U.S. patent application number 14/988815 was filed with the patent office on 2016-06-30 for gas turbine engine with accessory gear box.
The applicant listed for this patent is United Technologies Corporation. Invention is credited to Thomas G. Cloft, Robert L. Gukeisen, Claude Mercier.
Application Number | 20160186598 14/988815 |
Document ID | / |
Family ID | 50929329 |
Filed Date | 2016-06-30 |
United States Patent
Application |
20160186598 |
Kind Code |
A1 |
Cloft; Thomas G. ; et
al. |
June 30, 2016 |
Gas Turbine Engine With Accessory Gear Box
Abstract
A nacelle for incorporation into a gas turbine engine has an
inner wall defining a bypass duct, and an outer wall, at least one
drive shaft extending through said inner wall, said at least one
drive shaft to be connected to a gas turbine engine receiving the
nacelle, said at least one drive shaft being connected to drive at
least two accessory gear boxes, with said at least two accessory
gear boxes being received between said inner and outer walls of
said nacelle. A gas turbine engine is also disclosed.
Inventors: |
Cloft; Thomas G.;
(Glastonbury, CT) ; Gukeisen; Robert L.;
(Middletown, CT) ; Mercier; Claude; (South
Windsor, CT) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
United Technologies Corporation |
Farmington |
CT |
US |
|
|
Family ID: |
50929329 |
Appl. No.: |
14/988815 |
Filed: |
January 6, 2016 |
Related U.S. Patent Documents
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Application
Number |
Filing Date |
Patent Number |
|
|
13719303 |
Dec 19, 2012 |
9297314 |
|
|
14988815 |
|
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|
Current U.S.
Class: |
60/786 ; 415/144;
60/805 |
Current CPC
Class: |
F01D 25/24 20130101;
F02C 7/32 20130101; F05D 2260/85 20130101; Y02T 50/60 20130101;
F02C 7/275 20130101; F02C 3/04 20130101; F01D 17/105 20130101; F01D
15/12 20130101; F01D 25/28 20130101; F02C 7/06 20130101; F05D
2220/32 20130101; F05D 2240/60 20130101; F05D 2260/4031 20130101;
F02K 3/06 20130101; F02C 7/36 20130101; F01D 25/20 20130101; F02C
7/268 20130101; F02C 7/22 20130101 |
International
Class: |
F01D 15/12 20060101
F01D015/12; F02C 7/268 20060101 F02C007/268; F01D 25/24 20060101
F01D025/24; F02C 7/06 20060101 F02C007/06; F01D 17/10 20060101
F01D017/10; F02C 3/04 20060101 F02C003/04; F02C 7/22 20060101
F02C007/22 |
Claims
1. A nacelle for incorporation into a gas turbine engine,
comprising: a nacelle having an inner wall defining a bypass duct,
and an outer wall, at least one drive shaft extending through said
inner wall, said at least one drive shaft to be connected to a gas
turbine engine receiving the nacelle, said at least one drive shaft
being connected to drive at least two accessory gear boxes, with
said at least two accessory gear boxes being received between said
inner and outer walls of said nacelle.
2. The nacelle as set forth in claim 1, wherein at least one of
said gear boxes includes a starter for the gas turbine engine.
3. The nacelle as set forth in claim 1, wherein at least one of
said at least two gear boxes drives a fuel pump and a lubricant
pump to support the gas turbine engine.
4. The gas turbine engine as set forth in claim 1, wherein said at
least one drive shaft includes at least two drive shafts, with one
of said at least two drive shafts being connected to each said at
least two accessory gear boxes.
5. The nacelle as set forth in claim 1, wherein said at least two
accessory gear boxes are received axially between an upstream end
and a downstream end of said nacelle.
