U.S. patent application number 16/260543 was filed with the patent office on 2020-01-23 for gas turbine engine combustor with tailored temperature profile.
The applicant listed for this patent is United Technologies Corporation. Invention is credited to James B. Hoke, Randal G. McKinney, Timothy S. Snyder.
Application Number | 20200025376 16/260543 |
Document ID | / |
Family ID | 50979266 |
Filed Date | 2020-01-23 |
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United States Patent
Application |
20200025376 |
Kind Code |
A1 |
Snyder; Timothy S. ; et
al. |
January 23, 2020 |
GAS TURBINE ENGINE COMBUSTOR WITH TAILORED TEMPERATURE PROFILE
Abstract
A method of tailoring a combustor flow for a gas turbine engine
includes controlling an airflow into a swirler to be generally
uniform and controlling an airflow into a quench zone to provide a
desired pattern factor.
Inventors: |
Snyder; Timothy S.;
(Glastonbury, CT) ; McKinney; Randal G.; (Tolland,
CT) ; Hoke; James B.; (Tolland, CT) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
United Technologies Corporation |
Farmington |
CT |
US |
|
|
Family ID: |
50979266 |
Appl. No.: |
16/260543 |
Filed: |
January 29, 2019 |
Related U.S. Patent Documents
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Application
Number |
Filing Date |
Patent Number |
|
|
13725137 |
Dec 21, 2012 |
10260748 |
|
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16260543 |
|
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Current U.S.
Class: |
1/1 |
Current CPC
Class: |
F23R 3/26 20130101; F23R
3/002 20130101; F23R 3/286 20130101; Y02T 50/60 20130101; F23R 3/12
20130101; Y02T 50/675 20130101; F02C 7/2365 20130101; F23R 3/06
20130101 |
International
Class: |
F23R 3/00 20060101
F23R003/00; F23R 3/26 20060101 F23R003/26; F23R 3/12 20060101
F23R003/12; F02C 7/236 20060101 F02C007/236; F23R 3/06 20060101
F23R003/06; F23R 3/28 20060101 F23R003/28 |
Claims
1. A method of tailoring a combustor flow for a gas turbine engine,
comprising: controlling an airflow into a swirler; and controlling
an airflow into a quench zone to provide a desired pattern
factor.
2. The method as recited in claim 1, further comprising:
controlling the airflow into the swirler to be generally
uniform.
3. The method as recited in claim 1, further comprising:
controlling a fuel-air mixture to be generally uniform downstream
of the swirler.
4. The method as recited in claim 1, further comprising:
controlling the airflow into the swirler to increase fuel
uniformity.
5. The method as recited in claim 1, further comprising: trimming
the airflow into the quench zone to achieve a desired radial
profile.
6. The method as recited in claim 5, further comprising:
controlling the airflow into the quench zone to reduce said desired
radial profile in a 5%-25% span region.
7. The method as recited in claim 1, further comprising:
controlling the airflow into the quench zone with a multiple of
trim holes downstream of a quench hole.
8. A method of tailoring a combustor flow for a gas turbine engine,
comprising: controlling interaction of a fuel-air mixture from a
swirler with an airflow into a quench zone.
9. The method as recited in claim 8, further comprising:
controlling the swirler interaction with a quench zone to increase
fuel uniformity.
10. The method as recited in claim 8, further comprising: uniformly
distributing airflow to the swirler.
11. The method as recited in claim 8, further comprising: trimming
an airflow into the quench zone.
12. The method as recited in claim 8, further comprising: providing
a pattern factor below approximately 0.3.
13. The method as recited in claim 8, further comprising: providing
a profile factor below approximately 0.12.
Description
[0001] This application is a divisional of U.S. patent application
Ser. No. 13/725,137 filed on Dec. 21, 2012, which is hereby
incorporated herein by reference in its entirety
BACKGROUND
[0002] The present disclosure relates to a gas turbine engine and,
more particularly, to a combustor section therefor.
[0003] Gas turbine engines, such as those that power modern
commercial and military aircraft, generally include a compressor
section to pressurize an airflow, a combustor section for burning a
hydrocarbon fuel in the presence of the pressurized air, and a
turbine section to extract energy from the resultant combustion
gases.
[0004] High fuel-air ratio combustors generate high thermal loads
for prolonged time periods that may challenge turbine section
durability. The thermal loads may also be non-uniform in some
locations such that the turbine section service life and/or thrust
generation may be at least partially compromised through combustor
section restrictions.
SUMMARY
[0005] A method of tailoring a combustor flow for a gas turbine
engine according to one disclosed non-limiting embodiment of the
present disclosure includes controlling an airflow into a swirler,
and controlling an airflow into a quench zone to provide a desired
pattern factor.
