Note: Descriptions are shown in the official language in which they were submitted.
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INTERMEDIATE FAN STAGE
BACKGROUND OF THE INVENTION
This invention relates generally to gas turbine engines, and, more
specifically,
to a gas turbine engine with an intermediate pressure fan stage having a tip
fan located
on a blade driven by a pressurized airflow.
In a turbofan aircraft gas turbine engine, air is pressurized in a fan module,
an
optional booster module and a compression module during operation. A portion
of the
air passing through the fan module is passed into a by-pass stream and used
for
generating a portion of the thrust needed for propelling an aircraft in
flight. The air
channeled through the optional booster module and compression module is mixed
with fuel in a combustor and ignited, generating hot combustion gases which
flow
through turbine stages that extract energy therefrom for powering the fan,
booster and
compressor rotors. The fan, booster and compressor modules have a series of
rotor
stages and stator stages. The fan and booster rotors are typically driven by a
low-
pressure turbine (LPT) and the compressor rotor is driven by a high-pressure
turbine
(HPT). The fan and booster rotors are aerodynamically coupled to the
compressor
rotor although the fan rotor and compressor rotor normally operate at
different
mechanical speeds.
It is often desirable to use an engine core comprising the compressor,
combustor, high-pressure turbine (HPT) and low-pressure turbine (LPT) from a
high
bypass commercial engine or a medium bypass engine with a moderate fan
pressure
ratio as a building block for lower bypass ratio engines with higher fan
pressure ratios.
The boost pressure and temperature into the high-pressure compressor (HPC) is
usually significantly higher in the low-bypass derivative engine than in the
original
high-bypass engine. This typically requires that the maximum operating airflow
in the
core be limited below its full design corrected airflow capacity due to
mechanical
limitations of the maximum physical core speed and/or the maximum compressor
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discharge temperature capability of the core. It is desirable to find a way to
operate the
original engine core airflow at its full potential while significantly
increasing the fan
pressure ratio to the bypass stream to maximize the thrust potential of the
derivative
engine.
Accordingly, it would be desirable to have a fan system that makes it possible
to operate the original engine core near its full airflow capability while
significantly
increasing the fan pressure ratio to the bypass stream to maximize the thrust
potential
of the derivative engine.
BRIEF DESCRIPTION OF THE INVENTION
The above-mentioned need or needs may be met by exemplary embodiments
which provide a fan system having a forward fan stage configured to pressurize
an
airflow and an aft fan stage having a tip-fan configured to pressurize a first
portion of
a pressurized air flow from the forward fan stage wherein the aft fan stage is
driven by
a second portion of the pressurized airflow.
In one aspect of the invention, the aft fan stage rotates independently from
the
forward fan stage.
In another aspect of the invention, the aft fan stage has an air turbine blade
comprising a turbine airfoil adapted to extract energy from a pressurized flow
of air
and a tip-fan blade adapted to pressurize a flow of air.
BRIEF DESCRIPTION OF THE DRAWINGS
The subject matter which is regarded as the invention is particularly pointed
out and distinctly claimed in the concluding part of the specification. The
invention,
however, may be best understood by reference to the following description
taken in
conjunction with the accompanying drawing figures in which:
Figure 1 is a schematic cross-sectional view of a portion of a gas turbine
engine with an exemplary embodiment of an intermediate fan stage according to
the
present invention.
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Figure 2 is a schematic cross-sectional view of an exemplary gas turbine
engine according to the present invention having an exemplary embodiment of an
intermediate fan stage.
DETAILED DESCRIPTION OF THE INVENTION
Referring to the drawings wherein identical reference numerals denote the
same elements throughout the various views, FIG. 1 shows an exemplary turbofan
gas
turbine engine 10 incorporating an exemplary embodiment of the present
invention.
The exemplary gas turbine engine 10 comprises an engine centerline axis 11, a
fan 12
which receives an inflow of ambient air 1, an optional booster or low-pressure
compressor (LPC) (not shown in FIG. 1), a high-pressure compressor (HPC) 18, a
combustor 20 which mixes fuel with the air pressurized by the HPC 18 for
generating
combustion gases which flow downstream through a high-pressure turbine (HPT)
22,
and a low-pressure turbine (LPT) 24 from which the combustion gases are
discharged
from the engine 10. The HPT 22 is coupled to the HPC 18 using a HPT shaft 23
to
substantially form a high-pressure rotor 29. A low-pressure shaft 25 joins the
LPT 24
to the fan 12 (and the optional booster if present) to substantially form a
low-pressure
rotor 28. The second or low-pressure shaft 25 is rotatably disposed co-axially
with
and radially inwardly of the high-pressure rotor 29. The low-pressure rotor 28
and the
high-pressure rotor 29 are aerodynamically coupled but rotate independently
since
they are not mechanically coupled.