6. A gas turbine engine comprising: a fan, a compressor section, a
combustor section, and a turbine section, said turbine section for
driving a shaft to in turn drive said fan and said compressor, at
least one accessory shaft to be connected for rotation with said
turbine section; a nacelle having an inner wall defining a bypass
duct receiving bypass air from said fan, and an outer wall, said at
least one drive shaft extending through said inner wall, and being
connected to drive at least two accessory gear boxes, with said at
least two accessory gear boxes being received between said inner
and outer walls of said nacelle.
7. The engine as set forth in claim 7, wherein at least one of said
gear boxes includes a starter for the gas turbine engine.
8. The engine as set forth in claim 7, wherein at least one of said
at least two gear boxes drives a fuel pump and a lubricant pump to
support the gas turbine engine.
9. The engine as set forth in claim 6, wherein said at least one
drive shaft includes at least two drive shafts, with one of said at
least two drive shafts being connected to each said at least two
accessory gear boxes.
10. The engine as set forth in claim 6, wherein said at least two
accessory gear boxes are received axially between an upstream end
and a downstream end of said nacelle.
11. The engine as set forth in claim 6, wherein said engine can be
divided into two halves, and a mount bracket for mounting said
engine to an aircraft is in one of the halves, and said at least
two accessory gear boxes are in an opposed half.
12. A gas turbine engine comprising: a fan, a compressor section, a
combustor section, and a turbine section, said turbine section for
driving a shaft to in turn drive said fan and said compressor, at
least one accessory shaft to be connected for rotation with said
turbine section; a nacelle having an inner wall defining a bypass
duct receiving bypass air from said fan, and an outer wall, said at
least one drive shaft extending through said inner wall, and being
connected to drive at least two accessory gear boxes, with said at
least two accessory gear boxes being received between said inner
and outer walls of said nacelle; at least one of said gear boxes
includes a starter for the gas turbine engine; at least another of
said at least two gear boxes drives a fuel pump and a lubricant
pump to support the gas turbine engine; said at least two accessory
gear boxes are received axially between an upstream end and a
downstream end of said nacelle; and said engine can be divided into
two halves, and a mount bracket for mounting said engine to an
aircraft is in one of the halves, and said at least two accessory
gear boxes are in an opposed half.
13. The engine as set forth in claim 12, wherein said at least one
drive shaft includes at least two drive shafts, with one of said at
least two drive shafts being connected to each said at least two
accessory gear boxes.
Description
CROSS-REFERENCE TO RELATED APPLICATION
[0001] This application is a continuation of U.S. application Ser.
No. 13/719,303, filed Dec. 19, 2012.
BACKGROUND OF THE INVENTION
[0002] This application relates to a gas turbine engine, wherein an
accessory gear box is split into at least two portions driven by
the main gas turbine engine, and positioned within a nacelle.
[0003] Gas turbine engines are known, and typically include a fan
delivering air into a bypass duct, and into a core engine. Air in
the core engine passes through a compressor which compresses the
air and delivers it into a combustor section. The air is mixed with
fuel and ignited. Products of this combustion pass downstream over
turbine rotors.
[0004] The turbine rotors are driven to rotate, and drive the fan
and compressor. In addition, power from the turbine rotors rotation
is utilized to generate electricity, and to drive accessories to
support the operation of the gas turbine engine.
[0005] As an example, pumps for supplying liquid to the gas turbine
engine are driven as accessories.
[0006] Another accessory is a starter. The starter is typically
provided with a fluid drive, which drives turbine rotors to rotate
the starter, and the starter begins to rotate other accessories and
the gas turbine engine thru a drive shaft. The same drive shaft
operates to drive other accessories, all through a gear box.
[0007] One known location for the accessory gear box is between an
inner and outer wall of the nacelle. Typically, a single gear box
has been provided at one circumferential location in the nacelle.
This has resulted in the radial thickness of the nacelle being
relatively large.
