Note: Descriptions are shown in the official language in which they were submitted.
21 84~1
WO 9~/25932 PCT/US94103138
TUR51NE COMBUSTOR COOLING SYSTEM
TECHNICAL FIELD
This invention relates generally to power plants in which
combustion products are used as the motive fluid le.g. gas turbine
S enginesl and more ,,ue~ifi~u:'y to a cooled porous combustor liner for
the gas generator portion of such power pl3nts.
BACKGROUND OF THE INVENTION
Gas turbine power plants are used as the primary propulsive
power source for aircraff, in the fomms of jet engines, and turboprop
10 engines as auxiliary power sources for driving air CO~ `a hydraulic
pumps, etc. on aircraff, and as stationary power supplies such as
backup electrical ç,er,~,ulors or hospitals and the like. The same basic
power ~en~:,u,iol~ principles apply for all of these types of gas turbine
power plants. A gas turbine engine in its basic fomm includes a
5 c~l"~,.=..ol section, a combustion section and a turbine section
arran3ed to provide a generally axially extending flow path for the
working gases. Cc." l~n~=aaed air is mixed with fuel and bumed, and the
expanding hot combustion gases are directed asainst stationary
turbine guide vanes in the one or more turbine stages of the.engine.
Z The vanes tum the high velocity gas flow partially sideways to impinge
at the proper angle upon turbine blades mounted on a turbine disk or
wheel that is free to rotate. The force of the impinging gas causes the
turbine disk to spin at high speed. The power so g~n_,ul~d is then used
to draw more air into the engine, in the case of the jet propulsion
25 engine, and both draw more air into the engine and also supply shaff
power to tum the propeller, an electric generator, or for other uses, in
the cases of the other I, " " na. The high velocity combustion gas is
then passed out the aff end of the gas turbine which, in the propulsion
engine ~ F'' ' ~ ns, supplies a forward reaction force to the aircraft.
3û As is well known, the themmal efficiency, and therefore power,
produced by any engine is a function of, among other parameters, the
temperature of the working gases admitted into the turbine section.
That is, all other things being equal, an increase in power from a given
engine can be obtained by increasing the combustion gas
WO 9512S932 2 1 8 ~ 8 2 1 PCT/US94/03~38
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temperature. This is particularly true for small turboshaft or turboprop
engines where very small changes In the operating temperature can
suL,,Iu,,ti .,~ affect the engine output. For example, it has been
det~ " ,;"ed in a typical engine of this type that a single de3ree
5 cer,li~,u~e increase in the temperature of the working gases can
increase the engine power by as much as 15 I ùrsep~ ~r~, . However, as
a practical matter, the maximum ~easible gas temperature is llmited by
the useful operating temperature of the cv""~ ,,l parts in contact
with the motive gas and/or the ability to cool these parts below the hot
gas temperature.
The maximum gas temperatures occur In the combustion
section. A turbine engine conventionally employs e-lther an annular
combustor or several cylindrical combustor cans arranged around the
circ ", ~f~,erlce of the turbine to contain the buming fuel and air and to
lS produce energetic hot gases for introduction to the turbine section. A
transition duct cv" ,il~3 guide vanes is typically disposed between
the combustors and the first turbine stage to properly direct the flow of
hot gases onto the turbine wheel blades.
Various methods for cooling the walls of these combustor
20 COIll~uCJllell~ have been tried in order to allow ever higher ~aas
temperatures to be used. Most methods utilke relatively cool
uncombusted air from the engines cv~ e"ul to both passively cool
the exterior of thê walls by convection and to actively protect the
Interior of the walls by fi~m cooling.
The temm film cooling as used herein refers to the technique of
cooling a surface by r~ui~lluill;~lg a relatively slow movin3 layer or film
of cool air near the surface so that the layer acts as an insulative barrier
to prevent or retard unwanted heating of the surface by the adjacent
hot gas stream. In this context, film coollng Is distlnquished from the
more common convection cooling which operates on the cor"plel :ly
different principle of Illuilllvill;~lg a relatively high velocity flow of
coolin3 air at a surface to carry hea~ away from the surface rather than
insulating the surface from an adjacent heat source.
