Language selection

Search

Patent 2756914 Summary

Third-party information liability

Some of the information on this Web page has been provided by external sources. The Government of Canada is not responsible for the accuracy, reliability or currency of the information supplied by external sources. Users wishing to rely upon this information should consult directly with the source of the information. Content provided by external sources is not subject to official languages, privacy and accessibility requirements.

Claims and Abstract availability

Any discrepancies in the text and image of the Claims and Abstract are due to differing posting times. Text of the Claims and Abstract are posted:

  • At the time the application is open to public inspection;
  • At the time of issue of the patent (grant).
(12) Patent: (11) CA 2756914
(54) English Title: TURBOFAN MOUNTING ARRANGEMENT
(54) French Title: DISPOSITIF D'INSTALLATION DE TURBINE
Status: Granted
Bibliographic Data
(51) International Patent Classification (IPC):
  • F02C 7/20 (2006.01)
  • F02K 3/06 (2006.01)
(72) Inventors :
  • HEYERMAN, JEFFREY BERNARD (Canada)
  • OLVER, BRYAN (Canada)
  • CAULFEILD, STEPHEN (Canada)
(73) Owners :
  • PRATT & WHITNEY CANADA CORP. (Canada)
(71) Applicants :
  • PRATT & WHITNEY CANADA CORP. (Canada)
(74) Agent: NORTON ROSE FULBRIGHT CANADA LLP/S.E.N.C.R.L., S.R.L.
(74) Associate agent:
(45) Issued: 2019-01-22
(22) Filed Date: 2011-11-04
(41) Open to Public Inspection: 2013-05-04
Examination requested: 2016-10-27
Availability of licence: N/A
(25) Language of filing: English

Patent Cooperation Treaty (PCT): No

(30) Application Priority Data: None

Abstracts

English Abstract

A gas turbine engine has a rear mounting structure incorporating a mounting apparatus attached to a bypass duct wall with a link device of six rods for transferring core portion related inertia-induced loads, from an inner case of the core portion in a short circuit across an annular bypass air passage to the bypass duct wall.


French Abstract

Une turbine à gaz comporte une structure de montage arrière comprenant un appareil de montage fixé à la paroi dun conduit de dérivation avec un dispositif de liaison de six tiges pour transférer des charges induites par inertie liées à la partie centrale, à partir dun boîtier intérieur de la partie centrale dans un court circuit à travers un passage dair en dérivation annulaire vers la paroi du conduit de dérivation.

Claims

Note: Claims are shown in the official language in which they were submitted.



CLAIMS:

1. A turbofan gas turbine engine comprising:
a front mounting structure attached to an annular bypass duct wall at a front
axial
position adjacent an inlet of a bypass air passage, the bypass air passage
defined by and
radially between the bypass duct wall and an inner case of a core portion of
the engine, the
front mounting structure supporting the core portion within the bypass duct
wall;
a rear mounting structure haying a mounting apparatus attached to the bypass
duct
wall at a rear axial position adjacent to an outlet of the bypass air passage,
the rear mounting
structure including three pairs of first and second rods each haying opposed
inner and outer
ends, the rods extending across the bypass air passage and interconnecting the
bypass duct
wall and the inner case of the core portion, wherein the first rods extend
from the outer end to
the inner end thereof in substantially tangential directions with respect to
the core portion,
corresponding to a first circumferential direction, and the second rods extend
from the outer
end to the inner end thereof in substantially tangential directions with
respect to the core
portion, corresponding to a second circumferential direction opposite to the
first
circumferential direction; and
wherein the bypass duct wall includes two axially spaced flanges, radially and

outwardly extending from the bypass duct wall, three circumferentially spaced
apart
connecting brackets being attached to the bypass duct wall and positioned
axially between
and affixed to the two flanges, each pair of the rods being connected at the
outer ends thereof
to the bypass duct wall by one of the connecting brackets.
2. The engine as defined in claim 1 wherein the three pairs of rods are
substantially
identical.
3. The engine as defined in claim 1 wherein the three pairs of rods are
disposed in a
plane to which an engine central axis is substantially perpendicular.
4. The engine as defined in claim 1 wherein the three pairs of rods are
disposed to form
a substantially triangular structure within the bypass duct wall, the inner
case being disposed
within the triangular structure.

