U.S. patent application number 11/469187 was filed with the patent office on 2007-06-07 for turbofan engine assembly and method of assembling same.
Invention is credited to Thomas Ory Moniz, Robert Joseph Orlando.
Application Number | 20070125066 11/469187 |
Document ID | / |
Family ID | 37491531 |
Filed Date | 2007-06-07 |
United States Patent
Application |
20070125066 |
Kind Code |
A1 |
Orlando; Robert Joseph ; et
al. |
June 7, 2007 |
TURBOFAN ENGINE ASSEMBLY AND METHOD OF ASSEMBLING SAME
Abstract
A method for assembling a gas turbine engine includes providing
a core gas turbine engine including a high-pressure compressor, a
combustor, and a turbine, coupling a counter-rotating fan assembly
to the core gas turbine engine such that air discharged from the
counter-rotating fan assembly is channeled directly into an inlet
of the gas turbine engine compressor, and coupling a
counter-rotating low-pressure turbine assembly to the
counter-rotating fan assembly.
Inventors: |
Orlando; Robert Joseph;
(West Chester, OH) ; Moniz; Thomas Ory; (Loveland,
OH) |
Correspondence
Address: |
JOHN S. BEULICK (12729);C/O ARMSTRONG TEASDALE LLP
ONE METROPOLITAN SQUARE
SUITE 2600
ST. LOUIS
MO
63102-2740
US
|
Family ID: |
37491531 |
Appl. No.: |
11/469187 |
Filed: |
August 31, 2006 |
Related U.S. Patent Documents
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Application
Number |
Filing Date |
Patent Number |
|
|
11254143 |
Oct 19, 2005 |
|
|
|
11469187 |
Aug 31, 2006 |
|
|
|
Current U.S.
Class: |
60/226.1 ;
60/268 |
Current CPC
Class: |
F05D 2230/60 20130101;
Y02T 50/60 20130101; F02K 3/072 20130101; F02C 7/36 20130101 |
Class at
Publication: |
060/226.1 ;
060/268 |
International
Class: |
F02K 3/072 20060101
F02K003/072 |
Claims
1. A method of assembling a turbofan engine assembly comprises:
providing a core gas turbine engine including a high-pressure
compressor, a combustor, and a turbine; coupling a counter-rotating
fan assembly to the core gas turbine engine such that air
discharged from the counter-rotating fan assembly is channeled
directly into an inlet of the gas turbine engine compressor; and
coupling a counter-rotating low-pressure turbine assembly to the
counter-rotating fan assembly.
2. A method in accordance with claim 1 wherein coupling further
comprises coupling a counter-rotating fan assembly to the core gas
turbine engine such that compressed air is discharged from the
counter-rotating fan assembly at a first operating pressure and
received at the core gas turbine high-pressure compressor at
approximately the first operational pressure.
3. A method in accordance with claim 1 wherein coupling further
comprises coupling a counter-rotating fan assembly including a
first fan assembly and a second fan assembly to the core gas
turbine engine such that the first fan assembly rotates in a first
direction and the second fan assembly rotates in an opposite second
direction.
4. A method in accordance with claim 3 further comprising: coupling
a first shaft between the first fan assembly and a first turbine
rotor that is configured to rotate in a first rotational direction;
and coupling a second shaft between the second fan assembly and a
second turbine rotor that is configured to rotate in a second
rotational direction that is opposite the first rotational
direction.
5. A method in accordance with claim 1 wherein coupling further
comprises coupling a counter-rotating fan assembly that discharges
a predetermined quantity of air based on the gas turbine engine
compression ratio to the core gas turbine engine.
6. A method in accordance with claim 1 wherein providing a core gas
turbine engine comprises providing a core gas turbine engine that
includes a predetermined quantity of compressor stages based on the
quantity of compressed air discharged from the counter-rotating fan
assembly.