Several problems exist with the known cooling methods when
applied in smaller high pe,~v",,u,Icé gas turbine engines. Simple film
2 1 8482 ~
WO 95125932 PCINS94/03138
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cooling through slots and/or louvers in the combustor walls does not
utilke the full heat sink potential of the cooling air. Also the amount of
air so used leple~é~ a significant portion of the total air flow from the
COIll~le~ which would otherwise be available to support combustion
5 and control the bumer exit temperature profile, i.e. eliminate hot spots.
To use cooling air more effhciently. recent attempts have focused
on providing hlm cooling through arrays of holes or passages, as
opposed to continuous slots. and constnucting the passages to provide
more active intemal wall cooling by cu"v..,lion or ill"cil~yé",e"l. or
10 both. See, for example, U.S. Patent Nos. 3,420,058: 3,623,711 ; 3,737,1 52:
4,242,871: 4,622,821 and 4,773,'i93. Such cu""' '~ ' cooling
schemes raise new problems to be solved. For example, the uniform
hole pattems nommally employed can result in wall sections that are
u,,de,.uoled on the leading (upstream) edge, well cooled in some
5 central regions, and u~...ooled on the trailing edge as the cooling film
effectiveness increases from row to row in the ~heu~ direction. In
addition, the tuwing ratios much larger than the ideal value near 0.4.
Hence, the effusion jets can separate from the hot surface and mix with
the bulk flow rather than fomming a ~ tC ' .S! film near the surface.
In view of the foresoing, it is an object of the present invention to
provide an improved cooling system for gas turbine combustor walls.
More ,,..e.i~i.ull~, it is an object of the invention to provide a
durable but liUi,t~ combustor liner having a more effective array
of cooling passages therein.
It is a further object of the invention to provide an effhcient
method of making a porous combustor liner for advanced gas turbine
engines.
SUMMARY OF THE INVENTION
The present invention aims to overcome some of the
disadvantages of the prior art as well as offer certain other advantages
by providing a novel c~,,,Linuliu~ of a contoured combustor liner
having rib shaped thickened wall portions and an array of sl,uleui~ul~y
2 1 8~2 1
PCr/Uss410313s
WO 9s/25932 _4
positioned Gnd shaped cooling passases laser-drilled through these
thickened wall portions.
Thickened ribs are fommed around the circ~""~,lce of the
exterior surface of the otherwise thin metallic liner to add structural
5 strength and to increase the effective heat transfer area of the cold
side and of coolins passages drilled at an angle therethrou3h.
To increase the c~ . heat transfer from the combustor
liner the cooling passages are long and narrow with a length to
diameter ratio greater than about 5 and a slightly roughened intemal
1 surface.
To reduce non-unifomm cooling, the distance between the
cooling passages is adjusted, in both the circ~",' ~"" ' and axial
directions simultaneously, to maintain a relatively unifomm cooling effect
over the entire surface of the liner. That is, a large number of these
5 cooling passages are arranged in a row in the circ~""~ _"tial direction
and a number of rows are ir,.,~,i,,u'y spaced apart in the axial
direction such that locally the holes are about eaually spaced from
one another by a distance which increases in the, ' ~v/. I,~leul " rows.
The diameter of each cooling passa_e is reduced as much as
20 possible to minimke the effective area per hole but without increasing
the risk of blockage by debris.