-12-


5. The engine as defined in claim 1 wherein the connecting brackets are
mounted to an
outer side of the bypass duct wall, at least one of the connecting brackets
having a radially
outwardly extending mounting portion with at least one mounting opening
defined therein.
6. The engine as defined in claim 5 wherein the connecting brackets are
accessible from
an inside of the bypass air passage through respective openings defined in the
bypass duct
wall for connection with the respective rods.
7. The engine as defined in claim 1 wherein each of the rods is connected
at the inner
end thereof to the inner case by an adjustable connecting device in order to
ensure a
concentricity of the bypass duct wall and the core portion.
8. The engine as defined in claim 1 wherein each of the rods has an
aerodynamic profile
defined by side surfaces extending between leading and trailing edges with
respect to the
bypass air passage, the profile having a dimension between the side surfaces
smaller than a
dimension between the leading and trailing edges.
9. A turbofan gas turbine engine comprising:
a core portion including an inner case;
an annular bypass duct wall surrounding and supporting the core portion, the
bypass
duct wall and the inner case of the core portion defining a bypass air passage
radially between
the core portion and the bypass duct for directing a bypass air flow passing
therethrough;
a rod frame structure including three pairs of first and second rods defined
in a radial
plane defined by the inner case, the rod frame structure interconnecting the
bypass duct wall
and the inner case for transferring core portion related inertia-induced loads
from the inner
case in a short circuit across the bypass air passage in the radial plane to
the bypass duct wall,
thereby reducing distortion of the core portion caused by the inertia-induced
loads and
reducing carcass bending of the core portion; and
wherein the bypass duct wall includes two axially spaced flanges, radially and

outwardly extending from the bypass duct wall, three circumferentially spaced
apart
connecting brackets being attached to an outer side of the bypass duct wall
and being
positioned axially between and affixed to the two flanges.

-13-


10. The turbofan gas turbine engine as defined in claim 9 wherein the
radial plane is
substantially superposed on a vertical rear mounting plane of the engine, the
rear mounting
plane being located close to an outlet of the annular bypass air passage with
respect to a
substantially vertical front mounting plane of the engine which is located
close to an inlet of
the bypass air passage.
11. The turbofan gas turbine engine as defined in claim 10 wherein at least
one of the
connecting brackets comprises a mounting portion defining at least one
mounting opening
therein, the mounting portion extending radially outwardly from the bypass
duct wall and
disposed substantially in the vertical rear mounting plane of the engine.
12. The turbofan gas turbine engine as defined in claim 9 wherein the rod
frame structure
forms a substantially triangular structure within the bypass duct wall, the
inner case being
disposed within the triangular structure.
13. The turbofan gas turbine engine as defined in claim 9 wherein the first
and second
rods each comprise opposed inner and outer ends, the first rods extending from
the outer end
to the inner end thereof across the bypass air passage in substantially
tangential directions
with respect to the core portion, corresponding to a first circumferential
direction and the
second rods extending from the outer end to the inner end thereof across the
bypass air
passage in substantially tangential directions with respect to the core
portion, corresponding
to a second circumferential direction opposite to the first circumferential
direction.
14. The turbofan gas turbine engine as defined in claim 9 wherein the three
pair of rods
are substantially identical.
15. The turbofan gas turbine engine as defined in claim 9 wherein the
radial plane is
defined between vertical front and rear mounting planes of the engine, the
rear mounting
plane being located close to an outlet of the annular bypass air passage with
respect to a
vertical front mounting plane of the engine which is located close to an inlet
of the bypass air
passage.

-14-

Description

Note: Descriptions are shown in the official language in which they were submitted.