7. A method in accordance with claim 1 further comprising coupling
a gooseneck between the counter-rotating fan assembly and the core
gas turbine engine to facilitate channeling air discharged from the
counter-rotating fan assembly to the core gas turbine engine.
8. A method in accordance with claim 7 further comprising orienting
the gooseneck within the turbofan engine assembly to facilitate
preventing particles having a predetermined mass from being
channeled in a radially inward direction into the core gas turbine
engine.
9. A turbofan engine assembly comprising: a core gas turbine engine
including a high-pressure compressor, a combustor, and a
high-pressure turbine; a counter-rotating fan assembly coupled to
said core gas turbine engine such that air discharged from said
counter-rotating fan assembly is channeled directly into an inlet
of said gas turbine engine compressor; and a counter-rotating
low-pressure turbine assembly coupled to said counter-rotating fan
assembly.
10. A turbofan engine assembly in accordance with claim 9 wherein
said counter-rotating fan assembly is selectively sized to
discharge compressed air at a first operating pressure, said core
gas turbine engine is configured to receive the compressed air at
approximately the first operational pressure.
11. A turbofan engine assembly in accordance with claim 9 wherein
said counter-rotating fan assembly comprises a first fan assembly
that rotates in a first direction and a second fan assembly rotates
that rotates in an opposite second direction.
12. A turbofan engine assembly in accordance with claim 9 wherein
said core gas turbine engine comprises a predetermined quantity of
compressor stages based on the compression ratio of said
counter-rotating fan assembly and the overall compression ratio of
the gas turbofan engine assembly.
13. A turbofan engine assembly in accordance with claim 11 further
comprising: a first shaft coupled between said first fan assembly
and a first turbine rotor that is configured to rotate in a first
rotational direction; and a second shaft coupled between said
second fan assembly and a second turbine rotor that is configured
to rotate in a second rotational direction that is opposite the
first rotational direction.
14. A turbofan engine assembly in accordance with claim 9 wherein
said core gas turbine engine comprises a predetermined quantity of
high-pressure turbine stages based on the compression ratio of said
counter-rotating fan assembly and the overall compression ratio of
the gas turbofan engine assembly.
15. A turbofan engine assembly in accordance with claim 9 further
comprising a gooseneck coupled between the counter-rotating fan
assembly and the core gas turbine engine to facilitate channeling
air discharged from the counter-rotating fan assembly to the core
gas turbine engine.
16. A turbofan engine assembly in accordance with claim 16 wherein
said gooseneck is configured to prevent particles having a
predetermined mass from being channeled in a radially inward
direction into said core gas turbine engine.
Description
CROSS REFERENCE TO RELATED APPLICATION
[0001] This application is a continuation-in-part of U.S. patent
application Ser. No. 11/254,143 filed Oct. 19, 2005.
BACKGROUND OF THE INVENTION
[0002] This invention relates generally to gas turbine engines, and
more specifically to gas turbine engine assemblies and methods of
assembling the same.
[0003] At least some known gas turbine engines include a forward
fan, a core engine, and a power turbine. The core engine includes
at least one compressor, a combustor, a high-pressure turbine and a
low-pressure turbine coupled together in a serial flow
relationship. More specifically, the compressor and high-pressure
turbine are coupled through a shaft to define a high-pressure rotor
assembly. Air entering the core engine is mixed with fuel and
ignited to form a high energy gas stream. The high energy gas
stream flows through the high-pressure turbine to rotatably drive
the high-pressure turbine such that the shaft, in turn, rotatably
drives the compressor.
[0004] The gas stream expands as it flows through the low-pressure
turbine positioned forward of the high-pressure turbine. The
low-pressure turbine includes a rotor assembly having a fan coupled
to a drive shaft. The low-pressure turbine rotatably drives the fan
through the drive shaft. To facilitate increasing engine
efficiency, at least one known gas turbine engine includes a
counter-rotating low-pressure turbine that is coupled to a
counter-rotating fan and a booster compressor.
