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
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 6.0.
[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 22.0.
[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.
[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.
[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 3 1 positioned within the engine static structure 33.
[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 3 1 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 3 1 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 highbypass
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,
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 nonlimiting
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 Dl, whereas the low pressure turbine 39 rotates
in a second direction D2 that is opposite of the first direction Dl.
[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 Dl thereby turning the plurality of star gears
56 in a second direction D2 that is opposite of the first direction Dl. 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 Dl 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
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.
CLAIMS
What is claimed is:
1. A gas turbine engine comprising:
a fan section including a fan rotatable about an axis; and
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.
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 3, including a bypass ratio that is greater than about
6.0.
5. The gas turbine engine of claim 3, wherein the bypass ratio is between about 11.0 and
about 22.0.
6. 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.
7. The gas turbine engine of claim 6, wherein each of the plurality of star gears include
at least one bearing.
8. The gas turbine engine of claim 6, wherein the carrier is fixed from rotation.
9. The gas turbine engine of claim 6, wherein a low pressure turbine is mechanically
attached to the sun gear.
10. The gas turbine engine of claim 6, wherein a fan section is mechanically attached to
the ring gear.
11. 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.
12. 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.
13. The gas turbine engine of claim 12, wherein the fan blade tip speed of the fan is
greater than 1000 fps.
14. A method of improving performance of a gas turbine engine comprising:
determining fan tip speed boundary conditions for at least one fan blade of a fan
section;
determining rotor boundary conditions for a rotor of a low pressure turbine; and
utilizing stress level 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.
15. The method of claim 14, wherein 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.
16. The method of claim 15, wherein a fan pressure ratio is below 1.7.
17. The method of claim 15, wherein a fan pressure ratio is below 1.48.
18. The method of claim 16, wherein a bypass ratio is between about 11 and about 22.
19. The method of claim 14, wherein a fan blade tip speed of the at least one fan blade is
less than 1400 fps.
20. The method of claim 14, wherein 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.
21. The method of claim 14, wherein 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.