Note : Les descriptions sont présentées dans la langue officielle dans laquelle elles ont été soumises.
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EFFICIENTLY COOLED COMBUSTOR
FOR A COMBUSTION TURBINE
BACKGROUND OF THE INVENTION
The present invention relates to combustion
turbines as may be employed in a variety of uses, such as
industrial processes, electric power generation, or air-
craft engines. More particularly, the present inventionis directed to combustors employed in combustion turbines
to heat motive gases which drive the turbine. The des-
cription hereafter details the structure of a combustion
turbine used for power generation, but applies as well ~o
any type of combustion turbine.
A combustion turbine used for electric power
generation typically has some eight to sixteen combustors
peripherally arranged about the turbine longitudinal axis.
Combustion of fuel within each combustor heats pressurized
15 air supplied from a compressor section of the turbine.
The hot motive gas generated within ~ach combustor flows
by means of a transition duct to a turbine section of the
combustion turbine. Impingement of the hot motive gases
on turbine blades causes rotation of 'che turbine rotor and
generation of power.
Continuing ~fforts to obtain higher power gener-
ation at greater efficiency have resulted in a continuous
increase in gas operating temperatures. Even with the use
of impro~d stainless steel or nickel-based alloys such as
Hastalloy~or Inco ~ 7, it has become increasingly diffi-
cult to provide ade~uate cooling of combustors so as to
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assure long combustor life. At present, the temperature
of operating gases is typically 2300F while the tempera-
ture operating limit for special metals used in construc-
tion of combustors is about 1500F. It is expected that
turbine gas operating temperatures will increase even
further in the future.
Typical prior art combustors are presently
cooled by introduction of air in tha form of a thin film
along an interior surface of a ~ombustor wall. The pri-
mary purpose of the film of air is to prevent impingementof hot motive gas upon the interior surace of the com-
bustor wall. Such interior film cooling has generally
been adequate. However, it is becoming less so with
increasing turbine qas operating temperatures for three
reasons. First, depletion of the cooling film with hot
motive gas tends to defeat the insulating properties of
the cooling film. Second, inability of the cooling air to
persist as a jet film due to the force of impingement by
the hot mot:i~e gases destroys the film effect and reduces
the effectiveness of th~ cooling techniqu~. Finally, the
use of fuels which generate a highly radiative flame
increases the heat load on the combustor wall.
A second prior art combustor cooling technique
involves use of a porous, laminated liner within the
combustor, both to cool the liner as well as to provide a
film of cooling air between the liner and the hot motive
gases. The laminated liner comprises a plurality layers,
each having channels formed therein, the channels of
adjacent layers oriented generally perpendicular to one
another. The channels of adjacent layers oriented gener-
ally perpendicular to one another. The channels of adj-
acent layers are interconnected by means of holes in the
layers to provide flow communication through the liner to
the interior of the combustor.
Transpiration cooling has also generally been
adequate. However, increasing gas operating temperatures
has resulted in the same type of problems for a trans-
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piration-cooled combustor as were described above for a
film-cooled combustor. In addition, transpiration cooling
requires significant volumes of compr~ssor discharge air
to provide ade~uate cooling, depleting the air which would
normally be used to drive thP turbine end thereby reducing
turbine efficiency.
Hence, it would be advantageous to develop a
cooling apparatus which provides improved cooling effi-
ciency for a combustor wall, thereby permitting hisner gas
operating temperatures and resulting in longer combustor
life, and reducing the volume of cooling air drawn from
the compressor section, permitting higher turbine operat-
ing efficiency.
SU~ ~ RY OF THE INVENTION
Accordingly, a combustor for a combustion tur-
bine comprises a generally cylindrical combustor preer-
ably formed from a plurality of ring segments coupled
together with a conical segment at a fuel-injection end of
the combustor. Each segment comprises an outer shell
member and an inner skin member, a combination thereof
deining coolant channel means for dirscting flow of
cooling air. Cooling air enters from the combustor exter-
ior at one end of each segment through the shell member,
flows through the channel means between the shell and skin
members removing heat therefrom, and exits to the interior
of the combustor at the opposite end of the ring segment.
BhlEE DESCRIPI IN OF THE DRAWINGS
Fig. 1 shows a longitudinal section of a combus-
tion turbine in which a combustor is arranged to be cooled
in accordance with the principles of the invention;
Fig. 2 shows an elevational view of an upper
portion of the comh~lstor along a longitudinal section
thereof;
Fig. 3 shows a ring subassembly of the combustor
in section;
Fig. 4 shows an upstream end view o a cross-
~ection of the ring subasse~bly at a coolant inlet hole;
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Fig. 5 shows an upstream end view of a cross-
section o the ring subassembly at a point downstream of
the coolant inlet hole;
Fig. 6 shows a top view of a cross section of
the ring subassembly depicting spiral channels within a
skin member;
Fig. 7 shows an enlarged view of a portion of
Fig. 5 with an alternative configuration of channelsi
Fig. 8 shows an alternative embodiment of the
ring subassembly in section.
DESCRIPTION OF THE PREFERRED EMBODIMENT
More particularly, there is shown in Fig. 1 a
combustion turbine 10 having a plurality of generally
cylindrical combustors 12. Fuel is supplied to the com-
bustors 12 through a nozzle structure 14 and air is sup-
plied to the combustors 12 by a compressor 16 through air
flow space 18 within a combustion casing 23.
