Note: Descriptions are shown in the official language in which they were submitted.
CA 02849372 2014-04-11
METHOD FOR SETTING A GEAR RATIO OF A FAN DRIVE GEAR SYSTEM OF
A GAS TURBINE ENGINE
BACKGROUND
[0001] 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.
[0002] 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.
[0003] 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.
[0004] 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.
[0005] 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
[0006] A gas turbine engine according to an exemplary aspect of the
present
disclosure includes, among other things, a fan section including a fan
rotatable about an axis
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and a speed reduction device 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.
[0007] In a further non-limiting embodiment of the foregoing gas
turbine
engine, the speed reduction device includes a star gear system gear ratio of
at least 2.6.
[0008] In a further non-limiting embodiment of either of the
foregoing gas
turbine engines, the speed reduction device includes a system gear ratio less
than or equal to
4.1.
[0009] In a further non-limiting embodiment of any of the foregoing
gas
turbine engines, a bypass ratio is included that is greater than about 6Ø
[00010] In a further non-limiting embodiment of any of the foregoing
gas
turbine engines, the bypass ratio is between about 11.0 and about 22Ø
[00011] In a further non-limiting embodiment of any of the foregoing
gas
turbine engines, the star system includes a sun gear, a plurality of star
gears, a ring gear, and a
carrier.
[00012] In a further non-limiting embodiment of any of the foregoing
gas
turbine engines, each of the plurality of star gears include at least one
bearing.
[00013] In a further non-limiting embodiment of any of the foregoing
gas
turbine engines, the carrier is fixed from rotation.
[00014] In a further non-limiting embodiment of any of the foregoing
gas
turbine engines, a low pressure turbine is mechanically attached to the sun
gear.
[00015] In a further non-limiting embodiment of any of the foregoing
gas
turbine engines, a fan section is mechanically attached to the ring gear.
[00016] In a further non-limiting embodiment of any of the foregoing
gas
turbine engines, 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.
[00017] In a further non-limiting embodiment of any of the foregoing
gas
turbine engines, a low pressure turbine section is in communication with the
speed reduction
device, The low pressure turbine section includes at least three stages and no
more than four
stages.
[00018] In a further non-limiting embodiment of any of the foregoing
gas
turbine engines, the fan blade tip speed of the fan is greater than 1000 fps.
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[00019] A method of improving performance of a gas turbine engine
according
to another exemplary aspect of the present disclosure includes, among other
things,
determining fan tip speed boundary conditions for at least one fan blade of a
fan section and
determining rotor boundary conditions for a rotor of a low pressure turbine,
The stress level
utilizes constraints in the rotor of the low pressure turbine and the at least
one fan blade to
determine if the rotary speed of the fan section and the low pressure turbine
will meet a
desired number of operating cycles.
[00020] In a further non-limiting embodiment of the foregoing method,
a speed
reduction device connects the fan section and the low pressure turbine and
includes a star
gear ratio of at least about 1.5 and no more than about 4.1.
[00021] In a further non-limiting embodiment of either of the
foregoing
methods, a fan pressure ratio is below 1.7.
[00022] In a further non-limiting embodiment of any of the foregoing
methods,
a fan pressure ratio is below 1.48.
[00023] In a further non-limiting embodiment of any of the foregoing
methods,
a bypass ratio is between about 11 and about 22.
[00024] In a further non-limiting embodiment of any of the foregoing
methods,
a fan blade tip speed of the at least one fan blade is less than 1400 fps.
[00025] In a further non-limiting embodiment of any of the foregoing
methods,
if a stress level in the rotor or the at least one fan blade is too high to
meet a desired number
of operating cycles, a gear ratio of a gear reduction device is lowered and
the number of
stages of the low pressure turbine is increased.
[00026] In a further non-limiting embodiment of any of the foregoing
methods,
if a stress level in the rotor or the at least one fan blade is too high to
meet a desired number
of operating cycles, a gear ratio of a gear reduction device is lowered and an
annular area of
the low pressure turbine is increased.
[00027] The various features and advantages of this disclosure will
become
apparent to those skilled in the art from the following detailed description.
The drawings that
accompany the detailed description can be briefly described as follows.
BRIEF DESCRIPTION OF THE DRAWINGS
[00028] Figure 1 illustrates a schematic, cross-sectional view of an
example gas
turbine engine.
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[00029] Figure 2 illustrates a schematic view of one configuration of
a low
speed spool that can be incorporated into a gas turbine engine.
[00030] Figure 3 illustrates a fan drive gear system that can be
incorporated
into a gas turbine engine.
DETAILED DESCRIPTION
[00031] Figure 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.
[00032] 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.
[00033] 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 Figures 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.
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[00034] 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.
[00035] 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.
[00036] 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.
[00037] 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,
CA 02849372 2014-04-11
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.
[00038] 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.
[00039] 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 R) / (518.7 R)] 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).
[00040] Figure 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.
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[00041] Figure 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.
[00042] The sun gear 52 receives an input from the low pressure
turbine 39
(see Figure 2) and rotates in the first direction DI thereby turning the
plurality of star gears
56 in a second direction D2 that is opposite of the first direction D 1 .
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 Figure 2) in the second
direction D2.
[00043] 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.
[00044] 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.
[00045] 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
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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.
[00046] 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.
[00047] 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.
[00048] 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.
[00049] 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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