U.S. patent application number 16/730497 was filed with the patent office on 2020-06-11 for turbine engine case attachment and a method of using the same.
The applicant listed for this patent is United Technologies Corporation. Invention is credited to Darin S. Lussier, James T. Roach.
Application Number | 20200182153 16/730497 |
Document ID | / |
Family ID | 56367208 |
Filed Date | 2020-06-11 |
United States Patent
Application |
20200182153 |
Kind Code |
A1 |
Roach; James T. ; et
al. |
June 11, 2020 |
TURBINE ENGINE CASE ATTACHMENT AND A METHOD OF USING THE SAME
Abstract
The present disclosure relates generally to a turbine engine
case assembly, the turbine engine case assembly including a fan
case including an outer frame encircling an axis, the outer frame
including an outer frame exterior surface and an outer frame
interior surface, at least two mounts disposed on the outer frame
exterior surface, wherein the at least two mounts are
circumferentially spaced along the outer frame exterior surface,
and a compliant attachment device operably coupled to the at least
two mounts.
Inventors: |
Roach; James T.; (Vernon,
CT) ; Lussier; Darin S.; (Guilford, CT) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
United Technologies Corporation |
Farmington |
CT |
US |
|
|
Family ID: |
56367208 |
Appl. No.: |
16/730497 |
Filed: |
December 30, 2019 |
Related U.S. Patent Documents
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Application
Number |
Filing Date |
Patent Number |
|
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14874049 |
Oct 2, 2015 |
10519863 |
|
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16730497 |
|
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62087474 |
Dec 4, 2014 |
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Current U.S.
Class: |
1/1 |
Current CPC
Class: |
F01D 21/045 20130101;
Y02T 50/44 20130101; Y02T 50/671 20130101; F05D 2240/90 20130101;
F05D 2260/311 20130101; F02C 7/20 20130101; Y02T 50/40 20130101;
B64D 27/26 20130101; F01D 25/28 20130101; Y02T 50/60 20130101 |
International
Class: |
F02C 7/20 20060101
F02C007/20; F01D 25/28 20060101 F01D025/28; B64D 27/26 20060101
B64D027/26; F01D 21/04 20060101 F01D021/04 |
Goverment Interests
STATEMENT REGARDING FEDERALLY SPONSORED RESEARCH OR DEVELOPMENT
[0002] This invention was made with support of the government by
the United States Air Force under contract number
FA8650-09-D-2923-D00021. The government therefore has certain
rights in this invention.
Claims
1. A method of reducing load transfer from a turbine engine case to
a portion of an airframe, the method comprising: securing a
compliant attachment device to the turbine engine case to create a
turbine engine case assembly; and securing the turbine engine case
assembly to the portion of the airframe.
2. The method of claim 1, wherein the turbine engine case
comprises: an outer frame encircling an axis, the outer frame
including an outer frame exterior surface and an outer frame
interior surface; at least two mounts disposed on the outer frame
exterior surface, wherein the at least two mounts are
circumferentially spaced along the outer frame exterior
surface.
3. The method of claim 2, wherein securing the compliant attachment
device to the turbine engine case comprises: coupling the compliant
attachment device between the at least two mounts disposed on the
outer frame exterior surface of the outer frame.
4. The method of claim 1, wherein the compliant attachment device
comprises a spring.
5. The method of claim 4, wherein the spring is a leaf spring.
6. The method of claim 1, wherein the compliant attachment device
is frangible.
7. The method of claim 2 wherein the turbine engine case assembly
further comprises: a center frame encircling the axis, wherein the
center frame is positioned radially inward from the outer frame;
and a plurality of struts, each coupled at a first end to the
center frame, and at a second end to the outer frame interior
surface.
8. A method of reducing load transfer from a turbine engine case to
a portion of an airframe of a gas turbine engine, the method
comprising: securing a compliant attachment device to the turbine
engine case to create a turbine engine case assembly; and securing
the turbine engine case assembly to the portion of the airframe,
wherein the gas turbine engine comprises: a fan section; a low
pressure compressor; and the turbine engine case assembly is
disposed aft the fan section; and wherein the turbine engine case
assembly comprises: an outer frame encircling an axis, the outer
frame including an outer frame exterior surface and an outer frame
interior surface; at least two mounts operably coupled to the outer
frame exterior surface, wherein the at least two mounts are
circumferentially spaced along the outer frame exterior surface;
and the compliant attachment device is operably coupled to the at
least two mounts.
9. The method of claim 8, wherein the compliant attachment device
comprises a spring.
10. The method of claim 9, wherein the spring is a leaf spring.
11. The method of claim 8, wherein the compliant attachment device
is frangible.
12. The method of claim 8, wherein the turbine engine case assembly
further comprises: a center frame encircling the axis, wherein the
center frame is positioned radially inward from the outer frame;
and a plurality of struts, each coupled at a first end to the
center frame, and at a second end to the outer frame interior
surface.
Description
CROSS REFERENCE TO RELATED APPLICATIONS
[0001] This application is a divisional of U.S. patent application
Ser. No. 14/874,049 filed on Oct. 2, 2015, which claims the benefit
of U.S. Provisional Patent Application Ser. No. 62/087,474 filed on
Dec. 4, 2014, the entire contents each of which are incorporated
herein by reference thereto.
