U.S. patent application number 14/761559 was filed with the patent office on 2015-12-17 for gas turbine engine combustor liner assembly with convergent hyperbolic profile.
This patent application is currently assigned to UNITED TECHNOLOGIES CORPORATION. The applicant listed for this patent is UNITED TECHNOLOGIES CORPORATION. Invention is credited to Frank J. Cunha, Nurhak Erbas-Sen.
Application Number | 20150362192 14/761559 |
Document ID | / |
Family ID | 51209958 |
Filed Date | 2015-12-17 |
United States Patent
Application |
20150362192 |
Kind Code |
A1 |
Cunha; Frank J. ; et
al. |
December 17, 2015 |
GAS TURBINE ENGINE COMBUSTOR LINER ASSEMBLY WITH CONVERGENT
HYPERBOLIC PROFILE
Abstract
A liner assembly for a combustor of a gas turbine engine
according to one disclosed non-limiting embodiment of the present
disclosure includes a support shell with a convex profile which
faces the heat shield. A further embodiment of the foregoing
embodiment of the present disclosure is where the convex profile is
defined by a hyperbolic cosine function. A further embodiment of
any of the foregoing embodiments of the present disclosure is where
the convex profile provides an approximate 4.5 inlet-to-exit area
ratio. A further embodiment of any of the foregoing embodiments, of
the present disclosure wherein the convex profile provides a flow
acceleration toward approximately 0.5 Mach towards an end of a
convergent section.
Inventors: |
Cunha; Frank J.; (Avon,
CT) ; Erbas-Sen; Nurhak; (Manchester, CT) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
UNITED TECHNOLOGIES CORPORATION |
Hartford |
CT |
US |
|
|
Assignee: |
UNITED TECHNOLOGIES
CORPORATION
Hartford
CT
|
Family ID: |
51209958 |
Appl. No.: |
14/761559 |
Filed: |
January 17, 2013 |
PCT Filed: |
January 17, 2013 |
PCT NO: |
PCT/US13/21921 |
371 Date: |
July 16, 2015 |
Current U.S.
Class: |
60/772 ;
60/755 |
Current CPC
Class: |
Y02T 50/60 20130101;
F23R 3/002 20130101; F23R 2900/03043 20130101; F23R 2900/03045
20130101; F23R 3/04 20130101; F23R 3/50 20130101; F23R 2900/03042
20130101 |
International
Class: |
F23R 3/04 20060101
F23R003/04; F23R 3/00 20060101 F23R003/00 |
Claims
1. A liner assembly for a combustor of a gas turbine engine
comprising: a heat shield; and a support shell with a convex
profile which faces said heat shield.
2. The liner assembly as recited in claim 1, wherein said convex
profile is defined by a hyperbolic cosine function.
3. The liner assembly as recited in claim 1, wherein said convex
profile provides an approximate 4.5 inlet-to-exit area ratio.
4. The liner assembly as recited in claim 1, wherein said convex
profile provides a flow acceleration toward approximately 0.5 Mach
towards a end of a convergent section.
5. The liner assembly as recited in claim 1, further comprising an
exit splitter that extends from said heat shield.
6. The liner assembly as recited in claim 5, wherein said exit
splitter is zigzag in shape.
7. The liner assembly as recited in claim 5, further comprising a
film hole located in a valley on each side of said exit
splitter.
8. The liner assembly as recited in claim 1, further comprising a
plurality of studs which extend from said heat shield and are
received through said support shell, said stud include a
frustro-conical section.
9. The liner assembly as recited in claim 1, wherein said heat
shield includes a number of film holes which are approximately
equal to a number of impingement holes through said support
shell.
10. The liner assembly as recited in claim 1, wherein said heat
shield includes a multiple of pin fins.
11. The liner assembly as recited in claim 10, wherein said
multiple of pin fins are diamond-shaped.
12. The liner assembly as recited in claim 1, wherein said heat
shield includes a multiple of hemi-spherical dimples.
13. The liner assembly as recited in claim 12, wherein said
multiple of hemi-spherical dimples decrease in diameter toward an
exit splitter.
14. The liner assembly as recited in claim 12, wherein a center of
said sphere of each of said multiple of hemi-spherical dimples are
further displaced from an inner surface of said heat shield toward
an exit splitter.
15. A liner assembly for a combustor of a gas turbine engine
comprising: a heat shield; and a support shell non-parallel to said
heat shield.