[0008] Recently a gear reduction has been provided between a
turbine that drives the fan and the fan rotor. One result of this
gear reduction, is that the fan rotor can be made much larger, and
the volume of bypass air can be greatly increased to increase
propulsion. To increase the fan diameter, the inner diameter of a
nacelle also increases. With such an increased inner diameter, it
becomes desirable to limit the radial thickness of the nacelle.
SUMMARY OF THE INVENTION
[0009] In a featured embodiment of this invention, a nacelle for
incorporation into a gas turbine engine has an inner wall defining
a bypass duct, and an outer wall, at least one drive shaft
extending through said inner wall, said at least one drive shaft to
be connected to a gas turbine engine receiving the nacelle, said at
least one drive shaft being connected to drive at least two
accessory gear boxes, with said at least two accessory gear boxes
being received between said inner and outer walls of said
nacelle.
[0010] In another embodiment according to the previous embodiment,
at least one of said gear boxes includes a starter for the gas
turbine engine.
[0011] In another embodiment according to any of the previous
embodiments, at least one of the at least two gear boxes drives a
fuel pump and a lubricant pump to support the gas turbine
engine.
[0012] In another embodiment according to any of the previous
embodiments, the at least one drive shaft includes at least two
drive shafts, with one of the at least two drive shafts being
connected to each of the at least two accessory gear boxes.
[0013] In another embodiment according to any of the previous
embodiments, at least two accessory gear boxes are received axially
between an upstream end and a downstream end of the nacelle.
[0014] In another featured embodiment, a gas turbine engine has a
fan, a compressor section, a combustor section, and a turbine
section. The turbine section drives a shaft to in turn drive the
fan and compressor. At least one accessory shaft is connected for
rotation with the turbine section. A nacelle has an inner wall
defining a bypass duct receiving bypass air from the fan, and an
outer wall. The at least one drive shaft extends through the inner
wall and is connected to drive at least two accessory gear boxes.
At least two accessory gear boxes are received between the inner
and outer walls of the nacelle.
[0015] In another embodiment according to the previous embodiment,
at least one of the gear boxes includes a starter for the gas
turbine engine.
[0016] In another embodiment according to any of the previous
embodiments, at least one of the at least two gear boxes drives a
fuel pump and a lubricant pump to support the gas turbine
engine.
[0017] In another embodiment according to any of the previous
embodiments, at least one drive shaft includes at least two drive
shafts, with one of the at least two drive shafts connected to each
of the at least two accessory gear boxes.
[0018] In another embodiment according to any of the previous
embodiments, at least two accessory gear boxes are received axially
between an upstream end and a downstream end of the nacelle.
[0019] In another embodiment according to any of the previous
embodiments, the engine can be divided into two halves. A mount
bracket mounts the engine to an aircraft in one of the halves. The
at least two accessory gear boxes are in an opposed half.
[0020] In another featured embodiment, a gas turbine engine has a
fan, a compressor section, a combustor section, and a turbine
section. The turbine section drives a shaft to in turn drive the
fan and compressor. At least one accessory shaft is connected for
rotation with the turbine section. A nacelle has an inner wall
defining a bypass duct receiving bypass air from the fan, and an
outer wall. The at least one drive shaft extends through the inner
wall, and is connected to drive at least two accessory gear boxes.
The at least two accessory gear boxes are received between the
inner and outer walls of the nacelle. At least one of the gear
boxes includes a starter for the gas turbine engine. At least
another of the said at least two gear boxes drives a fuel pump and
a lubricant pump to support the gas turbine engine. At least two
accessory gear boxes are received axially between an upstream end
and a downstream end of the nacelle. The engine can be divided into
two halves. A mount bracket mounts the engine to an aircraft in one
of the halves. At least two accessory gear boxes are in an opposed
half.
[0021] In another embodiment according to any of the previous
embodiments, at least one drive shaft includes at least two drive
shafts, with one of the at least two drive shafts connected to each
of the at least two accessory gear boxes.
[0022] These and other features of the invention may be best
understood from the following specification and drawings, the
following which is a brief description.
