Note: Descriptions are shown in the official language in which they were submitted.
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GAS TURBINE ENGINE WITH FORWARD MOMENT ARM
BACKGROUND
[0002] A gas turbine engine typically includes a fan section, a
compressor section,
a combustor section and a turbine section. Air entering the compressor section
is
compressed and delivered into the combustion section where it is mixed with
fuel and
ignited to generate a high-speed exhaust gas flow. The high-speed exhaust gas
flow expands
through the turbine section to drive the compressor and the fan section. The
compressor
section typically includes low and high pressure compressors, and the turbine
section
includes low and high pressure turbines.
[0003] A speed reduction device such as an epicyclical gear assembly
may be
utilized to drive the fan section such that the fan section may rotate at a
speed different than
the turbine section so as to increase the overall propulsive efficiency of the
engine. In such
engine architectures, a shaft driven by one of the turbine sections provides
an input to the
epicyclical gear assembly 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.
[0004] Structures of a gas turbine engine contribute to an overall
weight of the
engine and balance point is defined at center of gravity. The location of the
center of gravity
of a gas turbine engine influences how an engine is mounted and how
surrounding nacelle
structures are configured. A center of gravity moved forward along an engine
axis increases
an internal moment arm and thereby increases load on engine mounting
structures. The
location of the engine center of gravity is influenced by selections of
materials and
component configurations.
[0005] Although geared architectures have improved propulsive
efficiency,
turbine engine manufacturers continue to seek further improvements to engine
performance
including improvements to thermal, transfer and propulsive efficiencies.
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SUMMARY
[0006] A gas
turbine engine according to an exemplary embodiment of this
disclosure, among other possible things includes a plurality of fan blades
rotatable about an
axis, each of the plurality of fan blades including a leading edge, a turbine
section including
an aft most turbine blade having a trailing edge, and a geared architecture
driven by the
turbine section for rotating the plurality of fan blades about the axis,
wherein a center of
gravity of the gas turbine engine is located a first axial distance from the
trailing edge of the
aft most turbine blade that is between about 35% and about 75% of a total
length between the
leading edge of the plurality of fan blades and the trailing edge of the aft
most turbine blade.
[0007] In a
further embodiment of the foregoing gas turbine engine, the center of
gravity is disposed substantially along the axis.
[0008] In a
further embodiment of any of the foregoing turbine engine
embodiments, the center of gravity is determined including weights of
structures comprising
the gas turbine engine not including engine mounting structures, engine
cowling structures
and nacelle structures.
[0009] In a
further embodiment of any of the foregoing turbine engine
embodiments, the center of gravity is determined including weights of fluids
contained within
operating systems of the gas turbine engine.
[0010] In a
further embodiment of any of the foregoing turbine engine
embodiments, the first axial distance is between about 40% and about 70% of
the total length
between the leading edge of the plurality of fan blades and the trailing edge
of the aft most
turbine blade.
[0011] In a
further embodiment of any of the foregoing turbine engine
embodiments, the plurality of fan blades are supported on a rotor with the
rotor and fan
blades having a density of between about 0.0094 lbs/ lbs/in3 and about 0.01540
lbs/ lbs/in3.
[0012] In a
further embodiment of any of the foregoing turbine engine
embodiments, the geared architecture comprises a gearbox having a density of
between about
0.22 lbs/in3 and about 0.30 lbs/in3.
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[0013] In a further embodiment of any of the foregoing turbine engine
embodiments, the center of gravity is located at an intersection of a vertical
line extending
through a hoist point of the gas turbine engine and the axis with the axis
normal to the
vertical line.
[0014] Another gas turbine engine according to an exemplary embodiment
of this
disclosure, among other possible things includes a plurality of fan blades
rotatable about an
axis, each of the plurality of fan blades including a leading edge, a turbine
section including
an aft most rotating turbine blade having a trailing edge, and a geared
architecture driven by
the turbine section for rotating the plurality of fan blades about the axis,
wherein an internal
moment arm of the turbofan engine comprises a ratio of first distance from a
center of gravity
of the turbofan engine to the trailing edge of the aft most rotating turbine
blade to a total
length between the leading edge of the plurality of fan blades and the
trailing edge of the aft
most turbine blade that is between about 35% and about 75%.
[0015] In a further embodiment of the foregoing gas turbine engine,
the ratio is
between about 40% and 70%.