16. The liner assembly as recited in claim 15, wherein said support
shell defines a convex profile defined by a hyperbolic cosine
function.
17. A method of increasing pressure in a liner assembly of a
combustor for a gas turbine engine, comprising: directing an
airflow in a generally circumferential direction along a convergent
flow channel within a cavity between a heat shield and a support
shell.
18. The method as recited in claim 17, further comprising: defining
the convergent flow channel by a hyperbolic cosine function.
19. The method as recited in claim 17, further comprising: defining
the convergent flow channel to provide an approximate 4.5
inlet-to-exit area ratio.
20. The method as recited in claim 17, further comprising:
accelerating the airflow toward approximately 0.5 Mach towards an
end of said convergent section.
Description
BACKGROUND
[0001] The present disclosure relates to a gas turbine engine and,
more particularly, to a combustor section therefor.
[0002] Gas turbine engines, such as those that power modern
commercial and military aircraft, generally include a compressor to
pressurize an airflow, a combustor for burning a hydrocarbon fuel
in the presence of the pressurized air, and a turbine to extract
energy from the resultant combustion gases.
[0003] As engine requirements increase for improved thrust specific
fuel consumption (TSFC), compressor discharge pressure and
temperature along with combustor exit temperatures (CET) may also
increase. As a result, current combustor configurations emissions,
such as NOx, CO, unburned hydrocarbons (UHC), and smoke, may
increase relative to exceedingly stringent emissions standards.
SUMMARY
[0004] A liner assembly for a combustor of a gas turbine engine
according to one disclosed non-limiting embodiment of the present
disclosure includes a support shell with a convex profile which
faces the heat shield.
[0005] A further embodiment of the foregoing embodiment of the
present disclosure wherein the convex profile is defined by a
hyperbolic cosine function.
[0006] A further embodiment of any of the foregoing embodiments, of
the present disclosure wherein the convex profile provides an
approximate 4.5 inlet-to-exit area ratio.
[0007] A further embodiment of any of the foregoing embodiments, of
the present disclosure wherein the convex profile provides a flow
acceleration toward approximately 0.5 Mach towards a end of a
convergent section.
[0008] A further embodiment of any of the foregoing embodiments, of
the present disclosure includes an exit splitter that extends from
the heat shield.
[0009] In the alternative or additionally thereto, the foregoing
embodiment wherein the exit splitter is zigzag in shape.
[0010] In the alternative or additionally thereto, the foregoing
embodiment further comprising a film hole located in a valley on
each side of the exit splitter.
[0011] A further embodiment of any of the foregoing embodiments, of
the present disclosure includes a plurality of studs which extend
from the heat shield and are received through the support shell,
the stud include a frustro-conical section.
[0012] A further embodiment of any of the foregoing embodiments, of
the present disclosure wherein the heat shield includes a number of
film holes which are approximately equal to a number of impingement
holes through the support shell.
[0013] A further embodiment of any of the foregoing embodiments, of
the present disclosure wherein the heat shield includes a multiple
of pin fins.
[0014] In the alternative or additionally thereto, the foregoing
embodiment wherein the multiple of pin fins are diamond-shaped.
[0015] A further embodiment of any of the foregoing embodiments, of
the present disclosure wherein the heat shield includes a multiple
of hemi-spherical dimples.
[0016] In the alternative or additionally thereto, the foregoing
embodiment includes a multiple of hemi-spherical dimples decrease
in diameter toward an exit splitter.
[0017] In the alternative or additionally thereto, the foregoing
embodiment includes a center of the sphere of each of the multiple
of hemi-spherical dimples are further displaced from an inner
surface of the heat shield toward an exit splitter.
[0018] A liner assembly for a combustor of a gas turbine engine
according to one disclosed non-limiting embodiment of the present
disclosure includes a support shell non-parallel to a heat
shield.
[0019] A further embodiment of the foregoing embodiment of the
present disclosure wherein the support shell defines a convex
profile defined by a hyperbolic cosine function.
[0020] A method of increasing pressure in a liner assembly of a
combustor for a gas turbine engine according to one disclosed
non-limiting embodiment of the present disclosure includes
directing an airflow in a generally circumferential direction along
a convergent flow channel within a cavity between a heat shield and
a support shell.
[0021] A further embodiment of the foregoing embodiment of the
present disclosure includes defining the convergent flow channel by
a hyperbolic cosine function.