CA 02756914 2011-11-04


TURBOFAN MOUNTING ARRANGEMENT

CROSS REFERENCE TO RELATED APPLICATIONS
[0001] The present invention is a continuation-in-part of co-pending
application
number 12/466,426, filed on May 15, 2009.
TECHNICAL FIELD
[0002] The application relates generally to gas turbine engines and more
particularly, to a turbofan engine mounting system.
BACKGROUND OF THE ART
[0003] A turbofan gas turbine engine basically includes a core portion which
must
be mounted inside a bypass duct. A traditional engine mount system for a
fuselage
mount turbofan gas turbine engine reacts to thrust, lateral and vertical loads
at the front
mounting plane (on the intermediate case of the engine), and reacts to lateral
and
vertical loads at the rear mount. The rear mount is usually located either on
the bypass
duct, forming a cantilever core as schematically shown in FIG. 9, or on the
engine core,
typically near the turbine exhaust case, forming a rear core mount as
schematically
shown in FIG. 10. However, the cantilever core suffers from distortion due to
inertia
loads and tends to droop from the burden of these loads, resulting in tip
clearance loss
which is critical to the functioning of an axial compressor. The rear core
mount suffers
from significant bending of the core portion caused by thrust loads. The rear
mount
carries a load due to a moment created by the engine thrust line of action
being offset
from the thrust reaction plane. Thus, the core portion is loaded analogous to
a simply
supported beam with a point moment located at the front mount plane. This
effect is
particularly evident on an axial compressor, since the maximum deflection
occurs at the
rear compressor stages, where small tip clearances are needed to maintain
engine
operability. Increased carcass bending results in having to set larger intial
tip
clearances, which results in both loss of efficiency in the compressor and
turbine, and is
therefore critical for the operability of an all-axial compressor.
[0004] Accordingly, there is a need to provide an improved mounting system
for
turbofan gas turbine engines.

-1-

CA 02756914 2011-11-04


SUMMARY
[0005] In one aspect, the described subject matter provides a turbofan gas
turbine
engine comprising: a front mounting structure attached to an annular bypass
duct wall at
a front axial position adjacent an inlet of a bypass air passage, the bypass
air passage
defined radially between the bypass duct wall and a core portion of the
engine, the front
mounting structure supporting the core portion within the bypass duct wall;
and a rear
mounting structure having a mounting apparatus attached to the bypass duct
wall at a
rear axial position adjacent to an outlet of the bypass air passage, the rear
mounting
structure including three pairs of first and second rods each having opposed
inner and
outer ends, the rods extending across the bypass air passage and
interconnecting the
bypass duct wall and an inner case of the core portion, wherein the first rods
extend
from the outer end to the inner end thereof in substantially tangential
directions with
respect to the core portion, corresponding to a first circumferential
direction, and the
second rods extend from the outer end to the inner end thereof in
substantially tangential
directions with respect to the core portion, corresponding to a second
circumferential
direction opposite to the first circumferential direction.
[0006] In another aspect, the described subject matter provides a turbofan gas
turbine engine comprising: a core portion including an inner case; an annular
bypass
duct wall surrounding and supporting the core portion, to thereby define a
bypass air
passage radially between the core portion and the bypass duct for directing a
bypass air
flow passing therethrough; and a rod frame structure including three pairs of
first and
second rods defined in a radial plane defined by the inner case, the rod frame
structure
interconnecting the bypass duct wall and the inner case for transferring core
portion
related inertia-induced loads from the inner case in a short circuit across
the bypass air
passage in the radial plane to the bypass duct wall, thereby reducing
distortion of the
core portion caused by the inertia-induced loads and reducing carcass bending
of the
core portion.
[0007] Further details of these and other aspects of the described subject
matter will
be apparent from the detailed description and drawings included below.



- 2 -

CA 02756914 2011-11-04


BRIEF DESCRIPTION OF THE DRAWINGS
[0008] Reference is now made to the accompanying drawings depicting aspects
of
the described subject matter, in which:
[0009] FIG. 1 is a schematic cross-sectional view of a turbofan gas turbine
engine
as an exemplary application of the describe subject matter;
[0010] FIG. 2A is a perspective view of a rear mounting structure according
to one
embodiment, as used in the engine of FIG. 1;
[0011] FIG. 2B is a schematic cross-sectional view of a rear mounting
structure
having six links in comparison with a structure having eight links;
[0012] FIG. 3 is a partial perspective view of the rear mounting structure of
FIG 2A
in an enlarged portion, showing one of the connecting brackets with a mounting
portion;
[0013] FIG. 4 is a partial perspective view of the circled area 4 of the rear
mounting
structure of FIG. 2A, looking into the inside surface of a bypass duct wall in
an enlarged
scale, showing the attachment of link rods to the connecting brackets;
[0014] FIG. 5 is a cross-sectional view of the link rod taken along line 5-5
in FIG.
4, showing the aerodynamic profile of the link rod;
[0015] FIG. 6 is a partial perspective view (partially exploded) of the rear
mounting
structure of FIG. 2A in an enlarged scale, showing a lockable adjustment
device for
connection of the link rods to a mid turbine frame (MTF) of a core portion of
the
engine;
[0016] FIG. 7A is a top plan view of a pin used in the lockable adjustment
device
of FIG. 6, showing an annular position of an eccentric distance between the
central axes
of the respective connecting section and base section of the pin;
[0017] Figure 7B is a side elevational view of the pin in FIG. 7A with a
connected
inner end of a link rod shown in broken lines;
[0018] FIG. 8 is a perspective view of a lockable adjustment device according
to
another embodiment;
[0019] FIG. 9 is a schematic illustration of a prior art turbofan gas turbine
engine
mounting system, showing a cantilever core portion; and