[0016] In a further embodiment of any of the foregoing gas turbine
engines, the
geared architecture comprises gearbox having a density of between about 0.22
lbs/in3 and
about 0.30 lbs/in3.
[0017] In a further embodiment of any of the foregoing gas turbine
engines,
including a rotor supporting the plurality of fan blades with the rotor and
fan blades having a
density of between about 0.0094 lbs/in3and about 0.01540 lbs/in3.
[0018] In a further embodiment of any of the foregoing gas turbine
engines, the
center of gravity is determined including weights of fluids contained within
operating
systems of the turbofan engine.
[0019] In a further embodiment of any of the foregoing gas turbine
engines, the
center of gravity is located at an intersection of a vertical line extending
through a hoist point
of the turbofan engine and the axis with the axis normal to the vertical line.
[0020] A method of assembling a gas turbine engine according to an
exemplary
embodiment of this disclosure among other possible things includes supporting
a plurality of
fan blades about an axis of rotation with each of the plurality of fan blades
including a
leading edge, supporting a turbine section including an aft most turbine blade
having a
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trailing edge about the axis of rotation, supporting a geared architecture
driven by the turbine
section for rotating the plurality of fan blades about the axis, and selecting
components of the
gas turbine engine structure to orientate a center of gravity of the gas
turbine engine located a
first axial distance from the trailing edge of the aft most turbine blade that
is between about
35% and about 75% of a total length between the leading edge of the plurality
of fan blades
and the trailing edge of the aft most turbine blade.
[0021] In a further embodiment of the foregoing method, including
supporting the
plurality of blades on a rotor with the rotor and fan blades having a density
of between about
0.0094 lbs/in3and about 0.01540 lbs/in3.
[0022] In a further embodiment of any of the foregoing methods,
including
assembling a geared architecture as a gearbox having a density of between
about 0.22 lbs/
lbs/in3 and about 0.30 lbs/in3.
[0023] In a further embodiment of any of the foregoing methods
including
selecting components of the gas turbine engine to orientate the center of
gravity within a
range of between about 40% and 70% of total length between the leading edge of
the
plurality of fan blades and the trailing edge of the aft most turbine blade.
[0024] Although the different examples have the specific components
shown in
the illustrations, embodiments of this disclosure are not limited to those
particular
combinations. It is possible to use some of the components or features from
one of the
examples in combination with features or components from another one of the
examples.
[0025] These and other features disclosed herein can be best
understood from the
following specification and drawings, the following of which is a brief
description.
BRIEF DESCRIPTION OF THE DRAWINGS
[0026] Figure 1 is a schematic view of a center of gravity of an
example gas
turbine engine.
[0027] Figure 2 is a schematic view of a center of gravity of an
example gas
turbine engine.
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DETAILED DESCRIPTION
[0028] Figure 1
schematically illustrates an example gas turbine engine 20 that
includes a fan section 22, a compressor section 24, a combustor section 26 and
a turbine
section 28. Alternative engines might include an augmenter section (not shown)
among other
systems or features. The fan section 22 drives air along a bypass flow path B
while the
compressor section 24 draws air in along a core flow path C where air is
compressed and
communicated to a combustor section 26. In the combustor section 26, air is
mixed with fuel
and ignited to generate a high pressure exhaust gas stream that expands
through the turbine
section 28 where energy is extracted and utilized to drive the fan section 22
and the
compressor section 24.
[0029] Although
the disclosed non-limiting embodiment depicts a turbofan gas
turbine engine, it should be understood that the concepts described herein are
not limited to
use with turbofans as the teachings may be applied to other types of turbine
engines; for
example a turbine engine including a three-spool architecture in which three
spools
concentrically rotate about a common axis and where a low spool enables a low
pressure
turbine to drive a fan via a gearbox, an intermediate spool that enables an
intermediate
pressure turbine to drive a first compressor of the compressor section, and a
high spool that
enables a high pressure turbine to drive a high pressure compressor of the
compressor
section.
[0030] The
example engine 20 generally includes a low speed spool 30 and a high
speed spool 32 mounted for rotation about an engine central longitudinal axis
A relative to an
engine static structure 36 via several bearing systems 38. It should be
understood that various
bearing systems 38 at various locations may alternatively or additionally be
provided.