[0022] A further embodiment of any of the foregoing embodiments, of
the present disclosure includes defining the convergent flow
channel to provide an approximate 4.5 inlet-to-exit area ratio.
[0023] A further embodiment of any of the foregoing embodiments, of
the present disclosure includes accelerating the airflow toward
approximately 0.5 Mach towards an end of the convergent
section.
BRIEF DESCRIPTION OF THE DRAWINGS
[0024] Various features will become apparent to those skilled in
the art from the following detailed description of the disclosed
non-limiting embodiment. The drawings that accompany the detailed
description can be briefly described as follows:
[0025] FIG. 1 is a schematic cross-section of a gas turbine
engine;
[0026] FIG. 2 is an expanded longitudinal schematic sectional view
of a combustor section according to one non-limiting embodiment
that may be used with the gas turbine engine shown in FIG. 1;
[0027] FIG. 3 is an expanded partial perspective longitudinal
schematic view of a combustor section according to one non-limiting
embodiment that may be used with the gas turbine engine shown in
FIG. 1;
[0028] FIG. 4 is an exploded view of a liner assembly of the
combustor;
[0029] FIG. 5 is an expanded circumferentially partial perspective
view of the combustor section associates with one pre-swirler;
[0030] FIG. 6 is an expanded lateral sectional view of a liner
assembly according to one non-limiting embodiment;
[0031] FIG. 7 is an expanded lateral sectional view of the liner
assembly of FIG. 6 with a relationship for a convex profile that
faces an inner surface of a heat shield of the liner assembly;
[0032] FIG. 8 is an expanded plan view of a heat shield of a liner
assembly according to one non-limiting embodiment;
[0033] FIG. 9 is an expanded plan view of a heat shield of a liner
assembly according to another non-limiting embodiment;
[0034] FIG. 10 is an expanded lateral sectional view of two
adjacent liner assemblies;
[0035] FIG. 11 is an expanded perspective view of an overlapping
interface between two adjacent liner assemblies;
[0036] FIG. 12 is an expanded lateral sectional view of two
adjacent liner assemblies;
[0037] FIG. 13 is a forward view of two adjacent combustor sections
facing a bulkhead heat shield illustrating cooling flow according
to one non-limiting embodiment;
[0038] FIG. 14 is a forward view of a combustor section facing a
bulkhead heat shield illustrating cooling flow according to another
non-limiting embodiment;
[0039] FIG. 15 is a forward view of a combustor section facing a
bulkhead heat shield illustrating cooling flow according to another
non-limiting embodiment; and
[0040] FIG. 16 is an expanded lateral sectional view of two
adjacent liner assemblies.
DETAILED DESCRIPTION
[0041] FIG. 1 schematically illustrates a gas turbine engine 20.
The gas turbine engine 20 is disclosed herein as a two-spool turbo
fan 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 while the compressor section 24 drives air
along a core flowpath for compression and communication into the
combustor section 26 then expansion through the turbine section 28.
Although depicted as a turbofan 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 such as a turbojets,
turboshafts, and three-spool (plus fan) turbofans wherein an
intermediate spool includes an intermediate pressure compressor
("IPC") between a Low Pressure Compressor ("LPC") and a High
Pressure Compressor ("HPC"), and an intermediate pressure turbine
("IPT") between the high pressure turbine ("HPT") and the Low
pressure Turbine ("LPT").
[0042] The engine 20 generally includes a low spool 30 and a high
spool 32 mounted for rotation about an engine central longitudinal
axis A relative to an engine static structure 36 via several
bearing structures 38. The low spool 30 generally includes an inner
shaft 40 that interconnects a fan 42, a low pressure compressor 44
("LPC") and a low pressure turbine 46 ("LPT"). The inner shaft 40
drives the fan 42 directly or through a geared architecture 48 to
drive the fan 42 at a lower speed than the low spool 30. An
exemplary reduction transmission is an epicyclic transmission,
namely a planetary or star gear system.
[0043] The high spool 32 includes an outer shaft 50 that
interconnects a high pressure compressor 52 ("HPC") and high
pressure turbine 54 ("HPT"). A combustor 56 is arranged between the
high pressure compressor 52 and the high pressure turbine 54. The
inner shaft 40 and the outer shaft 50 are concentric and rotate
about the engine central longitudinal axis A which is collinear
with their longitudinal axes.