- 3 -

CA 02756914 2011-11-04


[0020] FIG. 10 is a schematic illustration of a prior art turbofan gas turbine
engine
mounting system, showing a rear core portion mount.
[0021] It will be noted that throughout the appended drawings, like features
are
identified by like reference numerals.
DETAILED DESCRIPTION
[0022] Referring to FIG. 1 a long duct mixed flow (LDMF) turbofan gas turbine
engine (not numbered) includes an annular bypass duct wall 10, a low pressure
spool
assembly (not numbered) which includes a fan assembly 14 and a low pressure
turbine
assembly 18 connected by a shaft 12, and a high pressure spool assembly (not
numbered) which includes a high pressure compressor assembly 22 and a high
pressure
turbine assembly 24 connected by a shaft 20. The shafts 12 and 20 rotate about
a
central axis 11 of the engine. A core portion 13 accommodates the high
pressure
compressor 22 and the low and high pressure turbine assemblies 18, 24, to
define a
main fluid path (not numbered) therethrough. In the main fluid path there is
provided a
combustor 26 to generate combustion gases to power the high and low pressure
turbine
assemblies 24, 18. An inner core case 13 houses the engine's core, and in this
example
includes a mid turbine frame (MTF) 28 disposed between the high and low
pressure
turbine assemblies 24 and 18. The core portion 13 is coaxially positioned
within the
annular bypass duct wall 10 and an annular bypass air passage 30 is defined
radially
between the annular bypass duct wall 10 and the core portion 13 of the engine
for
directing a bypass air flow 32 driven by the fan assembly 14, to pass
therethrough.
[0023] Referring to FIGS. 1-5, a front mounting structure 34 is attached to
the
annular bypass duct wall 10 at a front axial position indicated by line 36
(representing a
front mounting plane) located close to an inlet (not numbered) of the annular
bypass air
passage 30, to mount the engine to an aircraft (not shown). Radial struts 38
are
provided near the axial location of the front mounting plane 36 and extend
between the
bypass duct wall 10 and the core portion 13 to support the core portion within
the
bypass duct 10, transferring thrust, lateral and vertical loads to the front
mounting
structure 34.
[0024] A rear mounting structure 40 is also attached to the annular bypass
duct
wall 10 at a rear axial position indicated by line 42 (representing a
substantially vertical
- 4 -

CA 02756914 2011-11-04


rear mounting plane to which the central axis 11 of the engine is
perpendicular), close to
an outlet (not numbered) of the bypass air passage 30. The rear mounting
structure 40
according to one embodiment includes a plurality of, such as three,
circumferentially
spaced apart connecting brackets 44 which are attached to the bypass duct wall
10, and
a plurality of, such as six link, rods 46 having opposed inner and outer ends
(not
numbered), extending across the annular bypass air passage 30, and
substantially
tangential to the core portion 13 of the engine. Each link rod 46 is connected
at the
outer end thereof to the bypass duct wall 10 by means of connecting brackets
44 and is
attached at the inner end thereof to the core portion 13, in this example at
the MTF 28,
thereby forming a six-link-rod structure. Generally, the six-link-rod
structure is formed
between the bypass duct and an inner case which is part of the core portion
13.
[0025] The six link rods 46 are in three pairs and each pair of link rods 46
includes
a first link rod 46a extending from the outer end to the inner end thereof in
a
substantially tangential direction to the core portion 13 corresponding to a
first
circumferential direction 48a, and a second link rod 46b extending from the
outer end to
the inner end thereof in a substantially tangential direction to the core
portion 13
corresponding to a second circumferential direction 48b opposite to the first
circumferential direction 48a.
[0026] Each of the connecting brackets 44 according to this embodiment, is
connected with two adjacent link rods 46, i.e. one pair including one first
link rod 46a
and one second link rod 46b. In particular, the connecting bracket 44 has a
generally U-
shaped cross-section formed by two spaced apart side walls (not numbered)
interconnected by a bottom wall 50 which is curved to match the configuration
of a
portion of a peripheral surface of the annular bypass duct wall 10. The
connecting
bracket 44 is mounted to the outer side of the bypass duct wall 10, and is
axially
positioned between and affixed to two axially spaced apart flanges 52 which
extend
radially and outwardly from the annular bypass duct wall 10. A cavity 56 with
a closed
top and open bottom is provided at the middle of each of the connecting
brackets 44,
defmed between the axially spaced apart side walls of the connecting brackets
44 and
between two circumferentially spaced apart end walls 58. The two
circumferentially
spaced apart end walls 58 extend divergently from each other, substantially in
the
tangential directions corresponding to one pair of the two adjacent link rods
46 (the first