[0031] The low
speed spool 30 generally includes an inner shaft 40 that connects
a fan 42 and a low pressure (or first) compressor section 44 to a low pressure
(or first) turbine
section 46. The inner shaft 40 drives the fan 42 through a speed change
device, such as a
geared architecture 48, to drive the fan 42 at a lower speed than the low
speed spool 30. The
high-speed spool 32 includes an outer shaft 50 that interconnects a high
pressure (or second)
compressor section 52 and a high pressure (or second) turbine section 54. The
inner shaft 40
and the outer shaft 50 are concentric and rotate via the bearing systems 38
about the engine
central longitudinal axis A.
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[0032] A
combustor 56 is arranged between the high pressure compressor 52 and
the high pressure turbine 54. In one example, the high pressure turbine 54
includes at least
two stages to provide a double stage high pressure turbine 54. In another
example, the high
pressure turbine 54 includes only a single stage. As used herein, a "high
pressure"
compressor or turbine experiences a higher pressure than a corresponding "low
pressure"
compressor or turbine.
[0033] The
example low pressure turbine 46 has a pressure ratio that is greater
than about 5. The pressure ratio of the example low pressure turbine 46 is
measured prior to
an inlet of the low pressure turbine 46 as related to the pressure measured at
the outlet of the
low pressure turbine 46 prior to an exhaust nozzle.
[0034] A mid-
turbine frame 58 of the engine static structure 36 is arranged
generally between the high pressure turbine 54 and the low pressure turbine
46. The mid-
turbine frame 58 further supports bearing systems 38 in the turbine section 28
as well as
setting airflow entering the low pressure turbine 46.
[0035] Airflow
through the core flow path C is compressed by the low pressure
compressor 44 then by the high pressure compressor 52 mixed with fuel and
ignited in the
combustor 56 to produce high speed exhaust gases that are then expanded
through the high
pressure turbine 54 and low pressure turbine 46. The mid-turbine frame 58
includes vanes
60, which are in the core airflow path and function as an inlet guide vane for
the low pressure
turbine 46. Utilizing the vane 60 of the mid-turbine frame 58 as the inlet
guide vane for low
pressure turbine 46 decreases the length of the low pressure turbine 46
without increasing the
axial length of the mid-turbine frame 58. Reducing or eliminating the number
of vanes in the
low pressure turbine 46 shortens the axial length of the turbine section 28.
Thus, the
compactness of the gas turbine engine 20 is increased and a higher power
density may be
achieved.
[0036] The
disclosed gas turbine engine 20 in one example is a high-bypass
geared aircraft engine. In a further example, the gas turbine engine 20
includes a bypass ratio
greater than about six (6), with an example embodiment being greater than
about ten (10).
The example geared architecture 48 is an epicyclical gear train, such as a
planetary gear
system, star gear system or other known gear system, with a gear reduction
ratio of greater
than about 2.3.
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[0037] In one
disclosed embodiment, the gas turbine engine 20 includes a bypass
ratio greater than about ten (10:1) and the fan diameter is significantly
larger than an outer
diameter of the low pressure compressor 44. It should be understood, however,
that the
above parameters are only exemplary of one embodiment of a gas turbine engine
including a
geared architecture and that the present disclosure is applicable to other gas
turbine engines.
[0038] A
significant amount of thrust is provided by the bypass flow B due to the
high bypass ratio. The fan section 22 of the engine 20 is designed for a
particular flight
condition -- typically cruise at about 0.8 Mach and about 35,000 feet. The
flight condition of
0.8 Mach and 35,000 ft., with the engine at its best fuel consumption - also
known as "bucket
cruise Thrust Specific Fuel Consumption ('TSFC')" - is the industry standard
parameter of
pound-mass (lbm) of fuel per hour being burned divided by pound-force (lbf) of
thrust the
engine produces at that minimum point.
[0039] "Low fan
pressure ratio" is the pressure ratio across the fan blade alone,
without a Fan Exit Guide Vane ("FEGV") system. The low fan pressure ratio as
disclosed
herein according to one non-limiting embodiment is less than about 1.50. In
another non-
limiting embodiment the low fan pressure ratio is less than about 1.45.
[0040] "Low
corrected fan tip speed" is the actual fan tip speed in ft/sec divided
by an industry standard temperature correction of [(Tram `R)/ (518.7 R)1 05.
