METHODS AND SYSTEM FOR REDUCING PRESSURE LOSSES IN GAS
TURBINE ENGINES
BACKGROUND OF THE INVENTION
This invention relates generally to gas turbine engines and more particularly, to
combustor assemblies for use with gas turbine engines.
At least some known gas turbine engines use cooling air to cool a combustion
assembly within the engine. Moreover, often the cooling air is supplied from a
compressor coupled in flow communication with the combustion assembly. More
specifically, in at least some known gas turbine engines, the cooling air is discharged
from the compressor into a plenum extending at least partially around a transition
piece of the combustor assembly. A first portion of the cooling air entering the
plenum is supplied to an impingement sleeve surrounding the transition piece prior to
entering a cooling channel defined between the impingement sleeve and the transition
piece. Cooling air entering the cooling channel is discharged into a second cooling
channel defined between a combustor liner and a flowsleeve. The remaining cooling
air entering the plenum is channeled through inlets defined within the flowsleeve prior
to also being discharged into the second cooling channel.
Within the second cooling channel, the cooling air facilitates cooling the combustor
liner. At least some known flowsleeves include inlets and thimbles that are
configured to discharge the cooling air into the second cooling channel at an angle
that is substantially perpendicular to the flow of the first portion of cooling air
entering the second cooling chamber. More specifically, because of the different flow
orientations, the second portion of cooling air loses axial momentum and may create a
barrier to the momentum of the first portion of cooling air. The barrier may cause
substantial dynamic pressure losses in the air flow through the second cooling
channel.
At least one known approach to decreasing the amount of pressure losses requires
resizing the inlets in the existing system. However, this approach may require
multiple inlets to be resized at multiple sections of the engine. As such, the
economics of this approach may outweigh any potential benefits.
BRIEF DESCRIPTION OF THE INVENTION
In one aspect, a method of assembling a combustor assembly is provided, wherein the
method includes providing a combustor liner having a centerline axis and defining a
combustion chamber therein, and coupling an annular flowsleeve radially outward
from the combustor liner such that an annular flow path is defined substantially
circumferentially between the flowsleeve and the combustor liner. The method also
includes orienting the flowsleeve such that a plurality of inlets formed within the
flowsleeve are positioned to inject cooling air in a substantially axial direction into the
annular flow path to facilitate increasing dynamic pressure recovery.
In another aspect, a combustor assembly is provided, wherein the combustor assembly
includes a combustor liner having a centerline axis and defining a combustion
chamber therein. The combustor liner also includes an annular flowsleeve coupled
radially outward from the combustor liner such that an annular flow path is defined
substantially circumferentially between the flowsleeve and the combustor liner. The
flowsleeve includes a plurality of inlets configured to inject cooling air therefrom in a
substantially axial direction into the annular flow path to facilitate increasing dynamic
pressure recovery.
In a further aspect, a gas turbine engine is provided, wherein the gas turbine engine
includes a combustor assembly including a combustor liner having a centerline axis
and defining a combustion chamber therein. The combustor assembly also includes
an annular flowsleeve coupled radially outward from the combustor liner such that an
annular flow path is defined substantially circumferentially between the flowsleeve
and the combustor liner. The flowsleeve includes a plurality of inlets configured to
inject cooling air therefrom in a substantially axial direction into the annular flow path
to facilitate increasing dynamic pressure recovery.
BRIEF DESCRIPTION OF THE DRAWINGS
Figure 1 is a schematic cross-sectional illustration of an exemplary gas turbine engine;
Figure 2 is an enlarged cross-sectional illustration of a portion of an exemplary
combustor assembly that may be used with the gas turbine engine shown in Figure 1;
Figure 3 is a perspective view of a known flowsleeve that may be used with the
combustor assembly shown in Figure 2;
Figure 4 is a perspective view of an exemplary flowsleeve that may be used with the
combustor assembly shown in Figure 2;
Figure 5 is a cross-sectional view of an exemplary flowsleeve and an impingement
sleeve/flowsleeve interface that may be used with the combustor assembly shown in
Figure 2; and
Figure 6 is a perspective view of an exemplary combustor liner that may be used with
the combustor assembly shown in Figure 2.