- 5 -

CA 02756914 2011-11-04


link rod 46a and the second link rod 46b) which are connected to said
connecting
bracket 44. At least one of the connecting brackets 44 includes a mounting
portion 54
with one or more mounting openings (not numbered) defmed therein, extending
radially
and outwardly from the annular bypass duct wall 10 for connection with a
mounting
device of the aircraft (not shown). Alternatively, one or more mounting
portions 54
may be separated from the connecting brackets 44, and mounted to other
circumferential
locations of the bypass duct 10, as illustrated in FIG. 2B.
[0027] A plurality of openings 60 in the annular bypass duct wall 10 are
provided
aligning with the cavities 56 of the respective connecting brackets 44, in
order to allow
the outer end of each link rod 46 to access the cavity 56 in the connecting
bracket 44
mounted to the outside of the bypass duct wall 10, from the inside of the
bypass air
passage 30. The inner ends of the two adjacent link rods 46 are secured to the

circumferentially spaced end walls 58 of each connecting bracket 44 by means
of screw
fasteners (not numbered), respectively.
[0028] Each of the link rods 46 may have an aerodynamic profile in cross-
section
(see FIG. 5), defined with side surfaces 62 extending between a leading edge
64 and a
trailing edge 66 with respect to the bypass air passage 30 of the engine. The
cross-
sectional profile of the link rod 46 may have a dimension "C" between the side
surfaces
62 smaller than a dimension "X" between the leading and trailing edges 64, 66
in order
to reduce air pressure loss in the bypass air flow 32 caused by the link rods
46. A
hollow configuration of the link rod 46 may also be an option.
[0029] The tangential link rods 46 may be connected at their inner ends
directly to
the core 13 or by means of any type of connector assemblies. For example, the
link rods
46 are usually fabricated in a same length for manufacturing economy and
installation
mistake-proofing. Therefore, an additional adjustability feature may be
required to
accommodate the eccentric condition of the bypass duct wall 10 and the core
portion 13
caused by manufacturing and assembly tolerances thereof. Therefore, the
tangential
link rods 46 may be connected to core portion 13 by means of a lockable
adjustment
device 68 in order to maintain the link rod 46 in the correct orientation to
the flow,
which will be further described hereinafter.
[0030] The tangential link rods 46 form a short circuit across the annular
bypass air
passage 30 in a radial plane which is substantially superposed on the vertical
rear
- 6 -

CA 02756914 2011-11-04


mounting plane indicated by line 42. The short circuit transfers the core
portion related
inertia-induced loads from, for example, the MTF 28 to the connecting brackets
44 and
the bypass duct wall 10.
[0031] The link rods 46 function as an effective load path to the rear
mounting
structure 40 for inertia-induced loads originating from the core portion 13,
thus reducing
core deflections from that source (inertia-induced meaning loads from gravity
or
acceleration). The core portion 13 is therefore supported at both mount planes

represented by lines 36, 42, rather than the "cantilever" mount of FIG. 9
which does not
support the core portion 13 at the rear and hence causes core droop effect.
[0032] It should be noted that if only engine thrust is applied to the
structure of an
engine which is of a rear core mount as shown in FIG. 10, the center of the
bypass
would shift laterally from the center of the engine core. This is because the
core is
bending like a simply supported beam and has a certain amount of bending
rotation at
the front mount. This rotation is then carried through to the bypass flange at
the outside
of the intermediate case and gives a slope to the bypass relative to the core,
which in
turn leads to a lateral shifting of bypass center relative to the core center
at the rear
mount. In contrast, the rear mounting structure 40 of this embodiment adds in
the link
rods 46, and moves the rear mount reaction point to the bypass duct wall 10.
This
relative centerline shift associated with the rear core mount of FIG. 10, is
largely
prevented by the tie-up with the link rods 46. The bypass duct wall 10 is a
stiffer load
path than the core portion 13, and thus the bypass duct wall 10 rather than
the core
portion 13, carries the bulk of the moment produced by the rear mount
reaction, thereby
reducing carcass bending of the core portion 13.
[0033] FIG. 2B schematically illustrates a comparison of the rear mounting
structure 40 which as six link rods 46, with an eight-link-rod structure of
another
embodiment, which is shown in broken lines. It is desirable to have all link
rods be
substantially tangential with respect to the core portion 13 in order to
effectively transfer
the core portion related inertia-induced loads without restriction of the core
portion 13
in thermal dilation. In some types of engines, small core portion size
combined with
large bypass ratio would result in link rods of an eight-link-rod structure
being "less
tangential" or "not substantially tangential", in comparison with the six-link-
rod
structure of this embodiment.
- 7 -