The HPC 18 that pressurizes the air flowing through the core has a rotor 19
that rotates about the longitudinal centerline axis 11. The HPC system
includes a
plurality of inlet guide vanes (IGV) 30 and a plurality of stator vanes 31
arranged in a
circumferential direction around the longitudinal centerline axis 11. The HPC
18
further includes multiple rotor stages 19 which have corresponding rotor
blades 40
extending radially outwardly from a rotor hub 39 or corresponding rotors in
the form
of separate disks, or integral blisks, or annular drums in any conventional
manner. The
high-pressure rotor 29 is supported in the engine static frames using known
support
methods using suitable bearings.
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Cooperating with each rotor stage 19 is a corresponding stator stage
comprising a plurality of circumferentially spaced apart stator vanes 31. An
exemplary
arrangement of stator vanes and rotor blades for an axial flow high-pressure
compressor 18 is shown in FIG. 1. The rotor blades 40 and stator vanes 31
define
airfoils having corresponding aerodynamic profiles or contours for
pressurizing a core
airflow 8 successively in axial stages. The rotor blades 40 rotate within an
annular
casing 38 that surrounds the rotor blade tips. In operation, pressure of the
core air flow
8 is increased as the air decelerates and diffuses through the stator and
rotor airfoils.
FIG. 1 shows a fan system 50 comprising a forward fan stage 52 that
pressurizes an airflow 1. The pressurized airflow 2 exits axially aft from the
forward
fan stage 52. A static annular splitter 46 that is coaxial with the centerline
axis 11 is
located axially aft from the forward fan stage 52. The annular splitter 46
divides the
pressurized airflow 2 into a first portion 3 and a second portion 4, as shown
in FIG. 1.
The fan system 50 has an aft fan stage 60 that is located axially aft from the
annular splitter 46. The aft fan stage 60 comprises an aft fan rotor 61 and
has a
circumferential row of aft fan blades 62. The aft fan stage 60 rotates about
the
centerline axis 11 but it is not mechanically coupled with the high-pressure
compressor 18 or the forward fan stage 52. Although the aft fan stage 60 is
aerodynamically coupled during operation of the engine 10 to the forward fan
stage 52
and the forward stages of the high-pressure compressor 18, the aft fan stage
60 rotates
mechanically independently from the low-pressure rotor 28 and the high-
pressure
rotor 29. Thus, the aft fan stage 60 rotates independently from the forward
fan stage
52 that is located upstream from it.
As shown in FIG. 1, the aft fan stage 60 comprises a row of aft fan blades 62
arranged circumferentially around the longitudinal axis 11. Each aft fan blade
62 has a
radially inner portion 63 and an outer portion 64. The radially inner portion
63 of the
aft fan blade 62 is configured to be driven as an air-turbine blade 82 that
can extract
energy from a pressurized airflow 7 that enters the inner portion 63. Known
air-
turbine airfoil shapes can be used in the construction of the inner portion 63
aft fan
blade 62. As the airflows over the inner portion 63, it expands to form an
outflow 57
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of air that has a lower pressure and lower temperature and imparts energy to
the aft fan
blades 62 to drive the aft fan stage 60.
As shown in FIG. 1, each aft fan blade 62 has an outer portion 64 and an
arcuate shroud 65 between the inner portion 63 and the outer portion 64. The
outer
portion 64 of the aft fan blade 62 is configured to be a tip-fan blade 72 that
can
pressurize an inflow of air 6. The arcuate shroud 65 supports the tip-fan
blade 72. The
outer portion of the aft fan blade 62 has known airfoil shapes for fan blades
that can
pressurize an inflow of air 6. In the assembled state of the aft fan stage 60,
the arcuate
shroud 65 of each blade 62 abuts the arcuate shrouds of the circumferentially
adjacent
fan blades 62 to form an annular platform and a tip-fan 70 comprising the tip-
fan
blades 72. In one embodiment, each aft fan blade 62 has one tip-fan blade 72
supported by the arcuate shroud 65. In alternative embodiments, each aft fan
blade 62
may have a plurality of tip-fan blades 72 supported by the arcuate shroud 65.
As shown in FIG. 1, the aft fan stage 60 has a tip-fan 70 configured to
pressurize a first portion 3 of a pressurized air flow 2 from the forward fan
stage 52.