[0005] An outer rotating spool, a rotating frame, a mid-turbine
frame, and two concentric shafts, are installed within the gas
turbine engine to facilitate supporting the counter-rotating
low-pressure turbine. The installation of the aforementioned
components also enables a first fan assembly to be coupled to a
first turbine and a second fan assembly to be coupled to a second
turbine such that the first fan assembly and the second fan
assembly each rotate in the same rotational direction as the first
turbine and the second turbine, respectively. Accordingly, the
overall weight, design complexity and/or manufacturing costs of
such an engine are increased.
BRIEF DESCRIPTION OF THE INVENTION
[0006] In one aspect, a method of assembling a gas turbine engine
is provided. The method includes providing a core gas turbine
engine including a high-pressure compressor, a combustor, and a
turbine, coupling a counter-rotating fan assembly to the core gas
turbine engine such that air discharged from the counter-rotating
fan assembly is channeled directly into an inlet of the gas turbine
engine compressor, and coupling a counter-rotating low-pressure
turbine assembly to the counter-rotating fan assembly.
[0007] In another aspect, a turbofan engine assembly is provided.
The turbofan engine assembly includes a core gas turbine engine
including a high-pressure compressor, a combustor, and a turbine,
and a counter-rotating fan assembly coupled to the core gas turbine
engine such that air discharged from the counter-rotating fan
assembly is channeled directly into an inlet of the gas turbine
engine compressor, and a counter-rotating low-pressure turbine
assembly coupled to the counter-rotating fan assembly.
BRIEF DESCRIPTION OF THE DRAWINGS
[0008] FIG. 1 is a cross-sectional view of a portion of an
exemplary turbofan engine assembly.
DETAILED DESCRIPTION OF THE INVENTION
[0009] FIG. 1 is a cross-sectional view of a portion of an
exemplary turbofan engine assembly 10 that includes a
counter-rotating fan assembly 11 that includes a forward or
upstream fan assembly 12 and an aft or downstream fan assembly 14
disposed about a longitudinal centerline axis 16. The terms
"forward fan" and "aft fan" are used herein to indicate that fan
assembly 12 is coupled axially upstream from fan assembly 14. In
one embodiment, fan assemblies 12 and 14 are positioned at a
forward end of turbofan engine assembly 10 as illustrated. In an
alternative embodiment, fan assemblies 12 and 14 are positioned
downstream from the core gas turbine engine. Fan assemblies 12 and
14 each include respective rotor disks 13 and 15 and a plurality of
fan blades 18 and 20 that are coupled to each respective rotor
disk. Fan assemblies 12 and 14 are each positioned within a nacelle
22.
[0010] Turbofan engine assembly 10 also includes a core gas turbine
engine 30 that is coupled downstream from fan assemblies 12 and 14.
Core gas turbine engine 30 includes a high-pressure compressor 32,
a combustor 34, and a high-pressure turbine 36 that is coupled to
high-pressure compressor 32 via a shaft 38. Core gas turbine engine
30 includes an outer casing 72 that defines an annular core engine
inlet 74.
[0011] Turbofan engine assembly 10 also includes a counter-rotating
low-pressure turbine 40. Low-pressure turbine 40 includes a
stationary outer casing 50 that is coupled to core gas turbine
engine 30 downstream from high-pressure turbine 36. Low-pressure
turbine 40 also includes a radially outer rotor section 52 that is
positioned radially inwardly of outer casing 50. Outer rotor
section 52 has a generally frusto-conical shape and includes a
plurality of circumferentially spaced rotor blades 54 that are
coupled to, and extend radially inwardly from, a respective rotor
disk 56. Although, the exemplary embodiment only illustrates four
rotor disks 56, it should be realized that outer rotor 52 may have
any quantity of rotor disks 56 without affecting the scope of the
method and apparatus described herein.
