U.S. patent application number 14/705459 was filed with the patent office on 2015-08-20 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 | 20150233301 14/705459 |
Document ID | / |
Family ID | 50337525 |
Filed Date | 2015-08-20 |
United States Patent
Application |
20150233301 |
Kind Code |
A1 |
Sheridan; William G. ; et
al. |
August 20, 2015 |
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 section including a fan
rotatable about an axis of rotation of the gas turbine engine. A
speed reduction device is in communication with the fan. The speed
reduction device includes a star drive gear system with a star gear
ratio of at least 1.5. A fan blade tip speed of the fan is less
than 1400 fps. A bypass ratio is between about 11.0 and about
22.0.
Inventors: |
Sheridan; William G.;
(Southington, CT) ; Hasel; Karl L.; (Manchester,
CT) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
United Technologies Corporation |
Hartford |
CT |
US |
|
|
Family ID: |
50337525 |
Appl. No.: |
14/705459 |
Filed: |
May 6, 2015 |
Related U.S. Patent Documents
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Application
Number |
Filing Date |
Patent Number |
|
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PCT/US2013/061115 |
Sep 23, 2013 |
|
|
|
14705459 |
|
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61706212 |
Sep 27, 2012 |
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Current U.S.
Class: |
415/124.1 |
Current CPC
Class: |
F05D 2210/12 20130101;
F02C 7/32 20130101; F05D 2220/36 20130101; F02C 3/04 20130101; F16H
1/28 20130101; F16H 1/36 20130101; F05D 2240/40 20130101; F02C 3/06
20130101; F02C 7/00 20130101; F02C 3/107 20130101; F05D 2220/32
20130101; F02C 3/113 20130101; F05D 2260/40311 20130101; F02C 7/36
20130101 |
International
Class: |
F02C 7/32 20060101
F02C007/32; F02C 3/107 20060101 F02C003/107; F02C 7/36 20060101
F02C007/36 |
Claims
1. A gas turbine engine comprising: a fan section including a fan
rotatable about an axis of rotation of the gas turbine engine; a
speed reduction device in communication with the fan, wherein the
speed reduction device includes a star drive gear system with a
star gear ratio of at least 1.5, wherein a fan blade tip speed of
the fan is less than 1400 fps; and a bypass ratio is between about
11.0 and about 22.0.
2. The gas turbine engine of claim 1, wherein the speed reduction
device includes a star gear system gear ratio of at least 2.6.
3. The gas turbine engine of claim 2, wherein the speed reduction
device includes a system gear ratio less than or equal to 4.1.
4. The gas turbine engine of claim 1, wherein the star system
includes a sun gear, a plurality of star gears, a ring gear, and a
carrier.
5. The gas turbine engine of claim 4, wherein each of the plurality
of star gears includes at least one bearing.
6. The gas turbine engine of claim 4, wherein the carrier is fixed
from rotation.
7. The gas turbine engine of claim 4, wherein a low pressure
turbine is mechanically attached to the sun gear.
8. The gas turbine engine of claim 4, wherein a fan section is
mechanically attached to the ring gear.
9. The gas turbine engine of claim 1, wherein an input of the speed
reduction device is rotatable in a first direction and an output of
the speed reduction device is rotatable in a second direction
opposite to the first direction.
10. The gas turbine engine of claim 1, including a low pressure
turbine section in communication with the speed reduction device,
wherein the low pressure turbine section includes at least three
stages and no more than four stages.
11. The gas turbine engine of claim 10, wherein the fan blade tip
speed of the fan is greater than 1000 fps.
12. The gas turbine engine as recited in claim 1, further
comprising a low pressure turbine which is one of three turbine
rotors, and the low pressure turbine drives the fan section and the
other two of the three turbine rotors each drive a compressor
section.
13. The gas turbine engine as recited in claim 1, further
comprising a high pressure turbine, with each of a low pressure
turbine and the high pressure turbine driving a compressor
rotor.
14. The gas turbine engine as recited in claim 13, wherein the
speed reduction device is positioned intermediate a compressor
rotor driven by the low pressure turbine and the fan section.
15. The gas turbine engine as recited in claim 13, wherein the
speed reduction device is positioned intermediate the low pressure
turbine and the compressor rotor driven by the low pressure
turbine.
Description
CROSS-REFERENCE TO RELATED APPLICATIONS
[0001] This disclosure is a continuation in part of
PCT/US2013/061115 filed on Sep. 23, 2013, which claims priority to
U.S. Provisional Patent Application No. 61/706,212 filed on Sep.
27, 2012.
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 a
fan section including a fan rotatable about an axis of rotation of
the gas turbine engine. A speed reduction device is in
communication with the fan. The speed reduction device includes a
star drive gear system with a star gear ratio of at least 1.5. A
fan blade tip speed of the fan is less than 1400 fps. A bypass
ratio is between about 11.0 and about 22.0.
[0008] In a further embodiment of the above, the speed reduction
device includes a star gear system gear ratio of at least 2.6.
[0009] In a further embodiment of any of the above, the speed
reduction device includes a system gear ratio less than or equal to
4.1.
[0010] In a further embodiment of any of the above, the star system
includes a sun gear, a plurality of star gears, a ring gear, and a
carrier.
[0011] In a further embodiment of any of the above, each of the
plurality of star gears includes at least one bearing.
[0012] In a further embodiment of any of the above, the carrier is
fixed from rotation.
[0013] In a further embodiment of any of the above, a low pressure
turbine is mechanically attached to the sun gear.
[0014] In a further embodiment of any of the above, a fan section
is mechanically attached to the ring gear.
[0015] In a further embodiment of any of the above, an input of the
speed reduction device is rotatable in a first direction and an
output of the speed reduction device is rotatable in a second
direction opposite to the first direction.
[0016] In a further embodiment of any of the above, a low pressure
turbine section that is in communication with the speed reduction
device is included. The low pressure turbine section includes at
least three stages and no more than four stages.
[0017] In a further embodiment of any of the above, the fan blade
tip speed of the fan is greater than 1000 fps.
[0018] In a further embodiment of any of the above, there is a low
pressure turbine which is one of three turbine rotors. The low
pressure turbine drives the fan section and the other two of the
three turbine rotors each drive a compressor section.
[0019] In a further embodiment of any of the above, there is a high
pressure turbine. Each of a low pressure turbine and the high
pressure turbine drives a compressor rotor.
[0020] In a further embodiment of any of the above, the speed
reduction device is positioned intermediate a compressor rotor
driven by the low pressure turbine and the fan section.
[0021] In a further embodiment of any of the above, the speed
reduction device is positioned intermediate the low pressure
turbine and the compressor rotor driven by the low pressure
turbine.
BRIEF DESCRIPTION OF THE DRAWINGS
[0022] FIG. 1 illustrates a schematic, cross-sectional view of an
example gas turbine engine.
[0023] FIG. 2 illustrates a schematic view of one configuration of
a low speed spool that can be incorporated into a gas turbine
engine.
[0024] FIG. 3 illustrates a fan drive gear system that can be
incorporated into a gas turbine engine.
[0025] FIG. 4 shows another embodiment.
[0026] FIG. 5 shows yet another embodiment.
DETAILED DESCRIPTION
[0027] 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.
[0028] 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.
[0029] 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.
[0030] 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.
[0031] 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.
[0032] 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.
[0033] 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.
[0034] 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.
[0035] 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).
[0036] 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 Dl.
[0037] 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.
[0038] 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.
[0039] 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.
[0040] 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.
[0041] 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.
[0042] 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.
[0043] 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.
[0044] 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.
[0045] 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.
[0046] 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.
[0047] 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.
* * * * *