Note : Les descriptions sont présentées dans la langue officielle dans laquelle elles ont été soumises.
CA 02853694 2015-10-07
GAS TURBINE ENGINE WITH GEARED ARCHITECTURE
BACKGROUND
[0002] The present invention relates to a gas turbine engine, and more
particularly to
a turbofan engine having a variable geometry fan exit guide vane (FEGV) system
to change a fan
bypass flow path area thereof.
[0003] Conventional gas turbine engines generally include a fan
section and a core
section with the fan section having a larger diameter than that of the core
section. The fan section
and the core section are disposed about a longitudinal axis and are enclosed
within an engine
nacelle assembly. Combustion gases are discharged from the core section
through a core exhaust
nozzle while an annular fan bypass flow, disposed radially outward of the
primary core exhaust
path, is discharged along a fan bypass flow path and through an annular fan
exhaust nozzle. A
majority of thrust is produced by the bypass flow while the remainder is
provided from the
combustion gases.
[0004] The fan bypass flow path is a compromise suitable for take-off
and landing
conditions as well as for cruise conditions. A minimum area along the fan
bypass flow path
determines the maximum mass flow of air. During engine-out conditions,
insufficient flow area
along the bypass flow path may result in significant flow spillage and
associated drag. The fan
nacelle diameter is typically sized to minimize drag during these engine-out
conditions which
results in a fan nacelle diameter that is larger than necessary at normal
cruise conditions with less
than optimal drag during portions of an aircraft mission.
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SUMMARY
[0005] A gas turbine engine according to an exemplary aspect of the
present
disclosure includes a core nacelle defined about an engine centerline axis, a
fan nacelle mounted
at least partially around the core nacelle to define a fan bypass flow path
for a fan bypass airflow,
a fan variable area nozzle axially movable relative the fan nacelle to vary a
fan nozzle exit area
and adjust a pressure ratio of the fan bypass airflow during engine operation,
and a gear system
driven by a core engine within the core nacelle to drive a fan within the fan
nacelle, the gear
system defines a gear reduction ratio of greater than or equal to about 2.3.
[0006] In a further non-limiting embodiment of any of the foregoing
gas turbine
engine embodiments, the engine may further include a multiple of fan exit
guide vanes in
communication with the fan bypass flow path, the multiple of fan exit guide
vane rotatable about
an axis of rotation to vary the fan bypass flow path. Additionally or
alternatively, the multiple of
fan exit guide vanes may be simultaneously rotatable. Additionally or
alternatively, the multiple
of fan exit guide vanes may be mounted within an intermediate engine case
structure.
[0007] In a further non-limiting embodiment of any of the foregoing
gas turbine
engine embodiments, each of the multiple of fan exit guide vanes may include a
pivotable
portion rotatable about the axis of rotation relative a fixed portion.
Additionally or alternatively,
the pivotable portion may include a leading edge flap.
[0008] In a further non-limiting embodiment of any of the foregoing
gas turbine
engine embodiments, the controller may be operable to control the fan variable
area nozzle to
vary a fan nozzle exit area and adjust the pressure ratio of the fan bypass
airflow. Additionally
or alternatively, the controller may be operable to reduce the fan nozzle exit
area at a cruise flight
condition. Additionally or alternatively, the controller may be operable to
control the fan nozzle
exit area to reduce a fan instability.
[0009] In a further non-limiting embodiment of any of the foregoing
gas turbine
engine embodiments, the fan variable area nozzle may define a trailing edge of
the fan nacelle.
[0010] In a further non-limiting embodiment of any of the foregoing
gas turbine
engine embodiments, the gear system may define a gear reduction ratio of
greater than or equal
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to about 2.5. Additionally or alternatively, the gear system may define a gear
reduction ratio of
greater than or equal to 2.5.
[0011] In a further non-limiting embodiment of any of the foregoing
gas turbine
engine embodiments, the core engine may include a low pressure turbine which
defines a
pressure ratio that is greater than about five (5).
[0012] In a further non-limiting embodiment of any of the foregoing
gas turbine
engine embodiments, the core engine may include a low pressure turbine which
defines a
pressure ratio that is greater than five (5).
[0013] In a further non-limiting embodiment of any of the foregoing
gas turbine
engine embodiments, the bypass flow may define a bypass ratio greater than
about six (6).