DETAILED DESCRIPTION OF THE INVENTION
As used herein, "upstream" refers to a forward end of a gas turbine engine, and
"downstream" refers to an aft end of a gas turbine engine.
Figure 1 is a schematic cross-sectional illustration of an exemplary gas turbine engine
100. Engine 100 includes a compressor assembly 102, a combustor assembly 104, a
turbine assembly 106 and a common compressor/turbine rotor shaft 108. It should be
noted that engine 100 is exemplary only, and that the present invention is not limited
to engine 100 and may instead be implemented within any gas turbine engine that
functions as described herein.
In operation, air flows through compressor assembly 102 and compressed air is
discharged to combustor assembly 104. Combustor assembly 104 injects fuel, for
example, natural gas and/or fuel oil, into the air flow, ignites the fuel-air mixture to
expand the fuel-air mixture through combustion and generates a high temperature
combustion gas stream. Combustor assembly 104 is in flow communication with
turbine assembly 106, and discharges the high temperature expanded gas stream into
turbine assembly 106. The high temperature expanded gas stream imparts rotational
energy to turbine assembly 106 and because turbine assembly 106 is rotatably coupled
to rotor 108, rotor 108 subsequently provides rotational power to compressor
assembly 102.
Figure 2 is an enlarged cross-sectional illustration of a portion of combustor assembly
104. Combustor assembly 104 is coupled in flow communication with turbine
assembly 106 and with compressor assembly 102. Compressor assembly 102
includes a diffuser 140 and a discharge plenum 142, that are coupled to each other in
flow communication to facilitate channeling air downstream to combustor assembly
104 as discussed further below.
In the exemplary embodiment, combustor assembly 104 includes a substantially
circular dome plate 144 that at least partially supports a plurality of fuel nozzles 146.
Dome plate 144 is coupled to a substantially cylindrical combustor flowsleeve 148
with retention hardware (not shown in Figure 2). A substantially cylindrical
combustor liner 150 is positioned within flowsleeve 148 and is supported via
flowsleeve 148. A substantially cylindrical combustor chamber 152 is defined by
liner 15C. More specifically, liner 150 is spaced radially inward from flowsleeve 148
such that an annular combustion liner codling passage 154 is defined between
combustor flowsleeve 148 and combustor liner 150. Flowsleeve 148 includes a
plurality of inlets 156 which provide a flow path into cooling passage 154.
An impingement sleeve 158 is coupled substantially concentrically to combustor
flowsleeve 148 at an upstream end 159 of impingement sleeve 158, and a transition
piece 160 is coupled to a downstream end 161 of impingement sleeve 158. Transition
piece 160 facilitates channeling combustion gases generated in chamber 152
downstream to a turbine nozzle 174. A transition piece cooling passage 164 is
defined between impingement sleeve 158 and transition piece 160. A plurality of
openings 166 defined within impingement sleeve 158 enable a portion of air flow
from compressor discharge plenum 142 to be channeled into transition piece cooling
passage 164.
In operation, compressor assembly 102 is driven by turbine assembly 106 via shaft
108 (shown in Figure 1). As compressor assembly 102 rotates, it compresses air and
discharges compressed air into diffuser 140 as indicated in Figure 2 with a plurality of
arrows. In the exemplary embodiment, the majority of air discharged from
compressor assembly 102 is channeled through compressor discharge plenum 142
towards combustor assembly 104, and a smaller portion of air discharged from
compressor assembly 102 is channeled downstream for use in cooling engine 100
components. More specifically, a first flow leg 168 of the pressurized compressed air
within plenum 142 is channeled into transition piece cooling passage 164 via
impingement sleeve openings 166. The air is then channeled upstream within
transition piece cooling passage 164 and discharged into combustion liner cooling
passage 154. In addition, a second flow leg 170 of the pressurized compressed air
within plenum 142 is channeled around impingement sleeve 158 and injected into
combustion liner cooling passage 154 via inlets 156. Air entering inlets 156 and air
from transition piece cooling passage 164 is then mixed within passage 154 and is
then discharged from passage 154 into fuel nozzles 146 wherein it is mixed with fuel
and ignited within combustion chamber 152.