[0006] In a further embodiment of the foregoing embodiment, the
method includes controlling the airflow into the swirler to be
generally uniform.
[0007] In a further embodiment of any of the foregoing embodiments,
the method includes controlling a fuel-air mixture to be generally
uniform downstream of the swirler.
[0008] In a further embodiment of any of the foregoing embodiments,
the method includes controlling the airflow into the swirler to
increase fuel uniformity.
[0009] In a further embodiment of any of the foregoing embodiments,
the method includes trimming the airflow into the quench zone to
achieve a desired radial profile. In the alternative or
additionally thereto, the foregoing embodiment includes controlling
the airflow into the quench zone to reduce said desired radial
profile in a 5%-25% span region.
[0010] In a further embodiment of any of the foregoing embodiments,
the method includes controlling the airflow into the quench zone
with a multiple of trim holes downstream of a quench hole.
[0011] A method of tailoring a combustor flow for a gas turbine
engine, according to another disclosed non-limiting embodiment of
the present disclosure includes controlling interaction of a
fuel-air mixture from a swirler with an airflow into a quench
zone.
[0012] In a further embodiment of the foregoing embodiment, the
method includes controlling the swirler interaction with a quench
zone to increase fuel uniformity.
[0013] In a further embodiment of any of the foregoing embodiments,
the method includes uniformly distributing airflow to the
swirler.
[0014] In a further embodiment of any of the foregoing embodiments,
the method includes trimming an airflow into the quench zone.
[0015] In a further embodiment of any of the foregoing embodiments,
the method includes providing a pattern factor below approximately
0.3.
[0016] In a further embodiment of any of the foregoing embodiments,
the method includes providing a profile factor below approximately
0.12.
[0017] A combustor of a gas turbine engine according to another
disclosed non-limiting embodiment of the present disclosure
includes a liner assembly with a quench hole and a multiple of trim
holes downstream thereof.
[0018] In a further embodiment of the foregoing embodiment, the
multiple of trim holes are circumferentially offset from said
quench hole.
[0019] In a further embodiment of any of the foregoing embodiments,
the multiple of trim holes are axially aligned with a quench hole
in an opposed liner assembly. In the alternative or additionally
thereto, in the foregoing embodiment the opposed liner assembly is
an outer liner assembly. In the alternative or additionally
thereto, in the foregoing embodiment the multiple of trim holes
includes three trim holes.
[0020] In a further embodiment of any of the foregoing embodiments,
the combustor includes a bulkhead assembly transverse to said liner
assembly, a swirler mounted to said bulkhead assembly said swirler
without tabs.
[0021] In a further embodiment of any of the foregoing embodiments,
the combustor includes a bulkhead assembly transverse to said liner
assembly, a swirler mounted to said bulkhead assembly said swirler
circular in shape.
BRIEF DESCRIPTION OF THE DRAWINGS
[0022] 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:
[0023] FIG. 1 is a schematic cross-section of an example gas
turbine engine architecture;
[0024] FIG. 2 is a schematic cross-section of another example gas
turbine engine architecture;
[0025] FIG. 3 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 FIGS. 1 and
2;
[0026] FIG. 4 is an expanded perspective partial sectional view of
a combustor sector;
[0027] FIG. 5 is an exploded view of a swirler;
[0028] FIG. 6 is a cross-sectional view of the swirler of FIG.
5;
[0029] FIG. 7 is a plot of swirler passage flow between the swirler
of FIG. 5 and a conventional tanged swirler;
[0030] FIG. 8 is a rear perspective view of the swirler of FIG.
5;
[0031] FIG. 9 is a rear perspective view of a conventional RELATED
ART tanged swirler;
[0032] FIG. 10 is a schematic flow diagram of an airflow into the
swirler of FIG. 5 along line A-A in FIG. 7;
[0033] FIG. 11 is a schematic flow diagram of an airflow into the
swirler of FIG. 5 along line B-B in FIG. 7;
[0034] FIG. 12 is a schematic flow spectrograph of a fuel-air
mixture from the swirler of FIG. 5;
[0035] FIG. 13 is a schematic flow spectrograph of a fuel-air
mixture from the RELATED ART swirler of FIG. 9;
[0036] FIG. 14 is an expanded perspective partial sectional view of
a liner assembly;
[0037] FIG. 15 is an expanded perspective partial sectional view of
a combustor sector with trim holes according to one non-limiting
embodiment;
[0038] FIG. 16 is a thermograph of a combustor sector adjacent to
the trim holes;
[0039] FIG. 17 is a graphical representation of a pattern factor
for a combustor according the disclosure; and
[0040] FIG. 18 is a graphical representation of a radial profile
factor for a combustor according the disclosure.