Additionally or alternatively, the bypass flow may define a bypass ratio
greater than about ten
(10). Additionally or alternatively, the bypass flow may define a bypass ratio
greater than ten
(10).
[0014] In a further non-limiting embodiment of any of the foregoing
gas turbine
engine embodiments, the engine may further comprise a multiple of fan exit
guide vanes in
communication with the fan bypass flow path, the multiple of fan exit guide
vanes rotatable
about an axis of rotation to vary the fan bypass flow path.
BRIEF DESCRIPTION OF THE DRAWINGS
[0015] The various features and advantages of this invention will
become apparent to
those skilled in the art from the following detailed description of the
currently preferred
embodiment. The drawings that accompany the detailed description can be
briefly described as
follows:
[0016] Figure lA is a general schematic partial fragmentary view of an
exemplary
gas turbine engine embodiment for use with the present invention;
[0017] Figure 1B is a perspective side partial fragmentary view of a
FEGV system
which provides a fan variable area nozzle;
[0018] Figure 2A is a sectional view of a single FEGV airfoil;
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[0019] Figure 2B is a sectional view of the FEGV illustrated in Figure
2A shown in a
first position;
[0020] Figure 2C is a sectional view of the FEGV illustrated in Figure
2A shown in a
rotated position;
[0021] Figure 3A is a sectional view of another embodiment of a single
FEGV
airfoil;
[0022] Figures 3B is a sectional view of the FEGV illustrated in
Figure 3A shown in
a first position;
[0023] Figure 3C is a sectional view of the FEGV illustrated in Figure
3A shown in a
rotated position;
[0024] Figure 4A is a sectional view of another embodiment of a single
FEGV slatted
airfoil with a;
[0025] Figures 4B is a sectional view of the FEGV illustrated in
Figure 4A shown in
a first position; and
[0026] Figure 4C is a sectional view of the FEGV illustrated in Figure
4A shown in a
rotated position.
DETAILED DESCRIPTION OF THE DISCLOSED EMBODIMENT
[0027] Figure 1 illustrates a general partial fragmentary schematic
view of a gas
turbofan engine 10 suspended from an engine pylon P within an engine nacelle
assembly N as is
typical of an aircraft designed for subsonic operation.
[0028] The turbofan engine 10 includes a core section within a core
nacelle 12 that
houses a low spool 14 and high spool 24. The low spool 14 includes a low
pressure compressor
16 and low pressure turbine 18. The low spool 14 drives a fan section 20
directly or through a
gear train 22. The high spool 24 includes a high pressure compressor 26 and
high pressure
turbine 28. A combustor 30 is arranged between the high pressure compressor 26
and high
pressure turbine 28. The low and high spools 14, 24 rotate about an engine
axis of rotation A.
[0029] The engine 10 is a high-bypass geared architecture aircraft
engine. In one
disclosed, non-limiting embodiment, the engine 10 bypass ratio is greater than
about six (6), with
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an example embodiment being greater than about ten (10), the gear train 22 is
an epicyclic gear
train such as a planetary gear system or other gear system with a gear
reduction ratio of greater
than about 2.3 and the low pressure turbine 18 has a pressure ratio that is
greater than about five
(5). The engine 10 in the disclosed embodiment is a high-bypass geared
turbofan aircraft engine
in which the engine 10 bypass ratio is greater than ten (10), the turbofan
diameter is significantly
larger than that of the low pressure compressor 16, and the low pressure
turbine 18 has a pressure
ratio greater than five (5). Low pressure turbine 18 pressure ratio is
pressure measured prior to
inlet of low pressure turbine 18 as related to the pressure at the outlet of
the low pressure turbine
18 prior to exhaust nozzle. The gear train 22 may be an epicycle gear train
such as a planetary
gear system or other gear system with a gear reduction ratio of greater than
about 2.5. It should
be understood, however, that the above parameters are exemplary of only one
geared turbofan
engine and that the present invention is likewise applicable to other gas
turbine engines including
direct drive turbofans.
[0030] Airflow enters a fan nacelle 34, which may at least partially
surround the core
nacelle 12. The fan section 20 communicates airflow into the core nacelle 12
for compression by
the low pressure compressor 16 and the high pressure compressor 26. Core
airflow compressed
by the low pressure compressor 16 and the high pressure compressor 26 is mixed
with the fuel in
the combustor 30 then expanded over the high pressure turbine 28 and low
pressure turbine 18.