The tip-fan 70 is driven by the aft blade inner portion 63 that acts as an air
turbine
blade 82. The aft fan stage 60, with the tip-fan 70, is driven by a second
portion 4 of
the pressurized airflow 2. The inner portion 63 of the aft fan blade 62 is
configured to
work as an air-turbine blade that can extract energy from a pressurized air
stream
whereas the outer portion 64 of the aft fan blade 62 is configured to be a
compression-
type airfoil that can pressurize an airflow. The inner portion 63 is an air
turbine blade
82 having a turbine-type airfoil 84 adapted to extract energy from a
pressurized flow
of air. The outer portion 64 of the aft fan blade 62 is alternatively referred
to herein as
a tip-fan blade 72. The tip-fan blade 72 is capable of pressuring a flow of
air 6 to
create a pressurized tip flow 56 (see FIG. 1).
As shown in FIG. 1, the fan system 50 further comprises a circumferential row
of inlet guide vanes (IGV) 74 that are located axially forward from the tip-
fan 70 of
the aft fan stage 60. The IGVs 74 have known airfoil shapes that can re-orient
an
incoming airflow 3 to be an airflow 6 that suitably enters the tip-fan 70 for
further
pressurization. The IGVs 74 are suitably supported by an inner casing 68 (see
FIG. 1)
and/or by the splitter 46. For enhanced control of the operation of the aft
fan stage 60,
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the fan system 50 may have inlet guide vanes 74 that have variable vanes
configured
to modulate a flow of air 6 to the tip-fan 70. The amount and orientation of
the airflow
6 that is directed to the tip-fan 70 can be varied by suitably moving a
portion of the
IGVs 74 to vary the stagger angles using known actuators 75.
FIG. 2 shows an exemplary embodiment of a gas turbine engine 110
comprising a multistage fan 112 having multiple forward fan stages 152
configured to
pressurize an airflow 1. Although three forward fan stages 152 are shown in
the
exemplary engine 110 shown in FIG. 2, any suitable number of forward fan
stages for
a particular application can be selected. The forward fan stages pressurize
the flow
stream 1 entering the fan to generate a pressurized flow stream 2. The forward
fan
stages are driven by a low-pressure rotor 128 comprising a low-pressure
turbine 124
and a low-pressure turbine shaft 125. The gas turbine engine 110 further
comprises a
compressor 118 driven by a high-pressure rotor 129 having a high-pressure
turbine
112 and a high-pressure shaft 123. The HPC 118 has a rotor 19 that rotates
about the
longitudinal centerline axis 11 and pressurizes the air 8 flowing through the
core. The
HPC system includes a plurality of stator vanes arranged in a circumferential
direction
around the longitudinal centerline axis 11 (see FIG. 1 for example). The HPC
118
further includes multiple rotor stages 119 which have corresponding rotor
blades 140
extending radially outwardly from a rotor hub 139 or corresponding rotors in
the form
of separate disks, or integral blisks, or annular drums in any conventional
manner. The
high-pressure rotor 129 is supported in the engine static frames using known
support
methods using suitable bearings. The high-pressure turbine 122 and low-
pressure
turbine 124 are driven by combustion gases generated in the combustor 120 that
exit
as a hot exhaust stream 92.
The exemplary embodiment of a gas turbine engine 110 comprises an annular
splitter 146 (see FIG. 2) located axially aft from the axially last forward
fan stage 152.
The splitter 146 is adapted to bifurcate the pressurized flow stream 2 from
the forward
fan stage 152 to form the first portion 3 and the second portion 4 of the
pressurized
flow 2.
The exemplary embodiment of a gas turbine engine 110 comprises an aft fan
stage 160 located axially aft from the splitter 146, and axially forward from
the
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compressor 118, as shown in FIG. 2. As shown in FIG. 2, the aft fan stage 160
has a
tip-fan 170 configured to pressurize a first portion 3 of a pressurized air
flow 2 from
the forward fan stage 152. The tip-fan 170 is driven by the aft blade inner
portion 163
that acts as an air turbine blade 182. The aft fan stage 160, with the tip-fan
170, is
driven by a second portion 4 of the pressurized airflow 2. The inner portion
163 of the
aft fan blade 162 is configured to work as an air-turbine blade that can
extract energy
from a pressurized air stream whereas the outer portion 164 of the aft fan
blade 162 is
configured to be a compression-type airfoil that can pressurize an airflow.