TECHNICAL FIELD OF THE DISCLOSED EMBODIMENTS
[0003] The present disclosure is generally related to turbine
engine cases, in particular to a turbine engine attachment and a
method of using the same.
BACKGROUND OF THE DISCLOSED EMBODIMENTS
[0004] Stationary struts within a gas turbine engine function to
support the inner portion, or core. Additionally, the struts may
function as an airfoil. These struts, may be radially disposed
between an inner hub and an outer casing, and may be spaced around
the circumference of the rotor section in either a symmetrical or
an asymmetrical arrangement. The strut design provides the
structure with the stiffness required to maintain fit, form, and
function against loads, including but not limited to, those caused
by maneuvers, fan blade out, impinging gas loads, surge, and may
provide the ability to withstand both hard and soft body
impact.
[0005] Loads are generally transmitted through the hub-strut-case
structure through the mounts to the airframe via links or similar
features. Generally, in a situation where the loads are excessively
large (e.g., when the engine has suffered a fan blade-out event),
the attachment links transmit excessively large dynamic loads into
the aircraft. As a result, the system may experience a flight
safety event, such as, the aircraft must be taken out of service in
order to repair and/or replace the case and other necessary
components.
[0006] Improvements in turbine case attachments are therefore
needed in the art.
SUMMARY OF THE DISCLOSED EMBODIMENTS
[0007] In one aspect, a turbine engine case assembly is provided.
The turbine engine case assembly includes a fan case including an
outer frame encircling the axis. The outer frame includes an outer
frame exterior surface and an outer frame interior surface. At
least two mounts, circumferentially spaced along the outer frame
exterior surface, are disposed on the outer frame exterior surface.
In one embodiment, the turbine engine case assembly further
includes a center frame encircling the axis A, wherein the center
frame is positioned radially inward from the outer frame, and a
plurality of struts, each coupled at a strut first end to the
center frame and at a strut second end to the outer frame interior
surface.
[0008] The turbine engine case assembly further includes a
compliant attachment device operably coupled to the case at the at
least two mounts. In one embodiment, the compliant attachment
device is a spring. In one embodiment, the spring is a leaf spring.
In one embodiment, the compliant attachment device is
frangible.
[0009] In one aspect, a method for reducing load transfer on a
turbine engine case is provided. The method includes the step of
securing a compliant attachment device to a turbine engine case to
produce a turbine engine case assembly. In one embodiment, securing
a compliant attachment device to a turbine engine case includes
coupling the compliant attachment device between the at least two
mounts disposed on the outer frame exterior surface. The method
further includes the step of securing the turbine engine case to an
air frame.
[0010] Other embodiments are also disclosed.
BRIEF DESCRIPTION OF THE DRAWINGS
[0011] The embodiments and other features, advantages and
disclosures contained herein, and the manner of attaining them,
will become apparent and the present disclosure will be better
understood by reference to the following description of various
exemplary embodiments of the present disclosure taken in
conjunction with the accompanying drawings, wherein:
[0012] FIG. 1 is a sectional view of one example of a gas turbine
engine in which the presently disclosed embodiments may be
used;
[0013] FIG. 2 is a perspective view of a turbine engine case
assembly used in a gas turbine engine in one embodiment;
[0014] FIG. 3 is a sectional view of one example of a gas turbine
engine and a turbine engine case assembly in one embodiment;
[0015] FIG. 4 is a front view of a turbine engine case assembly
used in a gas turbine engine in one embodiment;
[0016] FIG. 5 is a front view of a turbine engine case assembly
used in a gas turbine engine in one embodiment; and
[0017] FIG. 6 is a schematic flow diagram of an embodiment of a
method of reducing load transfer on a turbine engine case in one
embodiment.
DETAILED DESCRIPTION OF THE DISCLOSED EMBODIMENTS
[0018] For the purposes of promoting an understanding of the
principles of the present disclosure, reference will now be made to
the embodiments illustrated in the drawings, and specific language
will be used to describe the same. It will nevertheless be
understood that no limitation of the scope of this disclosure is
thereby intended.
[0019] FIG. 1 shows a gas turbine engine 20, such as a gas turbine
used for power generation or propulsion, circumferentially disposed
about an engine centerline, or axial centerline axis A. 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 flow path
B in a bypass duct, while the compressor section 24 drives air
along a core flow path C for compression and communication into the
combustor section 26 then expansion through the turbine section 28.
Although depicted as a two-spool 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 two-spool
turbofans as the teachings may be applied to other types of turbine
engines including three-spool architectures.
[0020] The exemplary 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, and the location of bearing systems 38
may be varied as appropriate to the application.
[0021] 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 speed change mechanism, which in exemplary gas turbine
engine 20 is illustrated as 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 in exemplary gas turbine 20 between the high pressure
compressor 52 and the high pressure turbine 54. An engine static
structure 36 is arranged generally between the high pressure
turbine 54 and the low pressure turbine 46. The engine static
structure 36 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.