[0044] Core airflow is compressed by the LPC 44 then the HPC 52,
mixed with the fuel and burned in the combustor 56, then expanded
over the HPT 54 and the LPT 46. The turbines 54, 46 rotationally
drive the respective low spool 30 and high spool 32 in response to
the expansion. The main engine shafts 40, 50 are supported at a
plurality of points by bearing structures 38 within the static
structure 36. It should be understood that various bearing
structures 38 at various locations may alternatively or
additionally be provided.
[0045] In one non-limiting example, the gas turbine engine 20 is a
high-bypass geared aircraft engine. In a further example, the gas
turbine engine 20 bypass ratio is greater than about six (6:1). The
geared architecture 48 can include an epicyclic gear train, such as
a planetary gear system or other gear system. The example epicyclic
gear train has a gear reduction ratio of greater than about 2.3,
and in another example is greater than about 2.5:1. The geared
turbofan enables operation of the low spool 30 at higher speeds
which can increase the operational efficiency of the low pressure
compressor 44 and low pressure turbine 46 and render increased
pressure in a fewer number of stages.
[0046] A pressure ratio associated with the low pressure turbine 46
is pressure measured prior to the inlet of the low pressure turbine
46 as related to the pressure at the outlet of the low pressure
turbine 46 prior to an exhaust nozzle of the gas turbine engine 20.
In one non-limiting embodiment, the bypass ratio of the gas turbine
engine 20 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). 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.
[0047] In one embodiment, a significant amount of thrust is
provided by the bypass flow path due to the high bypass ratio. The
fan section 22 of the gas turbine engine 20 is designed for a
particular flight condition--typically cruise at about 0.8 Mach and
about 35,000 feet. This flight condition, with the gas turbine
engine 20 at its best fuel consumption, is also known as bucket
cruise Thrust Specific Fuel Consumption (TSFC). TSFC is an industry
standard parameter of fuel consumption per unit of thrust.
[0048] Fan Pressure Ratio is the pressure ratio across a blade of
the fan section 22 without the use of a Fan Exit Guide Vane system.
The low Fan Pressure Ratio according to one non-limiting embodiment
of the example gas turbine engine 20 is less than 1.45. Low
Corrected Fan Tip Speed is the actual fan tip speed divided by an
industry standard temperature correction of ("T"/518.7).sup.0.5 in
which "T" represents the ambient temperature in degrees Rankine.
The Low Corrected Fan Tip Speed according to one non-limiting
embodiment of the example gas turbine engine 20 is less than about
1150 fps (351 m/s).
[0049] With reference to FIG. 2, the combustor 56 generally
includes an outer combustor liner assembly 60, an inner combustor
liner assembly 62 and a diffuser case module 64. The outer
combustor liner assembly 60 and the inner combustor liner assembly
62 are spaced apart such that a combustion chamber 66 is defined
therebetween. The combustion chamber 66 is generally annular in
shape.
[0050] The outer combustor liner assembly 60 is spaced radially
inward from an outer diffuser case 64-O of the diffuser case module
64 to define an outer annular plenum 76. The inner combustor liner
assembly 62 is spaced radially outward from an inner diffuser case
64-I of the diffuser case module 64 to define an inner annular
plenum 78. It should be understood that although a particular
combustor is illustrated, other combustor types with various
combustor liner arrangements will also benefit herefrom. It should
be further understood that the disclosed cooling flow paths are but
an illustrated embodiment and should not be limited only
thereto.
[0051] The combustor liner assemblies 60, 62 contain the combustion
products for direction toward the turbine section 28. Each
combustor liner assembly 60, 62 generally includes a respective
support shell 68, 70 which supports one or more heat shields 72, 74
mounted to a hot side of the respective support shell 68, 70. Each
of the heat shields 72, 74 may be generally rectilinear and
manufactured of for example, a nickel based super alloy, ceramic or
other temperature resistant material and are arranged to form a
liner array. In one disclosed non-limiting embodiment, the liner
array includes a multiple of forward heat shields 72A and a
multiple of aft heat shields 72B that are circumferentially
staggered to line the hot side of the outer shell 68 (also shown in
FIG. 3). A multiple of forward heat shields 74A and a multiple of
aft heat shields 74B are circumferentially staggered to line the
hot side of the inner shell 70 (also shown in FIG. 3).