CA 02756914 2011-11-04


[0034] The short circuit, according to the embodiment shown in FIG. 2B, is
defmed
by the six link rods 46 which are substantially identical and are disposed to
form a
substantially triangular structure within the bypass duct wall 10, with the
MTF 28
disposed within the triangular structure.
[0035] In comparison with an eight-link-rod structure or a structure having
more
link rods, the six-link-rod structure of this embodiment will provide similar
or better
core tip clearances but at lower weight. The six-link-rod structure allows
link rods to lie
more tangentially to the core portion, thereby reducing thermal stresses. Due
to the
reduced number of link rods, manufacturing costs and bypass blockage/aero
losses are
also reduced. The reduced bypass blockage/aero losses will also improve engine

performance.
[0036] According to an alterative embodiment, the rear mounting structure 40
(shown in broken lines) may be attached to the bypass duct wall 10 at the rear
mount
plane located further aft relative to link rods 46, as indicated by line 42'
in FIG. 1. The
link arrangement in this embodiment has the same effect as in the previously
described
embodiment, but link loads are somewhat reduced because the axial portion of
the
bypass duct wall 10 between the link rods 46 and the rear mount plane 42',
allows for
the rear mount load to "diffuse" before reaching the link rods. In such an
embodiment,
the connecting brackets 44 as shown in FIG. 2A, will not be located at the
apex of a set
of the link rods.
[0037] Referring to FIGS. 1-2A and 5-7B, the lockable adjusting device 68
includes at least one pin 70 and a connecting base 72 to connect at least one
link rod 46
to an inner case such as the MTF 28. In the embodiment shown in FIGS. 2A and
6, two
pins 70 are provided to each connecting base 72 such that each connecting base
72 can
connect two adjacent link rods 46 to the MTF 28 (one rod 46a and the other rod
46b).
For convenience and precision of description, only one pin 70 and its
connection to the
connecting base 72 is described. It should be noted that the other pin 70 and
its
connection to the same connecting base 72 is substantially the same.
[0038] The connecting bases 72 are circumferentially spaced apart and attached
to
the core portion 13, for example to a flange 74 radially and outwardly
extending from
the MTF 28 of the core portion 13. Each of the connecting bases 72 defines two
holes
76 extending substantially radially therethrough. The pin 70 includes a
connecting
- 8 -

CA 02756914 2011-11-04


section 78 with a central axis 80 and a base section 82 with a central axis
84. The
central axis 80 of the connecting section 78 is eccentric to the central axis
84 of the base
section 82, at an eccentric distance "d". The connecting section 78 is
received in a hole
86 of a link rod 46 (FIG. 7B), and the base section 82 is received in one of
the holes 76
defined in the connecting base 72 (FIG. 6) . Therefore, an angular position
"A" of the
eccentric distance d with respect to a direction represented by line 88 which
is parallel
to the connected link rod 46, may be selected by rotating the pin 70 before
the pin 70 is
locked in position to secure the rod 46 to the connecting base 72. When the
angular
position A of the eccentric distance d changes within 180 degrees, a link
length "L"
which is measured in the direction of line 88 (or in the direction of the
connected link
rod 46) will change in a range of d x 2.
[0030] The base section 82 of the pin 70 and the hole 76 defined in the
connecting
base 72, may be tapered complimentarily to each other. The pin 70 may further
have a
threaded section 90 extending from the small end of the tapered base section
82, for
engagement with a locking nut 92 such that the tapered base section 82 of the
pin 70 is
secured within the tapered hole 76 of the connecting base 72 to lock the
selected angular
position of the pin 70 when the locking nut 92 is tightly engaged with the
threaded
section 90. The base section 82 of the pin 70 and the hole 76 of the
connecting base 72
may be tapered in an angle smaller than a self locking tapering angle such
that the
eccentric pin 70 is self-locked with the connecting base 72 against the
rotation resulting
from offset loads (torque) introduced by the link rods 46 even if the locking
nut 92
accidentally loosens from engagement with the threaded section 90.
[0040] The connecting section 78 may further have a threaded end portion (not
numbered) for engagement with a second locking nut 94 with a washer (not
numbered)
to prevent the connected link rod 46 from disconnecting from the connecting
section 78
of the pin 70.
[0041] The pin 70 may further define a hexagonal recess (not numbered) defined
in
the end of the connecting section 78 as a means to rotate and hold the pin to
maintain
the selected angular position of the pin 70 while tightening the nut 92. The
lockable
adjustment device 68 provides a compact configuration to ensure the
concentricity of
the bypass duct wall 10 and the core portion 13. This configuration can be
attached to
an inner case such as the MTF 28 and located outside of the annular bypass air
duct 30.
- 9 -