The inner
portion 163 is an air turbine blade 182 having a turbine airfoil 184 adapted
to extract
energy from a pressurized flow of air. The outer portion of the aft fan blade
162 is
alternatively referred to herein as a tip-fan blade 172. The tip-fan blade 172
is capable
of pressuring a flow of air 6 to create a pressurized tip flow 56 (see FIG. 1
for
example). The aft fan stage 160 reduces the pressure and temperature of the
pressurized airflow that drives the aft fan stage 160. Known air-turbine
airfoil shapes,
materials and manufacturing methods can be used in the construction of the
inner
portion 163 aft fan blade 162. As the air flows over the inner portion 163, it
expands
to form an outflow 57 of air that has a lower pressure and lower temperature
and
imparts energy to the aft fan blades 162 to drive the aft fan stage 160.
The exemplary gas turbine engine I10 shown in FIG. 2 further comprises a
circumferential row of inlet guide vanes (IGVs) 174 located axially forward
from the
tip-fan blades 172. Known airfoil shapes, materials and manufacturing methods
can be
used in constructing the IGVs 174. The IGVs 174 control the volume of flow of
air
into the tip-fan 170, similar to the arrangement shown in FIG. 1. For enhanced
control
of the flow of air into the tip-fan 170, the inlet guide vanes 174 are
variable vanes that
are configured to modulate the flow of air to the tip-fan 70. The amount and
orientation of the airflow that is directed to the tip-fan 710 can be varied
by varying
the stagger angles by suitably moving a portion of the IGVs 174 using known
actuators 175.
In one aspect of the present invention, the exemplary gas turbine engine 110
shown in FIG. 2 (and FIG. 1) further comprises an annular inner bypass passage
142
adapted to flow an inner bypass flow 56 and an annular outer bypass passage
144
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adapted to flow an outer bypass flow 5. The outer bypass flow 5 passes through
the
outer bypass passage 144 and is not pressurized by the tip-fan 170. The inner
bypass
flow 6 (see FIG. 1) is pressurized by the tip-fan 170 and exits as pressurized
tip now
56. A forward mixer 148 located downstream from the aft fan stage 160 is
provided to
enhance mixing of the higher pressure inner bypass flow 56 and the lower
pressure
outer bypass flow 5 to form a mixed bypass flow 9 and developing a static
pressure
balance. Known mixers (alternatively known as Variable Area Bypass Injectors,
or
VABI, in the art) can be used for the mixer 148. A reverse flow in the outer
bypass
passage 144 can be prevented by using a known blocker door 145 that is located
near
the forward area of the outer bypass passage 144. During operation of the
engine, the
blocker door is operated toward closure when the variable IGV 144 is opened to
cause
further pressurization by the tip-fan 170. The gas turbine engine 110 further
comprises
a rear mixer 94 (alternatively known as Variable Area Bypass Injectors, or
VABI, in
the art) located down-stream from the low-pressure turbine 24 that is adapted
to
enhance mixing of the hot exhaust 92 from the low-pressure turbine 24 and the
relatively cooler bypass air flow stream 91. Known mixers (VABIs) can be used
for
this purpose. During engine operation, the operability of the forward fan
stage 152 and
the aft fan stage 160 can be controlled as necessary by suitably scheduling,
using
known methods, the operation of the variable IGVs 144, blocker door 145,
forward
mixer 148 and the rear VABI 194.
As shown in FIGS. 1 and 2, the aft fan stage 60, 160 (alternatively referred
to
herein as an intermediate pressure fan stage or IPFS) is a separate,
independently
rotating, spool that incorporates a tip-fan 70, 170 unlike the core driven fan
stages that
are coupled to the core spools known in the art. Further, as described herein,
the IPFS
has a tip-fan blade 72, 172 in its outer portion and a air turbine blade 82,
182 in the
inner portion. The IPFS spool is located between the forward fan 52, 152 and
the HPC
18, 118 such that part of the fan air is delivered to the tip of the IPFS
where its
pressure is further increased by the IPFS tip-fan blade 72, 172 and then
delivered to
the inner bypass passage 42, 142. The inner portion 4 of the forward fan flow
2 is
delivered to the turbine blade 82, 182 in the inner portion of the IPFS where
it is
expanded to provide the power to drive the fan tip. The flow from the exit of
the
turbine is delivered to the entrance of the HPC 18, 118. The extraction of
energy by
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the IPFS turbine blade 82, 182 reduces the boost pressure and temperature into
the
HPC 18, 118 below those at the forward fan exit 52, 152. By judicious choice
of
forward fan 52, 152 and IPFS tip-fan 70, 170 pressure ratios, the inlet
conditions to
the high pressure compressor 18, 118 can be matched to the originating
(baseline)
engine design conditions and maximize the use of the core flow capability by
the
derivative engine. At the same time the forward fan 52, 152 and IPFS 60, 160
provide
the desired higher bypass air pressure for the bypass flow 9.