[0052] The combustor 56 further includes a forward assembly 80
immediately downstream of the compressor section 24 to receive
compressed airflow therefrom. The forward assembly 80 generally
includes an annular hood 82, a bulkhead assembly 84, a multiple of
fuel nozzles 86 (one shown) and a multiple of fuel nozzle
pre-swirlers 90 (one shown). Each of the fuel nozzle pre-swirlers
90 is circumferentially aligned with one of the hood ports 94 to
project through the bulkhead assembly 84. Each bulkhead assembly 84
includes a bulkhead support shell 96 secured to the combustor liner
assemblies 60, 62, and a multiple of circumferentially distributed
bulkhead heat shields 98 secured to the bulkhead support shell 96
around the central opening 92.
[0053] The annular hood 82 extends radially between, and is secured
to, the forwardmost ends of the combustor liner assemblies 60, 62.
The annular hood 82 includes a multiple of circumferentially
distributed hood ports 94 that accommodate the respective fuel
nozzle 86 and introduce air into the forward end of the combustion
chamber 66 through a central opening 92. Each fuel nozzle 86 may be
secured to the diffuser case module 64 and project through one of
the hood ports 94 and through the central opening 92 within the
respective fuel nozzle guide 90.
[0054] The forward assembly 80 introduces core combustion air into
the forward section of the combustion chamber 66 while the
remainder enters the outer annular plenum 76 and the inner annular
plenum 78. The multiple of fuel nozzles 86 and adjacent structure
generate a blended fuel-air mixture that supports stable combustion
in the combustion chamber 66.
[0055] Opposite the forward assembly 80, the outer and inner
support shells 68, 70 are mounted to a first row of Nozzle Guide
Vanes (NGVs) 54A in the HPT 54. The NGVs 54A are static engine
components which direct core airflow combustion gases onto the
turbine blades of the first turbine rotor in the turbine section 28
to facilitate the conversion of pressure energy into kinetic
energy. The core airflow combustion gases are also accelerated by
the NGVs 54A because of their convergent shape and are typically
given a "spin" or a "swirl" in the direction of turbine rotor
rotation. The turbine rotor blades absorb this energy to drive the
turbine rotor at high speed.
[0056] With reference to FIG. 4, a multiple of studs 100 extend
from the heat shields 72, 74 to mount the heat shields 72, 74 to
the respective support shells 68, 70 with fasteners 102 such as
nuts (also shown in FIG. 3). That is, the studs 100 project rigidly
from the heat shields 72, 74 and through the respective support
shells 68, 70 to receive the fasteners 102 at a threaded distal end
section thereof.
[0057] A multiple of cooling impingement holes 104 penetrate
through the support shells 68, 70 to allow air from the respective
annular plenums 76, 78 to enter cavities 106A, 106B (also shown in
FIG. 3) formed in the combustor liner assemblies 60, 62 between the
respective support shells 68, 70 and heat shields 72, 74. The
cooling impingement holes 104 are generally normal to the surface
of the heat shields 72, 74. The air in the cavities 106A, 106B
provides backside impingement cooling of the heat shields 72, 74
that is generally defined herein as heat removal via internal
convection.
[0058] A multiple of cooling film holes 108 penetrate through each
of the heat shields 72, 74. The geometry of the film holes, e.g.,
diameter, shape, density, surface angle, incidence angle, etc., as
well as the location of the holes with respect to the high
temperature main flow also contributes to effusion film cooling.
The combination of impingement holes 104 and film holes 108 may be
referred to as an Impingement Film Floatliner assembly.
[0059] The cooling film holes 108 allow the air to pass from the
cavities 106A, 106B defined in part by a cold side 110 of the heat
shields 72, 74 to a hot side 112 of the heat shields 72, 74 and
thereby facilitate the formation of a film of cooling air along the
hot side 112. The cooling film holes 108 are generally more
numerous than the impingement holes 104 to promote the development
of a film cooling along the hot side 112 to sheath the heat shields
72, 74. Film cooling as defined herein is the introduction of a
relatively cooler airflow at one or more discrete locations along a
surface exposed to a high temperature environment to protect that
surface in the immediate region of the airflow injection as well as
downstream thereof.