BRIEF DESCRIPTION OF THE DRAWINGS
[0023] FIG. 1 schematically shows a gas turbine engine.
[0024] FIG. 2 shows a first embodiment.
[0025] FIG. 3 is a side view of the first embodiment.
[0026] FIG. 4 is a front view of a second embodiment.
DETAILED DESCRIPTION
[0027] FIG. 1 schematically illustrates a gas turbine engine 20.
The gas turbine engine 20 is disclosed herein as a two-spool
turbofan that generally incorporates a fan section 22, a compressor
section 24, a combustor section 26 and a turbine section 28.
Alternative engines might include an augmentor section (not shown)
among other systems or features. The fan section 22 drives air
along a bypass flowpath B in a bypass duct defined within a nacelle
15, while the compressor section 24 drives air along a core
flowpath C for compression and communication into the combustor
section 26 then expansion through the turbine section 28. Although
depicted as a turbofan gas turbine engine in the disclosed
non-limiting embodiment, it should be understood that the concepts
described herein are not limited to use with turbofans as the
teachings may be applied to other types of turbine engines
including three-spool architectures.
[0028] The engine 20 generally includes a low speed spool 30 and a
high speed spool 32 mounted for rotation about an engine central
longitudinal axis A relative to an engine static structure 36 via
several bearing systems 38. It should be understood that various
bearing systems 38 at various locations may alternatively or
additionally be provided.
[0029] The low speed spool 30 generally includes an inner shaft 40
that interconnects a fan 42, a low pressure compressor 44 and a low
pressure turbine 46. The inner shaft 40 is connected to the fan 42
through a geared architecture 48 to drive the fan 42 at a lower
speed than the low speed spool 30. The high speed spool 32 includes
an outer shaft 50 that interconnects a high pressure compressor 52
and high pressure turbine 54. A combustor 56 is arranged between
the high pressure compressor 52 and the high pressure turbine 54. A
mid-turbine frame 57 of the engine static structure 36 is arranged
generally between the high pressure turbine 54 and the low pressure
turbine 46. The mid-turbine frame 57 further supports bearing
systems 38 in the turbine section 28. The inner shaft 40 and the
outer shaft 50 are concentric and rotate via bearing systems 38
about the engine central longitudinal axis A which is collinear
with their longitudinal axes.
[0030] The core airflow is compressed by the low pressure
compressor 44 then the high pressure compressor 52, mixed and
burned with fuel in the combustor 56, then expanded over the high
pressure turbine 54 and low pressure turbine 46. The mid-turbine
frame 57 includes airfoils 59 which are in the core airflow path.
The turbines 46, 54 rotationally drive the respective low speed
spool 30 and high speed spool 32 in response to the expansion.
[0031] The engine 20 in one example is a high-bypass geared
aircraft engine. In a further example, the engine 20 bypass ratio
is greater than about six (6), with an example embodiment being
greater than ten (10), the geared architecture 48 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 46 has a pressure ratio that is greater than about
5. In one disclosed embodiment, the engine 20 bypass ratio is
greater than about ten (10:1), the fan diameter is significantly
larger than that of the low pressure compressor 44, and the low
pressure turbine 46 has a pressure ratio that is greater than about
5:1. Low pressure turbine 46 pressure ratio is pressure measured
prior to inlet of low pressure turbine 46 as related to the
pressure at the outlet of the low pressure turbine 46 prior to an
exhaust nozzle. The geared architecture 48 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 2.5:1. It should be
understood, however, that the above parameters are only exemplary
of one embodiment of a geared architecture engine and that the
present invention is applicable to other gas turbine engines
including direct drive turbofans.