The turbines 28, 18 are coupled for rotation with respective spools 24, 14 to
rotationally drive the
compressors 26, 16 and, through the gear train 22, the fan section 20 in
response to the
expansion. A core engine exhaust E exits the core nacelle 12 through a core
nozzle 43 defined
between the core nacelle 12 and a tail cone 32.
[0031] A bypass flow path 40 is defined between the core nacelle 12
and the fan
nacelle 34. The engine 10 generates a high bypass flow arrangement with a
bypass ratio in
which approximately 80 percent of the airflow entering the fan nacelle 34
becomes bypass flow
B. The bypass flow B communicates through the generally annular bypass flow
path 40 and may
be discharged from the engine 10 through a fan variable area nozzle (FVAN) 42
which defines a
variable fan nozzle exit area 44 between the fan nacelle 34 and the core
nacelle 12 at an aft
segment 34S of the fan nacelle 34 downstream of the fan section 20.
CA 02853694 2015-10-07
[0032] Referring to Figure 1B, the core nacelle 12 is generally
supported upon a core
engine case structure 46. A fan case structure 48 is defined about the core
engine case structure
46 to support the fan nacelle 34. The core engine case structure 46 is secured
to the fan case 48
through a multiple of circumferentially spaced radially extending fan exit
guide vanes (FEGV)
50. The fan case structure 48, the core engine case structure 46, and the
multiple of
circumferentially spaced radially extending fan exit guide vanes 50 which
extend therebetween is
typically a complete unit often referred to as an intermediate case. It should
be understood that
the fan exit guide vanes 50 may be of various forms. The intermediate case
structure in the
disclosed embodiment includes a variable geometry fan exit guide vane (FEGV)
system.
[0033] Thrust is a function of density, velocity, and area. One or
more of these
parameters can be manipulated to vary the amount and direction of thrust
provided by the bypass
flow B. A significant amount of thrust is provided by the bypass flow B due to
the high bypass
ratio. The fan section 20 of the engine 10 is nominally 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 CTSFCT - is the industry standard parameter of lbm
of fuel being
burned divided by lbf of thrust the engine produces at that minimum point.
"Low fan pressure
ratio" is the pressure ratio across the fan blade alone, without the fan exit
guide vane (FEGV)
system. The low fan pressure ratio as disclosed herein according to one non-
limiting
embodiment is less than about 1.45. "Low corrected fan tip speed" is the
actual fan tip speed in
ft/sec divided by an industry standard temperature correction of [(Tambient
deg R) / 518.7)^0.5].
The "Low corrected fan tip speed" as disclosed herein according to one non-
limiting
embodiment is less than about 1150 ft / second.
[0034] As the fan section 20 is efficiently designed at a particular
fixed stagger angle
for an efficient cruise condition, the FEGV system and/or the FVAN 42 is
operated to adjust fan
bypass air flow such that the angle of attack or incidence of the fan blades
is maintained close to
the design incidence for efficient engine operation at other flight
conditions, such as landing and
takeoff. The FEGV system and/or the FVAN 42 may be adjusted to selectively
adjust the
pressure ratio of the bypass flow B in response to a controller C. For
example, increased mass
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flow during windmill or engine-out, and spoiling thrust at landing.
Furthermore, the FEGV
system will facilitate and in some instances replace the FVAN 42, such as, for
example, variable
flow area is utilized to manage and optimize the fan operating lines which
provides operability
margin and allows the fan to be operated near peak efficiency which enables a
low fan pressure-
ratio and low fan tip speed design; and the variable area reduces noise by
improving fan blade
aerodynamics by varying blade incidence. The FEGV system thereby provides
optimized engine
operation over a range of flight conditions with respect to performance and
other operational
parameters such as noise levels.
[0035] Referring to Figure 2A, each fan exit guide vane 50 includes a
respective
airfoil portion 52 defined by an outer airfoil wall surface 54 between the
leading edge 56 and a
trailing edge 58. The outer airfoil wall 54 typically has a generally concave
shaped portion
forming a pressure side and a generally convex shaped portion forming a
suction side. It should
be understood that respective airfoil portion 52 defined by the outer airfoil
wall surface 54 may
be generally equivalent or separately tailored to optimize flow
characteristics.