[0060] A multiple of dilution holes 116 penetrate through both the
respective support shells 68, 70 and heat shields 72, 74 along a
common axis D (FIG. 5). For example only, in a Rich-Quench-Lean
(R-Q-L) type combustor, the dilution holes 116 are located
downstream of the forward assembly 80 to quench the hot gases by
supplying cooling air into the combustor. The hot combustion gases
slow towards the dilution holes 116 and may form a stagnation point
at the leading edge which becomes a heat source. At the trailing
edge of the dilution hole, due to interaction with dilution jet,
hot gases form a standing vortex pair that may also become a heat
source.
[0061] With reference to FIG. 6, a lateral cross-section of the
support shells 68, 70 and heat shields 72, 74 with their respective
cavities 106A, 106B are illustrated with respect to the combustion
chamber 66. Although only one of the support shells 68, 70 and heat
shields 72, 74 is illustrated and described in detail hereafter, it
should be understood that each of the support shells 68, 70 and
heat shields 72, 74 are generally the same and need not be
described in detail herein.
[0062] An inner surface 120 of each support shell 68, 70 defines a
convex profile 122 such as a hyperbolic or catenary profile that
faces an inner surface 124 of the heat shields 72, 74 within the
respective cavities 106A, 106B. That is, the convex profile 122
results in the support shell 68, 70 being non-parallel to the
respective heat shields 72, 74. The inner surface 120 of each
support shell 68, 70 defines a relatively thin cavity zone 126
along a central portion of each combustor section 130 with respect
to the inner surface 124 of the heat shields 72, 74. That is, the
relatively thin cavity zone 126 is defined generally parallel to
the engine axis A and is flanked by relatively thicker cavity zones
128 of each combustor section 130 (FIG. 5).
[0063] With Reference to FIG. 7, the convex profile 122 may be
defined by a hyperbolic cosine function, cosh, provides an
approximate 4.5 inlet-to-exit area ratio. The inlet-to-exit area
ratio forces a flow acceleration toward approximately 0.5 Mach at
an end of a circumferential convergent flow section. A
corresponding increase in Reynolds number facilitates higher
internal heat transfer coefficients for cooling.
[0064] With reference to FIG. 8, the relatively thicker cavity
zones 128 receive airflow from the impingement holes 104. The
airflow within the cavities 106A, 106B is from the relatively
thicker cavity zones 128 toward the relatively thin cavity zone 126
to define the circumferential convergent flow section. That is, the
airflow is generally in the circumferential direction rather than
the axial direction.
[0065] In one disclosed non-limiting embodiment, the impingement
holes 104 direct airflow onto a multiple of pin fins 132. The pin
fins 132 in one example, may be diamond shaped pins that are
approximately 1/2-1/4 the height between the inner surfaces 120,
124 in the relatively thicker cavity zones 128. It should be
appreciated that other heights may be provided.
[0066] Inboard of the multiple of pin fins 132, a multiple of
hemispherical dimples 134 are located toward an exit splitter 136.
In one disclosed non-limiting embodiment, the hemispherical dimples
134 are of the same diameter but are progressively deeper into the
inner surface 124. That is, centers of the respective spheres which
define the hemispherical dimples 134 are progressively deeper into
the combustion chamber 66. In another disclosed non-limiting
embodiment, the hemispherical dimples 134-1 are progressively
smaller diameters toward the exit splitter 136 (FIG. 9). The
hemispherical dimples 134, 134-1 allow for less pressure resistance
(less friction) that facilitates convergent flow channel
acceleration capabilities. The hemispherical dimples 134, 134-1
reduce the frictional drag resistance to the cooling flow yet
augment cooling of the inner surface 124. It should be appreciated
that the hemispherical dimples 134, 134-1 may be arranged in
various patterns.
[0067] The exit splitter 136 is zigzag in shape along the axis A. A
film hole 108 is located in a valley 138 on each side of the zigzag
exit splitter 136. As defined herein "zigzag" includes, but is not
limited to, any serpentine, saw tooth or non-straight wall.
[0068] The exit splitter 136 also forms a base for a
frustro-conical stud 100 (only one shown). The stud 100 is received
within a corresponding aperture 140 in the heat shields 72, 74,
such that as the fastener 102 is tightened down on a threaded
interface 142, the aperture 140 seals and tightens onto the
frustro-conical stud portion 144 (FIG. 6).
[0069] With reference to FIG. 10 the threaded interface 142 also
forces sets of interleaved hooks 146, 148 along each edge 150, 152
of the heat shields 72, 74 to be forced together to facilitate a
seal between each adjacent combustor section 130-1, 130-2 (FIG.