[0012] Low-pressure turbine 40 also includes a radially inner rotor
section 60 that is aligned substantially coaxially with respect to,
and radially inward of, outer rotor section 52. Inner rotor 60
includes a plurality of circumferentially spaced rotor blades 62
that are coupled to, and extend radially outwardly from, a
respective rotor disk 64. Although, the exemplary embodiment only
illustrates five rotor disks 64, it should be realized that inner
rotor 60 may have any quantity of rotor disks 64 without affecting
the scope of the method and apparatus described herein. In the
exemplary embodiment, inner rotor 60 is rotatably coupled to aft
fan assembly 14 via shaft 44 and also to turbine midframe 66 which
provides structural support for inner rotor 60. Outer rotor 52 is
rotatably coupled to a forward fan assembly 12 via shaft 42 and
also to turbine rear-frame 68 which provides rotational support to
outer rotor 52.
[0013] In the exemplary embodiment, inner rotor blades 62 extending
from a respective rotor disk 64 are axially interdigitated with
outer rotor blades 54 extending from a respective rotor disk 56
such that inner rotor blades 62 extend between respective outer
rotor blades 54. The blades 54 and 62 are therefore configured for
counter-rotation of the rotors 52 and 60. In one preferred
embodiment, low-pressure turbine outer rotor 52 and forward fan
assembly 12 are configured to rotate in a first rotational
direction, and low-pressure turbine inner rotor 60 and aft fan
assembly 14 are configured to rotate in a second opposite
direction.
[0014] In operation, air flows through fan assemblies 12 and 14
supplying the high pressure compressor 32 wherein the airflow is
further compressed and delivered to combustor 34. Fuel is added to
the high pressure air in combustor 34 and ignited, expanding to
drive high-pressure turbine 36, and low-pressure turbine 40 is
utilized to drive fan assemblies 12 and 14 by way of shafts 42 and
44, respectively. Turbofan engine assembly 10 is operable at a
range of operating conditions between design operating conditions
and off-design operating conditions.
[0015] In the exemplary embodiment, the counter-rotating fan
assembly 11 is sized to discharge a predetermined quantity of air
based on the gas turbine engine compression ratio. More
specifically, high-pressure compressor 32 includes a plurality of
stages 70 wherein each stage further increases the pressure from
the previous stage such that core gas turbine engine 30 has a
compression ratio based on the quantity of stages 70 utilized
within high-pressure compressor 32. Moreover, although a single
core gas turbine is illustrated, it should be realized that the
core gas turbine engine 30 may include a compressor having any
quantity of compression stages, and thus a wide variety of
compression ratios.
[0016] Accordingly, in one embodiment, core gas turbine engine 30
includes a plurality of compression stages 70 that are
predetermined based on the quantity and/or pressure of the
compressed air discharged from the counter-rotating fan assembly
11. For example, a core gas turbine engine having a first
compression ratio may be coupled to a counter-rotating fan assembly
11 having a first compression ratio. If the compression ratio of
counter-rotating fan assembly 11 is increased, the fan assembly 11
may be utilized with a core gas turbine engine 30 having a reduced
compression ratio. Optionally, if the compression ratio of the
counter-rotating fan assembly 11 is reduced, fan assembly 11 may be
utilized with a core gas turbine engine 30 that includes an
increased quantity of stages and thus has an increased compression
ratio. In the exemplary embodiment, high-pressure compressor 32
includes at least six compression stages 70. Therefore,
counter-rotating fan assembly 11 may be selectively sized to be
coupled to a wide variety of core gas turbine engines. Optionally,
a single core gas turbine engine compressor may be modified by
either increasing or decreasing the quantity of compression stages,
i.e. greater or lesser than six stages, to facilitate coupling the
core gas turbine engine to counter-rotating fan assembly 11.