Cycle studies have shown that the thrust potential for an existing core can be
increased up to 20% over a mixed flow turbofan derivative at the same fan
airflow
size. Temperature levels into the HPC can readily be matched to the original
hardware
design conditions allowing maximum use of the corrected flow capability within
the
original core mechanical design limits. Those skilled in the art will
recognize that
flowpath architecture studies using known methods can be performed to
establish the
required mounting structure for the IPFS and the aerodynamic design properties
of the
fan tip and turbine hub. In the exemplary embodiments shown herein, the IPFS
is.
preferably mounted within the fan frame structure, thus requiring no
additional main
engine frames to mount the additional spool.
Referring to FIGS. 1 and 2, an exemplary method of operating the fan system
50, 150 comprises the following steps. An airflow 1 that is flowing into the
fan system
50, 150 is pressurized in a forward fan stage 52, 152 to generate a
pressurized flow 2
that exits from the forward fan stage. The pressurized flow 2 is bifurcated to
a first
portion 3 and a second portion 4 using a suitable means, such as for example,
using an
annular splitter 46, 146. The first portion 3 of the pressurized airflow 2 is
then directed
towards a tip-fan 70 of an aft fan stage 60. A portion of the pressurized flow
2 is
flown through an outer bypass passage 44, 144 creating an outer bypass flow 5.
The
aft fan stage 60 rotates independently from the forward fan stage 52. The
second
portion 4 of the pressurized airflow 2 is directed towards a circumferential
row of air-
turbine blades 82 of the aft fan stage 60 such that the aft fan stage 60 is
driven by the
pressurized air. At this time, a higher pressure inflow 7 entering the inner
portion 63,
163 of aft fan stage 60 is expanded to a lower pressure outflow 57. During
this
expansion, the temperature of the expanding air flow in the inner portion 63,
163 of
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the aft stage 60, 160 drops. Thus, the temperature and pressure of a core flow
8
entering a compressor 18, 118 is reduced.
The exemplary method further comprises the step of pressurizing a flow 6
entering the tip-fan 70, 170 to generate a pressurized tip flow 56 (See FIGS.
1 and 2).
The flow of air 6 entering the tip-fan 70 is modulated with an inlet guide
vane 74, 174.
Specifically, the amount of air flowing through the tip-fan 70, 170 is
independently
controlled by the inlet guide vanes 74, 174. More specifically, a stagger of
the inlet
guide vanes 74, 174 is varied to selectively control the quantity of airflow
through the
tip-fan 70, 170, based on the fan pressure ratio, thrust and performance
requirements
of the engine 10, 110. The modulating of air 6 between substantially zero air
flow and
a maximum discharge air flow is performed as required by varying a stagger of
the
inlet guide vanes 74, 174. In the exemplary embodiment, the inlet guide vanes
74, 174
are mechanically actuated by known actuators 75, 175 and operated by a known
main
engine control system (not shown). In alternative embodiments, the inlet guide
vanes
74, 174 are operated by any suitable mechanism. Further, the exemplary method
comprises the step of mixing the outer bypass flow 5 in an annular outer
bypass
passage 44,144 with a tip-flow 56 from the tip-fan 70, 170 in an annular inner
bypass
passage 42, 142 to create a mixed bypass flow 9. A blocker door 45, 145
located near
the outer bypass passage 44, 144 is operated by modulating it between
partially closed
and substantially fully open positions so as to prevent a reverse flow in the
outer
bypass passage 144. Mechanical actuators operated by a known main engine
control
system (not shown) are used in the exemplary embodiment shown herein. The
method
described herein optionally comprises the step of operating a forward mixer
48, 148 of
a known type to control the mixing of the outer bypass flow 5 and the tip-flow
56 and
achieve a suitable static pressure balance. Further, the method comprises
operating a
rear mixer 94, 194 of a known type to control the operating characteristics of
the
forward fan stage 152 and the aft fan stage 160 and engine 10, 110
performance. The
forward mixer 48, 148, rear mixer 94, 194, the blocker door 45, 145 and the
inlet
guide vanes 74, 174 are operated in a controlled manner using an engine
control
system (not shown) in order to optimize the operating characteristics and
performance
of the engine 10,110.
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This written description uses examples to disclose the invention, including
the
best mode, and also to enable any person skilled in the art to make and use
the
invention. The patentable scope of the invention is defined by the claims, and
may
include other examples that occur to those skilled in the art. Such other
examples are
intended to be within the scope of the claims if they have structural elements
that do
not differ from the literal language of the claims, or if they include
equivalent
structural elements with insubstantial differences from the literal languages
of the
claims.
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