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 engine architectures 200 might include an augmentor
section 202 and exhaust duct section 204 (FIG. 2) 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 low bypass augmented
turbofan (FIG. 2), 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.05 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. 3, 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 may be 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. 4). 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. 4). It should be
appreciated that other combustor liner assemblies 60, 62 such as
single wall liners will also benefit herefrom.
[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 swirlers 90 (one
shown). Each of the swirlers 90 is circumferentially aligned with
one of the annular hood ports 94 to project through the bulkhead
assembly 84. Each bulkhead assembly 84 generally includes a
bulkhead support shell 96 secured to the combustor liner assembly
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 swirler 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 the 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 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. 5, each swirler 90 generally includes
a capture plate 100, a nozzle guide 102, a guide housing 104, a
first swirler body 106 and a second swirler body 108. The capture
plate 100 is mounted to the guide housing 104 to retain the nozzle
guide 102, the nozzle guide 102 movable with respect to the guide
housing 104. It should be appreciated that alternative or
additional components may be utilized herewith and that the two
part swirler body shown is merely but one example assembly.
[0057] With reference to FIG. 6, the first swirler body 106
generally includes a base section 110 and a frustroconical section
112 that extends downstream of the base section 110. The base
section 110 includes a multiple of passages 114 generally radial
with respect to the swirler centerline F to receive and direct
primary combustion core airflow from within the hood 82 to generate
a fuel-air mixture. It should be appreciated that generally radial
as defined herein is transverse to said centerline F but inclusive
of a tangential component to impart a swirl to the primary
combustion core airflow about the centerline F.
[0058] An outer periphery of the swirler 90 is circular so as to
provide a uniform periphery and thereby a uniform airflow into the
swirler 90. The swirler may be brazed or otherwise attached to the
bulkhead support shell 96 to maintain the uniform periphery.
[0059] With reference to FIG. 7, the uniform airflow into the
swirler 90 (FIG. 8) facilitates fuel-air uniformity in contrast to
the fuel-air mixture of a conventional non-circular swirler with,
for example, with tangs (FIG. 9; RELATED ART). That is, the airflow
from within the annular hood 82 enters normal to the swirler
passages generally transverse to the fuel nozzle 86 (FIG. 10) and
enters the swirler 90 at a shallow angle generally parallel to the
fuel nozzle 86 (FIG. 11) to provide a uniform fuel-air mixture
(FIG. 12) as compared to a tanged swirler (FIG. 13; RELATED
ART).
[0060] With reference to FIG. 14, a multiple of cooling impingement
holes 116 penetrate through the support shells 68, 70 to allow air
from the respective annular plenums 76, 78 to enter cavities 118
formed in the combustor liner assemblies 60, 62 between the
respective support shells 68, 70 and heat shields 72, 74. The
cooling impingement holes 116 may be generally normal to the
surface of the heat shields 72, 74. The air in the cavities 118
provide backside impingement cooling of the heat shields 72, 74
that is generally defined herein as heat removal via internal
convection.
[0061] A multiple of cooling film holes 120 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 effusion film cooling. The
cooling film holes 120 are generally more numerous than the
impingement holes 116 to promote the development of a film cooling
along a hot side of 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. The
combination of impingement holes 116 and film holes 120 may be
referred to as an Impingement Film Floatliner assembly.
[0062] With reference to FIG. 15, a quench hole 122, 124 penetrates
the respective outer combustor liner assembly 60 and the inner
combustor liner assembly 62 in each segment. For example only, in a
Rich-Quench-Lean (R-Q-L) type combustor, the quench holes 122, 124
are located downstream of a respective fuel nozzle 86 to quench the
hot combustion gases through the supply of cooling air into the
combustor chamber 66. The quench hole 122 through the outer
combustor liner assembly 60 and the quench hole 124 through the
inner combustor liner assembly 62 are circumferentially offset but
axially opposed.
[0063] A multiple of trim holes 126 are located through the inner
combustor liner assembly 62 axially downstream of the quench hole
122. That is, the multiple of trim holes 126 are circumferentially
offset from the quench hole 124 in the inner combustor liner
assembly 62. The area of the multiple of trim holes 126 are
subtracted in proportion to the quench holes 122, 124 such that the
total mass quench airflow therethrough remains the same as prior to
addition of the multiple of trim holes 126. In the disclosed
non-limiting embodiment, three (3) trim holes 126 are located in
each panel of the inner combustor liner assembly 62. That is, for
each quench hole 124 there are three (3) trim holes 126 in the
disclosed non-limiting embodiment.