[0036] Each fan exit guide vane 50 is mounted about a vane
longitudinal axis of
rotation 60. The vane axis of rotation 60 is typically transverse to the
engine axis A, or at an
angle to engine axis A. It should be understood that various support struts 61
or other such
members may be located through the airfoil portion 52 to provide fixed support
structure
between the core engine case structure 46 and the fan case structure 48. The
axis of rotation 60
may be located about the geometric center of gravity (CG) of the airfoil cross
section. An
actuator system 62 (illustrated schematically; Figure 1A), for example only, a
unison ring
operates to rotate each fan exit guide vane 50 to selectively vary the fan
nozzle throat area
(Figure 2B). The unison ring may be located, for example, in the intermediate
case structure
such as within either or both of the core engine case structure 46 or the fan
case 48 (Figure 1A).
[0037] In operation, the FEGV system communicates with the controller
C to rotate
the fan exit guide vanes 50 and effectively vary the fan nozzle exit area 44.
Other control
systems including an engine controller or an aircraft flight control system
may also be usable
with the present invention. Rotation of the fan exit guide vanes 50 between a
nominal position
and a rotated position selectively changes the fan bypass flow path 40. That
is, both the throat
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area (Figure 2B) and the projected area (Figure 2C) are varied through
adjustment of the fan exit
guide vanes 50. By adjusting the fan exit guide vanes 50 (Figure 2C), bypass
flow B is increased
for particular flight conditions such as during an engine-out condition. Since
less bypass flow
will spill around the outside of the fan nacelle 34, the maximum diameter of
the fan nacelle
required to avoid flow separation may be decreased. This will thereby decrease
fan nacelle drag
during normal cruise conditions and reduce weight of the nacelle assembly.
Conversely, by
closing the FEGV system to decrease flow area relative to a given bypass flow,
engine thrust is
significantly spoiled to thereby minimize or eliminate thrust reverser
requirements and further
decrease weight and packaging requirements. It should be understood that other
arrangements as
well as essentially infinite intermediate positions are likewise usable with
the present invention.
[0038] By adjusting the FEGV system in which all the fan exit guide
vanes 50 are
moved simultaneously, engine thrust and fuel economy are maximized during each
flight regime.
By separately adjusting only particular fan exit guide vanes 50 to provide an
asymmetrical fan
bypass flow path 40, engine bypass flow may be selectively vectored to
provide, for example
only, trim balance, thrust controlled maneuvering, enhanced ground operations
and short field
performance.
[0039] Referring to Figure 3A, another embodiment of the FEGV system
includes a
multiple of fan exit guide vane 50' which each includes a fixed airfoil
portion 66F and pivoting
airfoil portion 66P which pivots relative to the fixed airfoil portion 66F.
The pivoting airfoil
portion 66P may include a leading edge flap which is actuatable by an actuator
system 62' as
described above to vary both the throat area (Figure 3B) and the projected
area (Figure 3C).
[0040] Referring to Figure 4A, another embodiment of the FEGV system
includes a
multiple of slotted fan exit guide vane 50" which each includes a fixed
airfoil portion 68F and
pivoting and sliding airfoil portion 68P which pivots and slides relative to
the fixed airfoil
portion 68F to create a slot 70 vary both the throat area (Figure 4B) and the
projected area
(Figure 4C) as generally described above. This slatted vane method not only
increases the flow
area but also provides the additional benefit that when there is a negative
incidence on the fan
exit guide vane 50" allows air flow from the high-pressure, convex side of the
fan exit guide
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vane 50" to the lower-pressure, concave side of the fan exit guide vane 50"
which delays flow
separation.
[0041] The foregoing description is exemplary rather than defined by
the limitations
within. Many modifications and variations of the present invention are
possible in light of the
above teachings. The preferred embodiments of this invention have been
disclosed, however,
one of ordinary skill in the art would recognize that certain modifications
would come within the
scope of this invention. It is, therefore, to be understood that within the
scope of the appended
claims, the invention may be practiced otherwise than as specifically
described. For that reason
the following claims should be studied to determine the true scope and content
of this invention.
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