The "Low
corrected fan tip speed", as disclosed herein according to one non-limiting
embodiment, is
less than about 1150 ft/second.
[0041] The
example gas turbine engine includes the fan 42 that comprises in one
non-limiting embodiment less than about 26 fan blades. In another non-limiting
embodiment,
the fan section 22 includes less than about 20 fan blades. Moreover, in one
disclosed
embodiment the low pressure turbine 46 includes no more than about 6 turbine
rotors
schematically indicated at 34. In another non-limiting example embodiment the
low pressure
turbine 46 includes about 3 turbine rotors. A ratio between the number of fan
blades 42 and
the number of low pressure turbine rotors is between about 3.3 and about 8.6.
The example
low pressure turbine 46 provides the driving power to rotate the fan section
22 and therefore
the relationship between the number of turbine rotors 34 in the low pressure
turbine 46 and
the number of blades 42 in the fan section 22 disclose an example gas turbine
engine 20 with
increased power transfer efficiency.
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[0042]
Referring to Figure 2, with continued reference to Figure 1, a center of
gravity indicated at 62 is transferred radially to an engine centerline
disposed along the axis
A. The center of gravity 62 is positioned a first distance XCG between forward
and aft ends of
the engine 20. In the disclosed example, the forward end is a leading edge 66
of the fan
blades 42 and the aft end is a trailing edge 68 of an aft most rotating
turbine blade 70. A total
length XL is defined between the leading edge 66 and the trailing edge 68.
[0043] The
position of the center of gravity 62 along the axis A is disposed at the
first distance XCG and influences the configuration for supporting engines on
an airframe and
is therefore of concern to engine manufactures and aircraft designers.
[0044] In the
disclosed example, a mounting structure schematically indicated at
64 supports the example gas turbine engine 20 on an airframe (not shown). It
should be
understood, that the location of the mounting structure 64 may vary for each
engine
application and such variations are within the contemplation of this
disclosure.
[0045] The
example gas turbine engine 20 includes the geared architecture 48 for
driving the fan section 22 at a speed different than that of a fan drive
turbine. In this
example, the fan drive turbine is the low pressure turbine 46.
[0046] The
selection of materials and component configurations define the overall
weight of the engine 20 along with the distribution of that weight to
determine the location of
the center of gravity 62. Structures such as the fan section 22 and the geared
architecture 48
along with the compressor section 24 and turbine section 28 combine to define
not only the
overall weight of the engine 20, but also the distribution of that weight that
determines the
location of the center of gravity 62.
[0047] The
geared architecture 48 is a structure located forward in the engine 20
along with the fan section 22 and therefore material selection and structure
configurations
influence the location of the center of gravity 62. Moreover, many structures
within the gas
turbine engine structures factor into and determine in part the positioning
and location of the
center of gravity 62.
[0048] In this
disclosed example, components of the geared architecture 48 such
as for example journal bearings, lubrication jets, and a gutter around
epicyclic components
are selected to provide a weight reduction and an axial size reduction that
define an overall
weight of the geared architecture 48 and thereby factor into the definition of
the center of
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gravity 62. In one disclosed example, the weight of the geared architecture is
stated as a
density to relate the overall size or volume of the geared architecture to a
weight. In this
example, the example geared architecture is a gearbox having a density between
about 0.22
lbs/in3 and about 0.30 lbs/in3. The part density of the gearbox is one
consideration that
influences the position of the center of gravity 62.
[0049]
Additionally, the fan section 22 is the one of the forward most components
of the example engine and therefore also has a significant influence on the
final location of
the center of gravity 62 in a completed engine 22. Accordingly, selection of
parts making up
the fan section 22 is considered not only in view of propulsive efficiencies,
but also in regard
to the overall structure of the gas turbine engine 20.
[0050] In the
disclosed example, the fan section 22 includes a rotor 72 that
supports the plurality of fan blades 42 for rotation about the axis A. The
rotor 72
arrangement that folds radially back around the bearings 38 provides a weight
benefit and
thereby provides another means for modifying and positioning the engine center
of gravity
62. Moreover, the rotor bearing 38 can include tapered roller bearings that
further provide a
beneficial impact on the structure of the fan section without adding
additional weight to
further influence the location of the center of gravity 62.