[0064] Generally, the disclosed non-limiting embodiment,
approximately 35% of the total airflow into the combustor section
26 enters the swirler 90; 25% is utilized for liner assembly
cooling; 35% is utilized as quench flow; and 5% is trim flow. The
quench hole 122 through the outer combustor liner assembly 60 and
the quench hole 124 through the inner combustor liner assembly 62
provide a generally two-thirds/one-thirds flow split of the
35%.
[0065] With reference to FIG. 16, the trim holes 126 are located
near the combustor exit and NGVs 54A to tailor the radial
temperature distribution factor, also known as the profile factor
of the combustion gases. The overall spatial distribution of
temperature within the combustor chamber 66 is referred to herein
as the engine "pattern factor," which can be defined as the
difference between maximum combustor temperature (T.sub.max) and
mean combustor exit temperature (T.sub.4) divided by the difference
between the mean combustor exit temperature (T.sub.4) and the
combustor inlet temperature (T.sub.3). Pattern
Factor=[(T.sub.max-T.sub.4)/(T.sub.4-T.sub.3)]. That is, The
Pattern Factor reflects the extent to which the maximum temperature
deviates from the average temperature rise across the combustor,
while the Profile Factor characterizes the extent to which the
maximum circumferential mean temperature deviates from the average
temperature rise across the combustor. In one example, the average
exit temperature is 2,000.degree. F. (1093.degree. C.) with the
maximum combustor temperature rise of 2,600.degree. F.
(1426.degree. C.) which provides a pattern factor of approximately
0.3.
[0066] A low pattern factor facilitates a relative increase in
service life and reduced fuel burn through reduced thermal effects
upon the Blade Outer Air Seal (BOAS) and turbine blade creep in the
downstream turbine section 54. That is, an increase in engine
fuel-air ratio facilitates increased thrust yet acceptable
combustion gas exit temperature distribution in the pattern factor
and profile factor for turbine component operations and service
life. A low pattern factor also facilitates usage of less cooling
air to cool stationary turbine structures such as the 1st and 2nd
vanes and blade outer air seals (BOAS). If airflow is reduced to
maintain the same metal temperature, a lower fuel burn (TSFC)
results for the same service life. If airflow is not reduced, the
longer turbine service life results for the same fuel burn
(TSFC).
[0067] The radial profile represents the average temperature at
each radial location at the combustor exit. This temperature is
what the rotating turbomachinery, such as the turbine blades
experience. Generally, the radial profile needs to be cooler toward
the inner radius to prevent 1st and 2nd blade creep. As the pattern
factor is reduced and the exit temperature becomes more uniform,
the radial profile typically becomes more uniform resulting in a
hotter than desired temperature toward the inner radius.
[0068] The circumferentially uniform fuel/air mixture, facilitated
by the swirler 90 that has uniform high swirl effectively interacts
and mixes with quench zone 122 and 124 (FIG. 15) to produces a low
pattern factor.
[0069] For operating conditions that are close to stoichiometric,
the lower pattern factor produced by the swirler 90 and quench
holes 122, 124 reduce the amount of temperature rise in the turbine
section 28 from fuel rich streaks that may convert CO to CO2.
Airflow from the quench holes 122, 124 is moved to the trim holes
126 in such a way to maintain an effective quench zone and low
pattern factor while cooling the inner radius to acceptable levels.
That is, the trim holes 126 tailor the radial profile of the
combustor gas flow to achieve desired results.
[0070] The swirler 90 and the trim holes 126 facilitate an increase
in engine fuel-air ratio facilitates increased thrust yet prevents
secondary heat release in the turbine section 28. Airflow that
feeds the swirler 90 is more uniformly distributed through the
removal of non-uniform mount features that increase swirler
interaction with the quench zone and improve fuel uniformity to
create a relatively low pattern factor on the order of less than
0.3 and a radial profile factor of approximately 0.1. The radial
temperature profile is then readily adjusted with the arrays of
trim holes 126.
[0071] With reference to FIGS. 17 and 18, the lower pattern factor
produced by the swirler 90 meets the desired target limit but,
without the trim holes 126, the radial profile is too hot towards
the inner span (FIG. 18). The trim holes 126 thereby further tailor
the radial profile in the desired region to, for example,
facilitate a reduction in turbine blade "creep". For example, the
trim holes 126 may be located and sized to trade-off a fuel-air
factor reduction in the 5%-25% span region for a slight increase in
the 55%-100% span region.
[0072] It should be understood that various other trade-offs and
tailoring may alternatively or additionally be provided so as to
form a desired profile. Typically, a "too flat" profile is
undesirable as cooling airflow may be more specifically controlled
in the mid-span regions of the turbine blades and vanes in the
turbine section 28. That is, cooling passages may be readily
refined in the mid-span regions rather than in the blade root
regions.
[0073] 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.
[0074] 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.
[0075] 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.
[0076] 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.
* * * * *