CA 02756914 2011-11-04


The adjustment of the eccentric pin 70 need not affect the orientation of the
aerodynamic profile of the link rods 46 in the bypass air flow 24. The self-
locking
tapering feature of the eccentric pin 70 provides a level of mistake-proofing
in the field.
Furthermore, there is no need to re-adjust the pins 70 once the engine is
assembled, and
the link rods 46 may be freely removed and re-installed in the field for
maintenance
purposes because the connecting base 72 which receives the respective link
rods 46 is
independently affixed to the MTF flange 74, thereby maintaining the
adjustment.
[0042] FIG. 8 shows a lockable adjustment device 68a according to another
embodiment in which similar components and features are indicated by numerals
similar to those used for the lockable adjustment device 68 of FIG. 6 for ease
of
description. The difference between devices 68 of FIG. 6 and 68a of FIG. 8,
lies in that
the pin 70 of adjustment device 68a further includes an extension 96 extending
from the
connecting section and is concentric with the base section 76. The extension
96 is
received in a hole 97 defined in a supporting member such as a plate 98. After
the pin
70 is locked in its adjusted position in the connecting base 72 and an inner
end of a link
rod 46 is attached to the connecting section 78 of the pin 70 (similar to that
shown in
FIG. 7B), the plate 98 is attached to the extension 96 of the pin 70 by
receiving the
extension 96 to extend through the hole 97 therein. The plate 98 is then
affixed by
fasteners (not shown) to the connecting base 72 or to the MTF 28. The
extension 96
may optionally have a threaded end portion 100 such that the locking nut 94
with a
bushing (not numbered), may be used to further secure the plate 98 to the pin
70. The
lockable adjustment device 68a provides the connecting base 72 and plate 98 as
two
spaced apart support elements flanking the connecting section 78 which
connects the
link rod 46, thereby forming a double-shear version of an adjustable pin
connecting
arrangement, in contrast to the device 68 of FIG. 6 which is a single-shear
version of an
adjustable pin connecting arrangement.
[0043] It should be understood that a support-link lockable adjustment
arrangement
as illustrated by devices 68 or 68a is described as a part of a support link
of a mounting
system for a long duct mixed flow (LDMF) turbofan gas turbine engine in the
above-
described embodiments. However this support-link lockable adjustment
arrangement
may be applicable to support links of other types for interconnecting an
annular outer
case and an annular inner case of a gas turbine engine. This compact cam-type
of

- 10 -

CA 02756914 2011-11-04


support-link lockable adjustment arrangement can be used at either end of the
link in its
attachment to an outer case or an inner case, conveniently located outside of
the annular
bypass air duct. This support-link lockable adjustment arrangement may be used
with
tangential links as described in this application, or with radial support
links. The
eccentric pin may extend either in a substantially radial direction as
described in the
embodiments or may extend in a substantially axial direction.
[0044] The above description is meant to be exemplary only, and one skilled in
the
art will recognize that changes may be made to the embodiments described
without
departing from the scope of the concept disclosed. For example, the short
circuit for
transferring inertia-induced loads directly from an inner case such as the MTF
to the
bypass duct casing may be configured differently from the particular
embodiments
described above and may be applicable to any bypass duct gas turbine engine
different
from the engine as described. The mounting structure incorporated with the
connector
for connecting the link rods to the bypass duct wall may be configured
differently from
the described embodiments of the connecting brackets. Link rods in each pair
may be
connected to the bypass duct wall separately, rather than by one connecting
bracket.
The link rods may be connected to the core portion individually by fasteners.
The
presence of an MTF is not necessary, and any suitable engine inner casing may
be
supported as described. Still other modifications which fall within the scope
of
described concept will be apparent to those skilled in the art, in light of a
review of this
disclosure, and such modifications are intended to fall within the appended
claims.