To help prevent the cooling film from s ,u,.,, ' ,9 from the interior
liner surface, th~ entering axial momentum of each jet is increased by
directing the cooling passages at a steep anUIe to the hot gas flow
25 direction. The exit of each passage may also be tapered to act as a
miniature diffuser to further reduce turbulence and the velocity of the
cooling air flowing into the combustor,
BRIEF DESCRIPTION OF THE DRAWINGS
While this ~ ;r~ ;OI ~ concludes with claims particularly
30 pointing out and distinctly claimina the subject matter which is
regarded as the invention, it is believed that the objects, features, and
advantages thereof may be better ul,cl~,,luod from the following
2l8~82~
WO 95125932 PCI'IUS94/~13138
_S _
detailed d~ , of a presently preFerred ~",b~di",t"l when taken
in co~ e~l;ùn with the ac~G" ,pu"~ing drawing in which:
FIG. 1 is a partial cross-sectional view of a combustor section of a
gas turbine en3ine jI~CGI ~U~U~il ,9 the present invention:
FIG. 2A is an enlarged cross-sectional view through the
combustor liner of FIG. 1:
FIG. 23 is an altemate cross-sectional configuration for the
combustor liner; and
FiG. 3 is an enlarged plan view of the exterior surface of the liner
10 of FIG. 2A.
BEST MODE FOR CARRYING OuT THE INVENTION
As an exemplary er"~odi",t:"l of the present invention FIG. 1
illustrates a partial sectional view of a combustor section 10 of a gas
turbine engine. The combustor section 10 includes a generally axially
15 extending hollow annular (or s~",e'; "~s cylindrical) combustion
chamber 30 defined by a thin metallic liner 20 in which c~" l~le"ed air
is mixed with fuel and bumed near the upstream end 31 to provide hot
motive gases for the turbine engine. Fuel is supplied to the chamber 30
through several injectors or spray nozles 40 spaced around the
20 upstream end 31 of the chamber 30 and conne~lèd to a suitable fuel
control system (not shown). A stream of air 14 from the turbine
c~,,,~ ,,.,, !also not shown) flows via a duct 11 into either end of a
plenum 18 surrounding the combustion chamber 30 through the liner
20 as described below and into the chamber where it is heated before
25 being d;~ u~y~:d in an axial direction 39 from the dov/~ t u, l, end 32
to a turbine.
The combustor liner 20 contains several relatively large holes 17
or slots 19 for admitting combustion air 15 into the chamber 30. In
addltion the liner 20 of the present invention contains circul"'f ~Illiully
30 disposed ,~ir,~ i"g ribs 26 each of which have a row of small cooling
passages 28 drilled at an angle therethrough as shown in more detail in
FIGS. 2A and 2B.
.'' ~ ! .,
' .
WO 9~i/25932 2 1 8 4 8 2 I PCT/U594103138 ~
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The thin metallic liner 20 has a generally smooth interior or hot
surface 21 to avoid turbulence in the cooling film and a contoured
exterior or cold surface 22 to promote turbulence nearby. The cold
surface 22 has a number of parallel but spaced apart thickened
S portions or ribs 26 disposed circu~ around the liner 20 so as to
be s~L,Iu"liu::J pe.,u~l,di~ular to the direction of bulk gas flow in the
combustor 10. The thickness of the liner 20 is nominally about .5 to i
mm but is increased in the area of each ,~=i,.fu,~i,,g rib up to about 1~5
to 3.0 mm. Since the thickened ribs 26 provide structural strength as
10 well as sufficient material to provide for steeply analed cooling
passages as discussed below the nominal thickness of the liner 20 may
be ~educed to save weight in the combustor section passaaes, as
discussed belûw the nomirlal thickness of the liner 20 may be reduced
to save weight in the combustor section 10.
As shown in the enlarged sectional view of FIG. 2A, the preferred
shape of the ribs 26 is like a saw tooth with a steep upstream facing
edge 25, CGIIluillil"~ the cooling passages 28 which intersect the edge
25 at a steep angle 33 and intersect the hot surface 21 at a shallow
angle 34, and a shallower sloped ~'~w",l,~", facing edge 27. In an
20 altemate ~,,IL.od;,l,e~, useful with reverse flow combustors and shown
in FIG 2B each of the ribs 26 has an upstream edge 25 which intersects
the thickened cold surface 22 at a relatively steep angle 35 cooling
passages 28 inclined at an an~le 34B to the hot surface 21, and a
d~vn,ll_~,", ed~ae 27, which intersects the cold surface at a relatively
25 shallow angle 37. In both cases the steep angle 35 is preferably
between 90 and 120 and most preferably between 95 and 110.
