U.S. patent application number 12/870442 was filed with the patent office on 2011-06-23 for method of operating a fan system.
Invention is credited to JOHN LEWIS BAUGHMAN.
Application Number | 20110150627 12/870442 |
Document ID | / |
Family ID | 44151361 |
Filed Date | 2011-06-23 |
United States Patent
Application |
20110150627 |
Kind Code |
A1 |
BAUGHMAN; JOHN LEWIS |
June 23, 2011 |
METHOD OF OPERATING A FAN SYSTEM
Abstract
A method of operating a fan system is described comprising the
steps of pressurizing an airflow in a forward fan stage to generate
a pressurized flow, directing a first portion of the pressurized
airflow towards a tip-fan of an aft fan stage, and directing a
second portion of the pressurized airflow towards a circumferential
row of air-turbine blades of the aft fan stage to drive the aft fan
stage.
Inventors: |
BAUGHMAN; JOHN LEWIS;
(Cincinnati, OH) |
Family ID: |
44151361 |
Appl. No.: |
12/870442 |
Filed: |
August 27, 2010 |
Related U.S. Patent Documents
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Application
Number |
Filing Date |
Patent Number |
|
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61288366 |
Dec 21, 2009 |
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Current U.S.
Class: |
415/1 |
Current CPC
Class: |
F02C 9/20 20130101; F02K
3/075 20130101; F02K 3/077 20130101; F02C 3/064 20130101 |
Class at
Publication: |
415/1 |
International
Class: |
F04D 27/02 20060101
F04D027/02 |
Claims
1. A method of operating a fan system comprising the steps of:
pressurizing an airflow in a forward fan stage to generate a
pressurized flow; directing a first portion of the pressurized
airflow towards a tip-fan of an aft fan stage; and directing a
second portion of the pressurized airflow towards a circumferential
row of air-turbine blades of the aft fan stage to drive the aft fan
stage.
2. A method according to claim 1 wherein the tip-fan is driven by
the circumferential row of air-turbine blades.
3. A method according to claim 1 further comprising the step of
pressurizing a flow entering the tip-fan to generate a pressurized
tip flow.
4. A method according to claim 1 further comprising the step of
expanding a higher pressure inflow entering the aft fan stage to a
lower pressure outflow.
5. A method according to claim 1 further comprising the step of
expanding a higher pressure inflow entering the aft fan stage such
that the temperature of a core flow entering a compressor is
reduced.
6. A method according to claim 1 further comprising the step of
modulating the flow of air entering the tip-fan with an inlet guide
vane.
7. A method according to claim 5 further comprising modulating the
flow of air between substantially zero air flow and a maximum
discharge air flow.
8. A method according to claim 1 wherein the aft fan stage rotates
independently from the forward fan stage.
9. A method according to claim 1 further comprising the step of
flowing a portion of the pressurized flow in an outer bypass
passage to create an outer bypass flow.
10. A method according to claim 9 further comprising the step of
mixing the outer bypass flow with a tip-flow from the tip-fan to
create a mixed bypass flow.
11. A method of operating a gas turbine engine comprising the steps
of: pressurizing an airflow 1 using a forward fan stage to generate
a pressurized flow; directing a first portion 3 of the pressurized
airflow towards a tip-fan of an aft fan stage; and expanding a
second portion 4 of the pressurized airflow in a circumferential
row of air-turbine blades of the aft fan stage such that the
pressure of a core flow entering a compressor is reduced.
12. A method according to claim 11 wherein the aft fan stage
rotates independently from the forward fan stage.
13. A method according to claim 11 further comprising the step of
modulating a flow of air entering the tip-fan with an inlet guide
vane.
14. A method according to claim 13 further comprising modulating
the flow of air between substantially zero air flow and a maximum
discharge air flow.
15. A method according to claim 11 further comprising flowing an
inner bypass flow in an annular inner bypass passage and flowing an
outer bypass flow in an annular outer bypass passage.
16. A method according to claim 15 further comprising the step of
operating a blocker door to prevent a reverse flow in the outer
bypass passage.
17. A method according to claim 15 further comprising the step of
mixing the outer bypass flow with a tip-flow from the tip-fan to
create a mixed bypass flow.
18. A method according to claim 17 further comprising operating a
forward mixer to control the mixing of the outer bypass flow and
the tip-flow.
19. A method according to claim 11 further comprising operating a
rear mixer to control the operation of the forward fan stage and
the aft fan stage.
20. A method according to claim 19 further comprising operating a
blocker door located near an outer bypass passage.
Description
CROSS-REFERENCE TO RELATED APPLICATIONS
[0001] This Application claims priority to U.S. Provisional
Application Ser. No. 61/288,366, filed Dec. 21, 2009, which is
herein incorporated by reference in its entirety.
BACKGROUND OF THE INVENTION
[0002] This invention relates generally to gas turbine engines,
and, more specifically, to a method of operating a fan system
having an intermediate pressure fan stage having a tip fan located
on blades driven by a pressurized airflow.
[0003] 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.
[0004] 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 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.
[0005] Accordingly, it would be desirable to have a method of
operating 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
[0006] The above-mentioned need or needs may be met by exemplary
embodiments which provide a method of operating a fan system
comprising the steps of pressurizing an airflow in a forward fan
stage to generate a pressurized flow, directing a first portion of
the pressurized airflow towards a tip-fan of an aft fan stage, and
directing a second portion of the pressurized airflow towards a
circumferential row of air-turbine blades of the aft fan stage to
drive the aft fan stage. In one aspect of the invention, the aft
fan stage rotates independently from the forward fan stage.
BRIEF DESCRIPTION OF THE DRAWINGS
[0007] 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:
[0008] FIG. 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.
[0009] FIG. 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
[0010] 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.
[0011] 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.
[0012] 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.
[0013] 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.
[0014] 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.
[0015] 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 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.
[0016] 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 64 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
(item 165 in FIG. 2).
[0017] 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)
[0018] 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, 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.
[0019] 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.
[0020] 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.
[0021] 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 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.
[0022] The exemplary gas turbine engine 110 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.
[0023] 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 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 flow 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.
[0024] 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 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.
[0025] 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.
[0026] 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 the aft stage 60, 160 drops. Thus, the
temperature and pressure of a core flow 8 entering a compressor 18,
118 is reduced.
[0027] 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.
[0028] 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.
* * * * *