non-limiting embodiment of the example gas turbine engine 20, the
Fan Pressure Ratio is less than 1.38 and greater than 1.25. In
another non-limiting embodiment, the fan pressure ratio is less
than 1.48. In another non-limiting embodiment, the fan pressure
ratio is less than 1.52. In another non-limiting embodiment, the
fan pressure ratio is less than 1.7. Low Corrected Fan Tip Speed is
the actual fan tip speed divided by an industry standard
temperature correction of [(Tram .degree. R)/(518.7.degree.
R)].sup.0.5, where T represents the ambient temperature in degrees
Rankine. The Low Corrected Fan Tip Speed according to one
non-limiting embodiment of the example gas turbine engine 20 is
less than about 1150 fps (351 m/s). The Low Corrected Fan Tip Speed
according to another non-limiting embodiment of the example gas
turbine engine 20 is less than about 1400 fps (427 m/s). The Low
Corrected Fan Tip Speed according to another non-limiting
embodiment of the example gas turbine engine 20 is greater than
about 1000 fps (305 m/s).
[0051] FIG. 2 schematically illustrates the low speed spool 30 of
the gas turbine engine 20. The low speed spool 30 includes the fan
36, the low pressure compressor 38, and the low pressure turbine
39. The inner shaft 34 interconnects the fan 36, the low pressure
compressor 38, and the low pressure turbine 39. The inner shaft 34
is connected to the fan 36 through the fan drive gear system 50. In
this embodiment, the fan drive gear system 50 provides for
counter-rotation of the low pressure turbine 39 and the fan 36. For
example, the fan 36 rotates in a first direction D1, whereas the
low pressure turbine 39 rotates in a second direction D2 that is
opposite of the first direction D1.
[0052] FIG. 3 illustrates one example embodiment of the fan drive
gear system 50 incorporated into the gas turbine engine 20 to
provide for counter-rotation of the fan 36 and the low pressure
turbine 39. In this embodiment, the fan drive gear system 50
includes a star gear system with a sun gear 52, a ring gear 54
disposed about the sun gear 52, and a plurality of star gears 56
having journal bearings 57 positioned between the sun gear 52 and
the ring gear 54. A fixed carrier 58 carries and is attached to
each of the star gears 56. In this embodiment, the fixed carrier 58
does not rotate and is connected to a grounded structure 55 of the
gas turbine engine 20.
[0053] The sun gear 52 receives an input from the low pressure
turbine 39 (see FIG. 2) and rotates in the first direction D1
thereby turning the plurality of star gears 56 in a second
direction D2 that is opposite of the first direction D1. Movement
of the plurality of star gears 56 is transmitted to the ring gear
54 which rotates in the second direction D2 opposite from the first
direction D1 of the sun gear 52. The ring gear 54 is connected to
the fan 36 for rotating the fan 36 (see FIG. 2) in the second
direction D2.
[0054] A star system gear ratio of the fan drive gear system 50 is
determined by measuring a diameter of the ring gear 54 and dividing
that diameter by a diameter of the sun gear 52. In one embodiment,
the star system gear ratio of the geared architecture 45 is between
1.5 and 4.1. In another embodiment, the system gear ratio of the
fan drive gear system 50 is between 2.6 and 4.1. When the star
system gear ratio is below 1.5, the sun gear 52 is relatively much
larger than the star gears 56. This size differential reduces the
load the star gears 56 are capable of carrying because of the
reduction in size of the star gear journal bearings 57. When the
star system gear ratio is above 4.1, the sun gear 52 may be much
smaller than the star gears 56. This size differential increases
the size of the star gear 56 journal bearings 57 but reduces the
load the sun gear 52 is capable of carrying because of its reduced
size and number of teeth. Alternatively, roller bearings could be
used in place of journal bearings 57.
[0055] Improving performance of the gas turbine engine 20 begins by
determining fan tip speed boundary conditions for at least one fan
blade of the fan 36 to define the speed of the tip of the fan
blade. The maximum fan diameter is determined based on the
projected fuel burn derived from balancing engine efficiency, mass
of air through the bypass flow path B, and engine weight increase
due to the size of the fan blades.
[0056] Boundary conditions are then determined for the rotor of
each stage of the low pressure turbine 39 to define the speed of
the rotor tip and to define the size of the rotor and the number of
stages in the low pressure turbine 39 based on the efficiency of
low pressure turbine 39 and the low pressure compressor 38.
[0057] Constraints regarding stress levels in the rotor and the fan
blade are utilized to determine if the rotary speed of the fan 36
and the low pressure turbine 39 will meet a desired number of
operating life cycles. If the stress levels in the rotor or the fan
blade are too high, the gear ratio of the fan drive gear system 50
can be lowered and the number of stages of the low pressure turbine
39 or annular area of the low pressure turbine 39 can be
increased.
[0058] FIG. 4 shows an embodiment 100, wherein there is a fan drive
turbine 108 driving a shaft 106 to in turn drive a fan rotor 102. A
gear reduction 104 may be positioned between the fan drive turbine
108 and the fan rotor 102. This gear reduction 104 may be
structured and operate like the geared architecture 45 disclosed
above. A compressor rotor 110 is driven by an intermediate pressure
turbine 112, and a second stage compressor rotor 114 is driven by a
turbine rotor 116. A combustion section 118 is positioned
intermediate the compressor rotor 114 and the turbine section
116.
[0059] FIG. 5 shows yet another embodiment 200 wherein a fan rotor
202 and a first stage compressor 204 rotate at a common speed. The
gear reduction 206 (which may be structured as the geared
architecture 45 disclosed above) is intermediate the compressor
rotor 204 and a shaft 208 which is driven by a low pressure turbine
section.
[0060] Although the different non-limiting embodiments are
illustrated as having specific components, the embodiments of this
disclosure are not limited to those particular combinations. It is
possible to use some of the components or features from any of the
non-limiting embodiments in combination with features or components
from any of the other non-limiting embodiments.
[0061] 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 and illustrated in these
exemplary embodiments, other arrangements could also benefit from
the teachings of this disclosure.
[0062] The foregoing description shall be interpreted as
illustrative and not in any limiting sense. A worker of ordinary
skill in the art would understand that certain modifications could
come within the scope of this disclosure. For these reasons, the
following claim should be studied to determine the true scope and
content of this disclosure.
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