- 11 -

Representative Drawing
A single figure which represents the drawing illustrating the invention.
Administrative Status

For a clearer understanding of the status of the application/patent presented on this page, the site Disclaimer , as well as the definitions for Patent , Administrative Status , Maintenance Fee  and Payment History  should be consulted.

Administrative Status

Title Date
Forecasted Issue Date 2019-01-22
(22) Filed 2011-11-04
(41) Open to Public Inspection 2013-05-04
Examination Requested 2016-10-27
(45) Issued 2019-01-22

Abandonment History

There is no abandonment history.

Maintenance Fee

Last Payment of $263.14 was received on 2023-10-19


 Upcoming maintenance fee amounts

Description Date Amount
Next Payment if standard fee 2024-11-04 $347.00
Next Payment if small entity fee 2024-11-04 $125.00

Note : If the full payment has not been received on or before the date indicated, a further fee may be required which may be one of the following

  • the reinstatement fee;
  • the late payment fee; or
  • additional fee to reverse deemed expiry.

Patent fees are adjusted on the 1st of January every year. The amounts above are the current amounts if received by December 31 of the current year.
Please refer to the CIPO Patent Fees web page to see all current fee amounts.

Payment History

Fee Type Anniversary Year Due Date Amount Paid Paid Date
Application Fee $400.00 2011-11-04
Maintenance Fee - Application - New Act 2 2013-11-04 $100.00 2013-11-04
Maintenance Fee - Application - New Act 3 2014-11-04 $100.00 2014-10-07
Maintenance Fee - Application - New Act 4 2015-11-04 $100.00 2015-09-29
Maintenance Fee - Application - New Act 5 2016-11-04 $200.00 2016-10-21
Request for Examination $800.00 2016-10-27
Maintenance Fee - Application - New Act 6 2017-11-06 $200.00 2017-10-23
Maintenance Fee - Application - New Act 7 2018-11-05 $200.00 2018-10-24
Final Fee $300.00 2018-12-07
Maintenance Fee - Patent - New Act 8 2019-11-04 $200.00 2019-10-22
Maintenance Fee - Patent - New Act 9 2020-11-04 $200.00 2020-10-21
Maintenance Fee - Patent - New Act 10 2021-11-04 $255.00 2021-10-20
Maintenance Fee - Patent - New Act 11 2022-11-04 $254.49 2022-10-24
Maintenance Fee - Patent - New Act 12 2023-11-06 $263.14 2023-10-19
Owners on Record

Note: Records showing the ownership history in alphabetical order.

Current Owners on Record
PRATT & WHITNEY CANADA CORP.
Past Owners on Record
None
Past Owners that do not appear in the "Owners on Record" listing will appear in other documentation within the application.
Documents

To view selected files, please enter reCAPTCHA code :



To view images, click a link in the Document Description column. To download the documents, select one or more checkboxes in the first column and then click the "Download Selected in PDF format (Zip Archive)" or the "Download Selected as Single PDF" button.

List of published and non-published patent-specific documents on the CPD .

If you have any difficulty accessing content, you can call the Client Service Centre at 1-866-997-1936 or send them an e-mail at CIPO Client Service Centre.


Document
Description 
Date
(yyyy-mm-dd) 
Number of pages   Size of Image (KB) 
Abstract 2011-11-04 1 8
Description 2011-11-04 11 563
Claims 2011-11-04 4 138
Drawings 2011-11-04 9 167
Representative Drawing 2012-09-24 1 19
Cover Page 2013-04-29 1 42
Examiner Requisition 2017-10-17 5 276
Amendment 2018-04-11 6 247
Claims 2018-04-11 3 140
Drawings 2018-04-11 9 170
Final Fee 2018-12-07 2 66
Representative Drawing 2019-01-02 1 15
Cover Page 2019-01-02 1 39
Assignment 2011-11-04 4 165
Request for Examination 2016-10-27 2 70