[0051] A
density of the fan rotor can be further selected to utilize light weight
structures that further correspond and effect the location of the center of
gravity 62.
In this example, the fan section 22, including the rotor 72 and the plurality
of fan blades
combine to provide a density within a range of between about 0.0094 lbs/in3
and about
0.01540 lbs/in3.
[0052]
Additionally, a fan containment case 16 is required to contain the blades
42 and is fabricated from a composite material to reduce weight and is a
selection that
determines the location of the center of gravity 62.
[0053]
Utilizing these and other configuration parameters and material selection
options, the example geared turbofan engine has an internal moment arm that is
a measure of
the location of the engine center of gravity 62. The example internal moment
arm is within a
range of between about 40% and 70% of the length of the engine 20 according to
the
following relationship.
[0054] The internal moment arm MA is defined as:
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EaxtaZ distance Between tho Center of Gravity (transferred nachaily t tke
center14'neof t gtneHthe location of the trailing edgF Cr the
tos.t
stage of the taitine zection)]
[0055] =
cLstertce Lle twe tite front e:lge or the prarit fan cnd the trailing
edge of iast stop? or the turbine section]
[0056] In this example MA= XcGAL.
[0057] Accordingly, for a geared turbofan gas turbine engine the
position of the
center of gravity 62 is related as a moment arm according to the above
equation and is within
a range of between about 35% and about 75% of the total length XL between the
leading edge
66 of the fan blades 42 and the trailing edge 68 of the aft most rotating
turbine blade 70. In
another disclosed embodiment the example moment arm MA for the example engine
20 is
within a range of between about 40% and about 70% of the total length XL
between the
leading edge 66 of the fan blades 42 and the trailing edge 68 of the aft most
rotating turbine
blade 70.
[0058] The location of the center of gravity 62 for the disclosed geared
gas
turbine engine is only about 30% longer than direct drive turbine engines in
the prior art.
Table 1 below includes further disclosed example embodiments of a geared
turbofan engine
moment arms MA.
MA as a
LPIiast percent of
CG :;?=:tion Fan LE isiade TE Ongine
(inches) station station knIgth
Engine
1 205.29 158 251 49.2%
2 203.31 152 253.3 49.3%
3 203.9 149.4 258 49.8%
Table 1
[0059] The disclosed moment arms MA represents the location of the
center of
gravity 62 as a percentage of the engine length XL (Figure 1) measured between
the leading
edge 66 of the fan blade 72 and the trailing edge 68 of the last rotating
turbine blade 70 in the
low pressure turbine 46.
[0060] Referring to Figure 2, although complex calculations can be
utilized to
determine the physical location of the center of gravity 62 of any engine,
another means
exists for determined this location for any engine. A hoist point 74 located
on the engine 20
will be axially located at the center of gravity 62 when the engine center
line or axis A is
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normal to a vertical line 76 extending through the hoist point 74 and
intersecting the axis A.
Accordingly, it this example the center of gravity 62 may in one example be
determined as
that point where the vertical line 76 intersects the engine center line or
axis A at a right angle
when supported at a single hoist point 74.
[0061] The
example engine center of gravity 62 is considered along the engine
centerline or axis A although the actual center of gravity 62 may be slightly
skewed from the
axis A due to locally mounted accessory components.
[0062] The
disclosed center of gravity 62 includes fluids contained within
operating systems of the turbofan engine 20. However, the example center of
gravity does
not include some structures such as for example typical tubes, brackets and
harness such as
those coming from the airframe which would have almost no effect on the
location of the
center of gravity 62. Moreover, the example center of gravity 62 is determined
for a bare
engine only that does not include engine mounts, a fan cowl, a thrust
reverser, an inlet, nozzle
or plug. In other words, the example center of gravity 62 is not determined
including weights
of structures comprising the gas turbine engine not including engine mounting
structures,
engine cowling structures and nacelle structures.
[0063]
Accordingly, through selection of materials and design of structures such
as the fan section and geared architecture, the center of gravity 62 can be
located in a
structurally desirable location to increase propulsive efficiencies and reduce
mounting
structure requirements.
[0064] Although
an example embodiment has been disclosed, a worker of
ordinary skill in this art would recognize that certain modifications would
come within the
scope of this disclosure. For that reason, the following claims should be
studied to determine
the scope and content of this disclosure.
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