[0017] In the exemplary embodiment, turbofan engine assembly 10
includes a gooseneck 78 that extends between and facilitates
coupling counter-rotating fan assembly 11 to core gas turbine
engine 30. Moreover, gooseneck 78 includes a structural strut
and/or aero strut 80 to facilitate channeling air discharged from
aft fan assembly 14, through gooseneck 78, to core gas turbine
engine 30. As such, the configuration of gooseneck 78 and the
structural strut facilitate substantially reducing and/or
eliminating ice and/or foreign particle ingestion into core gas
turbine engine 30 since core inlet gooseneck 78 substantially
"hides" the core gas turbine engine inlet from the main air
flowstream that is channeled axially past the exterior surface of
gooseneck 78 in an aftward or downstream direction. More
specifically, during operation, gooseneck 78 is configured or
oriented to divide the airstream discharged from the
counter-rotating fan assembly into a first airstream 82 and a
second airstream 84 to facilitate preventing particles having a
predetermined mass from flowing in a radially inward direction and
being ingested into the core gas turbine engine 30. Specifically,
gooseneck 78 is configured to channel particles having a heavier
mass, such as ice particles, or possible foreign object damage
(FOD) type material from being ingested into the core gas turbine
engine. As such, since the core gas turbine engine is "hidden" from
the airstream, gooseneck 78 channels the heavier particles around
the core engine to facilitate preventing damage to the core gas
turbine engine.
[0018] The turbofan engine assembly described herein includes a
counter-rotating fan assembly that is coupled to a counter-rotating
low-pressure turbine assembly. The turbofan engine assembly
described herein facilitates reducing at least some of the
complexities associated with known counter-rotating low-pressure
turbines. More specifically, the turbofan engine assembly described
herein includes a front fan assembly that is rotatably coupled to a
radially outer rotor section of the low-pressure turbine, and an
aft fan assembly that is rotatably coupled to a radially inner
rotor section of the low-pressure turbine.
[0019] Additionally, the above-described gas turbine engine does
not include a booster compressor. As a result, eliminating the
booster compressor results in a simpler, lower cost, and lower
weight engine than at least one known counter-rotating engine. More
specifically, the booster can be eliminated because a high-pressure
ratio core is used in conjunction with the increased core stream
pressure ratio that can be obtained with the two counter rotating
fans. The systems described herein facilitate optimizing the speed
ratio between the two counter-rotating fans to increase
performance. Moreover, since no booster stage count issues exists,
the interaction loss between the high-pressure turbine (HPT) and
the low-pressure turbine (LPT) is substantially eliminated thus
resulting in approximately 0.8% increase in LPT efficiency, the
two-stage HPT is approximately 3% more efficient than the known
single stage HPT thus increasing overall pressure ratio for
additional thermodynamic improvements. Additionally, no variable
bleed valves (VBV) bleed doors are utilized, and ice and foreign
particle ingestion is substantially eliminated because the
booster-less engine will allow the core inlet gooseneck to be
hidden.
[0020] Further, the two-stage HPT facilitates increasing the
capability of power extraction off the HP spool. The LPT power
requirements (Aero Dynamic Loading) are reduced by about 10%
resulting in either an improvement in efficiency and/or reduced
weight, a simpler thrust reverser design can be utilized,
additional space under the core cowl may he available to locate the
accessory gearbox and larger multiple generators, a shorter fan
case, and a simpler, lighter, thinner inlet fan duct.
[0021] Exemplary embodiments of a turbofan engine assembly and
method of assembling the turbofan engine assembly are described
above in detail. The assembly and method arc not limited to the
specific embodiments described herein, but rather, components of
the assembly and/or steps of the method may be utilized
independently and separately from other components and/or steps
described herein. Further, the described assembly components and/or
the method steps can also be defined in, or used in combination
with, other assemblies and/or methods, and are not limited to
practice with only the assembly and/or method as described
herein.
[0022] While the invention has been described in terms of various
specific embodiments, those skilled in the art will recognize that
the invention can be practiced with modification within the spirit
and scope of the claims.
* * * * *