Flowsleeve 148 substantially isolates combustion chamber 152 and its associated
combustion processes from the outside environment, for example, surrounding turbine
components. The resultant combustion gases are channeled from chamber 152
towards and through a transition piece combustion gas stream guide cavity 160 that
channels the combustion gas stream towards turbine nozzle 174.
Figure 3 is a perspective view of a known flowsleeve 200 that may be used with
combustor assembly 104. Flowsleeve 200 is substantially cylindrical and includes an
upstream end 202 and a downstream end 204. Upstream end 202 is coupled to dome
plate 144 (shown in Figure 2) and downstream end 204 is coupled to impingement
sleeve 158 (shown in Figure 2). Combustor liner 150 (shown in Figure 2) is coupled
radially inward from flowsleeve 200 such that cooling passage 154 (shown in Figure
2) is defined between flowsleeve 200 and combustor liner 150.
Flowsleeve 200 also includes a plurality of inlets 206 and thimbles 208 defined
adjacent downstream end 204. Inlets 206 and thimbles 208 are substantially circular
and are oriented substantially perpendicular to a flowsleeve center axis 210.
Furthermore, thimbles 208 extend substantially radially inward from flowsleeve 200
such that airflow is discharged from thimbles 208 and inlets 206 from around
impingement sleeve 158, radially inward through flowsleeve 200, and into
combustion liner cooling passage 154. The radial flow direction of airflow entering
passage 154 through inlets 206 and thimbles 208 substantially reduces the axial
momentum of airflow and creates a barrier to air flowing within passage 154 from
transition piece cooling passage 164. Furthermore, the radial length of thimbles 208
creates an obstruction to airflow channeled from transition piece cooling passage 164.
As such, a pressure drop of the airflow results within combustion cooling passage
154. The resulting pressure drop may cause disproportional cooling around
combustor liner 150.
Figure 4 is a perspective view of an exemplary embodiment of a flowsleeve 250 that
may be used with combustor assembly 104. Flowsleeve 250 is substantially
cylindrical and includes an upstream end 252 and a downstream end 254. Upstream
end 252 is coupled to dome plate 144 (shown in Figure 2) and downstream end 254 is
coupled to impingement sleeve 158 (shown in Figure 2). Combustor liner 150 (shown
in Figure 2) is coupled radially inward from flowsleeve 250 such that combustion
liner cooling passage 154 (shown in Figure 2) is defined between flowsleeve 250 and
combustor liner 150.
Flowsleeve 250 also includes a plurality of injectors 256 spaced circumferentially
about flowsleeve 250 at a distance 258 upstream from downstream end 254. In the
exemplary embodiment, injectors 256 are substantially circular and each has a large
length/diameter ratio. In an alternative embodiment, injectors 256 are substantially
rectangular slots having a width that is larger than a slot height. Moreover, injectors
256 are configured to substantially axially eject airflow from around impingement
sleeve 158 through flowsleeve 250 and into combustion liner cooling passage 154.
More specifically, airflow ejected from injectors 256 enters passage 154 in a generally
axial direction that is substantially tangential to a direction of flow discharged into
passage 154 from airflow channeled into passage 154 from passage 164, and in
substantially the same direction as airflow channeled into passage 154 from passage
164. Furthermore, injectors 256 are configured to accelerate airflow ejected
therefrom. An annular gap (not shown) is defined between flowsleeve 250 and
combustor liner 150 within distance 258. Injectors 256 and the annular gap facilitate
regulating pressure in airflow entering combustion liner cooling passage 154.
Figure 5 is a cross-sectional view of flowsleeve 250 and an impingement
sleeve/flowsleeve interface 300. Specifically, Figure 5 illustrates the interface 300
defined between the coupling of flowsleeve 250 and impingement sleeve 158.