11). It should be appreciated that the frustro-conical studs 100
may alternatively or additionally located in other locations such
as along the edges 150, 152. The interleaved hooks 146, 148 react
the force applied to the frustro-conical stud 100 to minimize
leakage.
[0070] With reference to FIG. 12, the film holes 108 along edge 150
of one combustor section 130-1 are directed toward edge 152 of the
adjacent combustor section 130-2 and vice-versa. The cross-flow
from the film holes 108 along edges 150, 152 protect the edges 150,
152 and further facilitates a seal between the interleaved hooks
146, 148.
[0071] The frustro-conical stud 100 and interleaved hooks 146, 148
facilitate a relatively higher pressure within the cavities 106A,
106B. In one example, an equal number of impingement holes 104 and
film holes 108 are located in each combustor section 130 to provide
a approximately 50:50 pressure split as compared to a more
conventional 80:20 pressure split with approximately half the
number of impingement holes 104 compared to the film holes 108. The
50:50 pressure split permits a relatively higher pressure within
the cavities 106A, 106B thereby permitting a relatively smaller
number of holes and thereby a more efficient usage of air by
spacing impingement holes 104 and film holes 108 further apart.
Reduced reaction flame temperatures are also avoid local
stoichiometric conditions and thereby reduce NOx formation
[0072] With reference to FIG. 13, the film holes 108 adjacent to
the zigzag exit splitter 136 may be directed across an interface
154 between circumferentially distributed bulkhead heat shields 98.
That is, the film holes 108 along one side of the exit splitter 136
are directed toward the opposite side and vice-versa. Such an
arrangement may be advantageous when the fuel nozzle pre-swirlers
90 are axially displaced from the film holes 108.
[0073] With reference to FIG. 14, in another disclosed non-limiting
embodiment, the film holes 108 on both sides of the zigzag exit
splitter 136 through the heat shields 72, 74 are directed in a
direction in coordination with the rotational direction of the fuel
nozzle pre-swirlers 90. Such an arrangement may be advantageous
when the fuel nozzle pre-swirlers 90 are positioned relatively
close to the film holes 108. It should be appreciated that the
rotational direction may be clockwise or counter-clockwise.
[0074] With reference to FIG. 15, in another disclosed non-limiting
embodiment, the film holes 108 on both sides of the zigzag exit
splitter 136 through the heat shields 72, 74 are directed in a
direction opposite the rotational direction of the fuel nozzle
pre-swirlers 90.
[0075] With reference to FIG. 16, the convex profile 122 may
include pre-drilled apertures 156 located in potential hot spots.
These apertures 156 are not initially drilled completely through
the support shell 68, 70. That is, the pre-drilled aperture 156 are
placed in the convergent section close to an area where hot-spots
may occur. Should the hot-spot prediction be realized, then
apertures 156 are drilled completely through the support shell 68,
70 to supply refresher air into the convergent section pre-drilled
apertures 156. This will effectively address the hot-spot by
maintaining the coolant heat pick-up low; while introducing more
convective flow into the circuit. Furthermore, even if not drilled
completely through, the pre-drilled apertures 156 provide weight
reduction.
[0076] It should be understood that relative positional terms such
as "forward," "aft," "upper," "lower," "above," "below," and the
like are with reference to the normal operational attitude of the
vehicle and should not be considered otherwise limiting.
[0077] Although the different non-limiting embodiments have
specific illustrated components, the embodiments of this invention
are not limited to those particular combinations. It is possible to
use some of the components or features from any of the non-limiting
embodiments in combination with features or components from any of
the other non-limiting embodiments.
[0078] It should be understood that like reference numerals
identify corresponding or similar elements throughout the several
drawings. It should also be understood that although a particular
component arrangement is disclosed in the illustrated embodiment,
other arrangements will benefit herefrom.
[0079] Although particular step sequences are shown, described, and
claimed, it should be understood that steps may be performed in any
order, separated or combined unless otherwise indicated and will
still benefit from the present disclosure.
[0080] The foregoing description is exemplary rather than defined
by the limitations within Various non-limiting embodiments are
disclosed herein, however, one of ordinary skill in the art would
recognize that various modifications and variations in light of the
above teachings will fall within the scope of the appended claims.
It is therefore to be understood that within the scope of the
appended claims, the disclosure may be practiced other than as
specifically described. For that reason the appended claims should
be studied to determine true scope and content.
* * * * *