Hot gaseous products of combustion are directed
from each combustor 12 through a transition duct 22 where
they are discharged into the annular space through which
turbine blades 24, 26 rotate under the driving force of
the expanding gases.
To provide high turbine operating efficiency and
extended combustor operating life, a combustor is ætruc-
tured with enhanced wall cooling in accordance with theprinciples of the invention. The preferred structure for
the combustor 12 is shown in Figs. 2 8.
The combustor 12 (Fig. 2) compri~es a plurality
of cylindrical ring subassemblies 30 and a conical ring
subassembly 32 all joined by appropriate means such as
welding. The lengkh of a single ring subass~mbly 30, 32
is shown as extending between reference numbers 34 and 36.
Cylindrical ring subassemblies 30 are joined end-to-end to
form a cylindrical combustor assembly 12. The conical
ring subassembly 32 is attached to the combustor assembly
12 at a fuel intake end of the assembly 12. Each ring
subassembly 30, 32 comprises a cooling mechanism complete
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within itself but cooperating with the cooling mechanism
of every other ring subassembly 30, 32 to provide an
efficiently cooled combustor assembly 12.
Fig. 3 shows a ring subassembly 30 of the com-
bustor 12 in section. Structure of a conical æubassembly32 is substantially like that described below for a cyl-
indrical subassembly 30. Each ring subassembly 30 com-
prises a cylindrical exterior shell 38 surrounding and
adjoined to a cylindrical interior skin 40. The exterior
shell 38 is continuous from one end of the ring sub~
assembly 30 to the other end about the entire exterior
surface of the subassembly 30 except for a plurality of
coolant inlet holes 42 positioned around the circumference
of the subassembly 30 at the end of each subassembly 30
nearest the fuel intake end of the combustor assembly 12.
The inlet holes 42 provide an entrance for
cooling air 44 into a cross channel 46 defined by an
inward facing groove in the shell 40 and extending around
the circumferenc~ of the ring subassembly 30. Fig. 4 de-
picts the inlet holes 42 and the cross channel 46 from adifferent view. The cross channel 46 is in flow communi-
cation with a plurality of coolant channels 48 between the
shell 38 and the skin 40. The coolant channels 48 shown
in greater detail in Figs. 4 through 7, extend from one
end of the subassembly 30 to the opposite end.
Fig. 6 depicts spiral grooves 50 which are
formed in an outer surface of the skin 40 by an appropri-
ate technique such as machining or etching. The grooved
surface of the skin 40 is attached to khe inner surface of
the shell 38 to define parallel coolant channels 48 which
æpiral longitudinally about the circumference of the
æubassembly 30 from the cross channel 46 to khe exit point
52. The coolant channels 48 can, for example, be .03
inches wide by .03 inches deep and spaced from each other
by .03 inches.
An alternative to the rectangular cross æec-
tional structure of coolant channels 48 is depicted in
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Fig. 7. In this arrangement, the coolant channels 48 are
formed by grooving the interior surface of the shell 38.
The walls 54 of the channels 48 are canted to provide a
decrease in actual surface contact between the shell 38
and the skin 40 and a resultant increase in surface con~
tact between the skin 38 and cooling air within the cool-
ant channels 48. Canting the channel walls 54 also in-
creases surface contact between the channel walls 54 and
the cooling air. The canted channel arrangement thus pro~
vides for more efficient transfer of h~at from the inter-
ior skin 40 and exterior shell 38 to the cooling air
within the channel 48.
The spiral arrangement of coolant channels 48 is
preferred over a strictly longitudinal arrangement.
Spiraling the channels 4~ increases the effective length
of the channels 48 without increasing the length of the
ring subassembly 30. The length of a spiral channel 48
may be controlled by the angle chosen for the spiral.
Spiraling the channels 48 also improves the distribution
of coolant channels 48 within the subassembly 30. For
example, the spiral channels 48 more effectively protect
the combustor assembly 12 from a streak of hot motive gas
flowing in a straight line through the combustor assembly
12 by providing the cooling effect of a plurality of
cooling channels 48 as opposed to the cooling effect of a
few longitudinal channels.
Control of cooling air flow within a ring sub-
asse~bly 30 may be accomplished by variation of fiv~
structural features: (l) the cross-sectional dimensions of
a coolant channel 48; (2) the number of coolant channels
48; (3) the angle of spiral of the channels 48; (4) the
length of the ring subassembly 30; and ~5) the dimensions
of the inlet holes. Each ring subassembly 30 may be
individually designed to achieve the cooling characteris-
35 tics required for its position within the combustor assem-
bly .
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An alternative structure o the ring subassambly
30 is depicted in Fig. 8. Each ring subassembly 30 com-
prises an exterior shell 38, an interior skin 40 and an
interior thermal barrier 60 bonded to the inner surface of
the skin 40. The relatively short leng~h of the ring
subassembly 30 and the non-porous nature of the skin 40
permits easy application of the thermal barrier 60 to the
interior of the combustor 12. The thermal barrier 60,
typically 0.015 to 0.025 inches in depth, may be any
ceramic or other thermal barrier coating, such as
yttrium-stabilized zirconium oxide, which effectively
insulates the skin 40 and shell 38 of the ring subassambly
30 from the harsh interior temperatures of the combustor
12. The thermal barr.ier 60 provides a feature with struc-
tural characteristics such as material type, grade andthickness which may be varied with each subassembly to
achieve desired insulating characteristics.