Furthermore Figure 5 illustrates a cross-sectional view of the axial injection geometry
of injectors 256. Specifically, flowsleeve 250 is oriented such that injectors 256 are
positioned an axial distance 302 upstream from interface 300. As such, an annular
gap 304 defined at the intersection region of flowsleeve 250 and impingement sleeve
158 has an axial length 302. Annular gap 304 facilitates regulating air flow from
transition piece cooling passage 164.
Figure 6 is a perspective view of an exemplary combustor liner 350 that may be used
with combustor assembly 104. Combustor liner 350 is substantially cylindrical and
includes an upstream end 352 and a downstream end 354. In the exemplary
embodiment, upstream end 352 has a radius RI that is substantially larger than a
radius RI of downstream end 354. Upstream end 352 receives a fuel/air mixture from
fuel nozzles 146 and discharges the fuel/air mixture into transition piece 160.
Combustor liner 350 is oriented within flowsleeve 250 such that flowsleeve 250 and
combustor liner 350 define combustion liner cooling passage 154. Cooling air
received in combustion liner cooling passage 154 is channeled upstream and across a
surface 356 of combustor liner 350 to facilitate cooling combustor liner 350.
Combustor liner surface 356 is configured with a plurality of grooves 358 defined
thereon that facilitate circumferentially distributing the airflow from injectors 256
across liner surface 356. In the exemplary embodiment, grooves 358 are configured
in a criss-crossed pattern across a length LI of combustor liner surface 356 such that
diamond shaped raised portions 359 are defined between grooves 358. In alternative
embodiments, grooves 358 may be configured in other geometrical patterns.
During operation of engine 100 cooling air is discharged from plenum 142 such that it
substantially surrounds impingement sleeve 158. First flow leg 168 enters transition
piece cooling passage 164 through openings 166. First flow leg 168 cools transition
pieoe 160 by traveling upstream through transition piece cooling passage 164. First
flow leg 168 continues through annular gap 304 and discharges into combustion liner
cooling passage 154. Second flow leg 170 flows around impingement sleeve 158 and
enters combustion liner cooling passage 154 through injectors 256. Within
combustion liner cooling passage 154, the first and second flow legs 168 and 170 mix
and continue upstream to facilitate cooling combustor liner 350.
The configuration of injectors 256 increases the velocity of cooling air within second
flow leg 170. The increased velocity facilitates enhanced heat transfer between the
cooling air and combustor liner 350. Annular gap 304 facilitates regulating flow of
first flow leg 168 into combustion cooling passage 154. As such, injectors 256 and
annular gap 304 facilitate balancing the pressure and velocity of the two flow legs 168
and 170 such that a balanced flow path results from the mixing of the two flow paths.
Furthermore, due to the axial configuration of injectors 256, the second flow leg 170
does not create an air dam which restricts the flow of first flow leg 168. As a result,
the axial configuration of injectors 256 facilitates increasing dynamic pressure
recovery within the resultant flow path. By balancing pressure loss and velocity
within combustion liner cooling passage 154, injectors 256 and annular gap 304
facilitate substantially uniform heat transfer between combustor liner 350 and the
cooling air.
Moreover, grooves 358 of combustor liner surface 356 facilitate enhancing the heat
transfer between cooling air and combustor liner 350. Specifically, grooves 358
facilitate circumferentially distributing cooling air from injectors 256 and facilitate
creating a uniform heat transfer coefficient distribution across the length and
circumference of combustor liner 350. In addition, grooves 358 facilitate allowing
high velocity cooling air to facilitate improving heat transfer.
The above-described apparatus and methods facilitate providing constant heat transfer
between cooling air and a combustor liner, while maintaining an overall pressure of
the gas turbine engine. Specifically, the injectors facilitate reducing pressure losses
by injecting the cooling air of the second flow leg axially such that dynamic pressure
recovery is increased between the first and second flow leg. Furthermore, the
enhancements to the combustor liner facilitate greater heat exchange between the
combustor liner and the cooling air.