[0022] 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 turbines 46,
54 rotationally drive the respective low speed spool 30 and high
speed spool 32 in response to the expansion. It will be appreciated
that each of the positions of the fan section 22, compressor
section 24, combustor section 26, turbine section 28, and fan drive
gear system 48 may be varied. For example, gear system 48 may be
located aft of combustor section 26 or even aft of turbine section
28, and fan section 22 may be positioned forward or aft of the
location of gear system 48.
[0023] 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 about 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 five. 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 five 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 2.3:1. It
should be understood, however, that the above parameters are only
exemplary of one embodiment of a geared architecture engine and
that the present disclosure is applicable to other gas turbine
engines including direct drive turbofans.
[0024] 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 (10,688 meters). The
flight condition of 0.8 Mach and 35,000 ft. (10,688 meters), 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 (350.5
m/sec).
[0025] FIG. 2 illustrates a turbine engine case assembly, generally
indicated at 100. The turbine engine case assembly 100, depicted in
FIG. 2, is configured to be disposed about the fan section 22. For
example, the turbine engine case assembly 100, shown in FIG. 1, is
aft of the fan section 22. It will be appreciated that other
turbine engine case assemblies 100 may be disposed about other
sections of the gas turbine engine 20, for example the compressor
section 24 to name one non-limiting example. The turbine engine
case assembly 100 includes a fan case 102 including an outer frame
104 encircling the axis A. The outer frame 104 includes an outer
frame exterior surface 106 and an outer frame interior surface 108.
At least two mounts 110, configured to attach the turbine engine
case assembly 100 to an air frame 126 (see FIG. 3), are disposed on
the outer frame exterior surface 106. The at least two mounts 110
are circumferentially spaced along the outer frame exterior surface
106. It will be appreciated that more than two mounts 110 may be
circumferentially spaced along the outer frame exterior surface
106. In one embodiment, the turbine engine case assembly 100
further includes a center frame 112 encircling the axis A, wherein
the center frame 112 is positioned radially inward from the outer
frame 104, and a plurality of struts 114, each coupled at a strut
first end 116 to the center frame 112, and at a strut second end
118 to the outer frame interior surface 108.
[0026] The turbine engine case assembly 100 further includes a
compliant attachment device 120. The compliant attachment device
120 is configured to reduce the loads transmitted to the supporting
aircraft during an extreme load condition, for example, a fan blade
out or ultimate maneuver, to name a couple of non-limiting
examples, by dissipating some of the load in the compliant
attachment device 120 before it is transferred to the airframe 126.
The compliant attachment device 120 is operably coupled to the case
at the at least two mounts 110 in one embodiment. For example, a
compliant attachment device first end 122 may be operably coupled
to the mount 110A, and a compliant attachment device second end 124
may be operably coupled to the mount 110B. It will be appreciated
that the compliant attachment device 120 may be coupled to each of
the mounts 110 by any suitable means, for example, a nut and bolt
to name one non-limiting example.
[0027] In one embodiment, the compliant attachment device 120 is a
spring. In one embodiment, the spring includes a leaf spring. For
example, as shown in FIGS. 3-5, the turbine engine case assembly
100 is a fan inlet case, and during extreme load conditions, the
compliant attachment device 120 deflects in an outward and/or
inward direction to provide a dampening effect to reduce the
resonance of the turbine engine case assembly 100 and, hence, the
magnitude of the load transferred to the airframe 126. In one
embodiment, the compliant attachment device 120 is frangible. In
this embodiment, during extreme load conditions, the compliant
attachment device 120 is capable of breaking, in whole or in part,
to reduce the load transferred from the compliant attachment device
120 to the supporting aircraft. In other embodiments, the compliant
attachment device 120 may be any device operative to dissipate some
of the load in the compliant attachment device 120 before it is
transferred to the airframe 126.
[0028] FIG. 6 illustrates a method, generally indicated at 200, for
reducing load transfer on a turbine engine case. The method 200
includes the step 202 of securing a compliant attachment device 120
to a turbine engine case 102 to produce a turbine engine case
assembly 100. In one embodiment, securing a compliant attachment
device 120 to a turbine engine case 102 includes coupling the
compliant attachment device 120 between the at least two mounts 110
disposed on the outer frame exterior surface 106.
[0029] The method further includes the step 204 of securing the
turbine engine case 102 to a portion of an airframe 126. For
example, as shown in FIG. 3, the turbine engine case assembly 100
may be mounted to a portion of the airframe 126 using the at least
two mounts 110 disposed on the outer frame exterior surface
106.
[0030] It will be appreciated that as the compliant attachment
device 120 may be composed of a lightweight material to reduce the
overall weight of the aircraft. It will also be appreciated that
the compliant attachment device 120 may be operably coupled to the
turbine engine case 102 to provide compliance by deflecting the
impact load from an airframe to a localized component.
[0031] While the disclosure has been illustrated and described in
detail in the drawings and foregoing description, the same is to be
considered as illustrative and not restrictive in character, it
being understood that only certain embodiments have been shown and
described and that all changes and modifications that come within
the spirit of the invention are desired to be protected.
* * * * *