[0032] A significant amount of thrust is provided by the bypass
flow B due to the high bypass ratio. The fan section 22 of the
engine 20 is designed for a particular flight condition--typically
cruise at about 0.8 Mach and about 35,000 feet. The flight
condition of 0.8 Mach and 35,000 ft, with the engine at its best
fuel consumption--also known as "bucket cruise Thrust Specific Fuel
Consumption (`TSFC`)"--is the industry standard parameter of 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 a Fan Exit Guide Vane ("FEGV")
system. The low fan pressure ratio as disclosed herein according to
one non-limiting embodiment is less than about 1.45. "Low corrected
fan tip speed" is the actual fan tip speed in ft/sec divided by an
industry standard temperature correction of [(Tram .degree.
R)/(518.7 .degree. R)].sup.0.5. The "Low corrected fan tip speed"
as disclosed herein according to one non-limiting embodiment is
less than about 1150 ft/second.
[0033] A drive shaft 84 is shown schematically, and is operable to
be driven by one of the turbine sections 54 or 46, and to in turn
drive the turbine sections when driven by a starter. Typically, the
drive shaft 84 communicates with a gear box which may be mounted
within the nacelle.
[0034] FIG. 2 shows the nacelle 15 having an outer wall 82 and
inner wall 80. The bypass duct 200 is defined inwardly of the inner
wall 80. As shown, a take-off or drive shaft 84 is driven by the
turbine in the core engine 100. This may operate as described above
with regard to FIG. 1. The take-off shaft 84 extends through inner
wall 80 and to a splitter gear box 86 such that it drives
accessories in a first gear box 88. The gear box 88 takes in
rotation from a split shaft 87, and drives several accessories. As
an example, fuel pumps, oil pumps, etc. for enabling operation of
the gas turbine engine may be driven by the gear box 88. A second
gear box 90 is shown driven by a second split shaft 91. The gear
boxes 88 and 90 are shown to be circumferentially spaced. Gear box
90 may incorporate other accessories for supporting operation of
the gas turbine engine, and in particular may include a starter for
the gas turbine engine. A supply of air is shown to the starter
through a duct 92.
[0035] By splitting the gear box into two sub gear boxes 88 and 90,
the radial thickness between the inner 80 and outer 82 walls may be
smaller.
[0036] As can be appreciated from FIG. 3, both gear boxes 88 and 90
sit within an axial length between an upstream end 104 of the
nacelle 15 and a downstream end 102.
[0037] FIG. 4 shows a second embodiment wherein the single take-off
shaft 84 is replaced by a pair of shafts 106A and 106B extending
through the inner wall 80. Shaft 106A drives an accessory gear box
108A, while shaft 106B drives an accessory gear box 108B. Accessory
gear box 108B may include a starter, and accessory gear box 108A
may include several pumps, as mentioned with regard to the first
embodiment.
[0038] The FIG. 4 embodiment would be received within the same
axial envelope as that shown in FIG. 3.
[0039] Both of the embodiments would thus allow the outer wall 82
of the nacelle to be radially inward from what would typically be
the case if the prior art single accessory gear box were received
between walls 80 and 82.
[0040] As can be appreciated from FIGS. 2, 3, and 4, the two
accessory gear boxes are mounted in what will be a vertically lower
portion of the nacelle 15 once it is mounted on an aircraft.
[0041] As can be appreciated, there is a mount bracket 301 which is
associated with the engine, and which is mounted to a pylon
302/303, which is typically part of the aircraft. As shown in FIG.
2, a horizontal center line C separates the engine into a
vertically upper half 307 and a vertically lower half 305. The
accessory gear boxes 88/90 or 108A/108B are both in the half 305,
which will be vertically lower when the gas turbine engine is
mounted on an aircraft. However, it can also be said that a center
line can divide the engine into two halves, with the mount bracket
301 mounted in one of the halves, and the accessory gear boxes
88/90 or 108A/108B are in a different half
[0042] Although embodiments of this invention have been disclosed,
a worker of ordinary skill in this art would recognize that certain
modifications would come within the scope of this invention. For
that reason, the following claims should be studied to determine
the true scope and content of this invention.
* * * * *