As used herein, an element or step recited in the singular and proceeded with the word
"a" or "an" should be understood as not excluding plural said elements or steps,
unless such exclusion is explicitly recited. Furthermore, references to "one
embodiment" of the present invention are not intended to be interpreted as excluding
the existence of additional embodiments that also incorporate the recited features.
Although the apparatus and methods described herein are described in the context of a
combustor assembly for a gas turbine engine, it is understood that the apparatus and
methods are not limited to combustor assemblies or gas turbine engines. Likewise,
the combustor assembly components illustrated are not limited to the specific
embodiments described herein, but rather, components of the combustor assembly can
be utilized independently and separately from other components described herein.
While the invention has been described in terms of various specific embodiments,
those skilled in the art will recognize that the invention can be practiced with
modification within the spirit and scope of the claims.

WHAT IS CLAIMED IS:
1. A combustor assembly (104) comprising:
a combustor liner (150,350) having a centerline axis and defining a combustion
chamber (152) therein; and
an annular flowsleeve (148,200,250) coupled radially outward from said combustor
liner such that an annular flow path is defined substantially circumferentially between
said flowsleeve and said combustor liner, said flowsleeve comprises a plurality of
inlets (156,206) configured to inject cooling air therefrom in a substantially axial
direction into said annular flow path to facilitate cooling said combustor liner.
2. A combustor assembly (104) in accordance with Claim 1 further comprising:
a transition piece (160) coupled to said combustor liner (150,350); and
an impingement sleeve (158) coupled radially outward from said transition piece such
that an annular transition piece cooling flow path is defined between said transition
piece and said impingement sleeve, said transition piece cooling flow path configured
facilitate increasing dynamic pressure recovery within said flow path.
3. A combustor assembly (104) in accordance with Claim 2 further comprising
an annular flow gap defined between said combustor liner (150,350) and said
flowsleeve (148,200,250), said annular flow gap configured to regulate flow from said
transition piece (160) cooling flow path into said annular flow path.
4. A combustor assembly (104) in accordance with Claim 1 wherein said
plurality of inlets (156,206) facilitate reducing inlet losses within said annular flow
path.
5. A combustor assembly (104) in accordance with Claim 1 wherein said
plurality of inlets (156,206) facilitate increasing cooling of said transition piece (160)
within said annular flow path.
6. A combustor assembly (104) in accordance with Claim 1 wherein said
plurality of inlets (156,206) are each substantially circular and facilitate increasing a
velocity of cooling air discharged therefrom.
7. A combustor assembly (104) in accordance with Claim 2 wherein an exterior
surface of said combustor liner (150,350) comprises surface enhancements that
facilitate increasing heat transfer between said combustor liner and cooling air
flowing through said annular flow path.
8. A gas turbine engine (100) comprising:
a combustor assembly (104) comprising:
a combustor liner (150,350) having a centerline axis and defining a
combustion chamber (152) therein; and
an annular flowsleeve (148,200,250) coupled radially outward from said
combustor liner such that an annular flow path is defined substantially
circumferentially between said flowsleeve and said combustor liner, said flowsleeve
comprises a plurality of inlets (156,206) configured to inject cooling air therefrom in a
substantially axial direction into said annular flow path to facilitate increasing
dynamic pressure recovery of said flow path.
9. A gas turbine engine (100) in accordance with Claim 8 wherein said
combustor assembly (104) further comprises
a transition piece (160) coupled to said combustor liner (150,350); and
an impingement sleeve (158) coupled radially outward from said transition piece such
that an annular transition piece cooling flow path is defined between said transition
piece and said impingement sleeve, said transition piece cooling flow path configured
to facilitate cooling said combustor liner.
10. A gas turbine engine (100) in accordance with Claim 9 wherein said
combustor assembly (104) further comprises an annular flow gap defined between
said combustor liner (150,250) and said flowsleeve (148,200,250), said annular flow
gap configured to regulate flow from said transition piece (160) cooling flow path into
said annular flow path.