Angle 37 is generally less than 30 but is not critical to the present
invention.
Each rib 26 contains a variable number of small coolina
30 passages 23 drilled at an angle through the thickened portion of the
liner wall 20. Each passage 28 preferably flares outwardly and intersects
the hot wall surface 21 at an angle 34 or 34E of less than 20 and
preferably at 5 to 15 with respect to the direction of hot sas flow 39. It
is important that the len3th-to-diameter ratio of the passage 28 be at
35 least 5 and ~ fe:lublt about 10 so that any air turbulance g~ d at
the passage entrance is not carried through the liner into the cooling
film and so that there is sufficient residence time for the air to absorb
2 1 8482 1
wo 9~/2~932 PCTIUS94/03138
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heat from the liner. The smal~est practical diameter for each passage
28 is about 0.5 mm and thus the length of each passage is about 2.5 to
5.0 mm. Preferably, the passage is made with a slightly roughened
- surface along its length by laser drilling to aid convective heat transfer.
As shown in FIG. 3. the passages 28 are drilled so that the
circ~""~e,e:r,liul distance 36 between adjacent passages 28 in one row
iS a~ u~ I lul~ly equal to the axial distance 38 between rows. Since the
passages in each row are preferably ,lugg~_d (or offset
cira.",~ s"l;ully by one half the spacing 36) from the passages in
10 adjacent rows, each passage 28 is surrounded by an equal voiume of
liner material so that its heat sink effect is evenly distributed to avoid hot
or cold spots. In a preferred cooling pattem the spacing between
holes 36, 38 is about four times the diameter of each hole (e.g. 2 mm) in
the hotter areas of the combustor 30 near the upstream end 31 and
5 about 20 times (e.g. 10 mm) in the cooler areas near the ' ~v/l l~heulll
end 32. This decrease in hole density whiie Illui,,lui,,i~g a locally
constant spacing prevents .,~ .' ,9 of the ~clvll~lleulll end 32 of
the liner 20 due to accumulationin cooling film thickness.
During operation of the turbine, a portion (about 10-15%) of the
20 relatively cool air from the plenum 18 surrounding the combustion
chamber 30 flows along the exterior surface 22 of the liner wall 20 and
throush the small cooling passages 28 thereby removing heat from the
wall 20 by co"~ ,n. The majority of the air 15 flows through the holes
17 and slots 19 to support combustion. The air flowing from the cooling
25 passages 28 is directed at a very shallow angle 34 along the hot interior
surface 21 of the liner 20 to provide a film of relatively cool air between
it and the much hotter combustion gases in chamber 30. Since
additional air is added to the film from passases 28 in the next
downstream rib 26, it is important that the location of the cooling
30 passages 28 be st~g~ or offset circ~",~:,e"li.,:'~ from the passages
in the ~.,e,cee~i"g and following rib as shown in FIG. 3 so as to more
evenly distribute the cooling air across the hot surface 21. In addition,
the number of cooling passages 28 is reduced and the spacing 38
between the ribs 26 is increased in the ~ u, " direction in order to
35 prevent overcoolins of those portions of the liner 20.
WO 95125932 2 1 8 4 8 2 ~ PCIIUS94/03138
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The liner 20 of the present invention is manufactured by hrst
formin3 the raised ribs 26 on one surface of a thin superalloy sheet by
any Or the well-known methods such as stampin3 or coining, etc. The
sheet or several sheets are then formed into the desired combustor
S shape, typically a generally cnnular hollow rin3, with the ribs on the
exterior surface 22.
The coolins passages 28 are drilled throu3h the ribs 26 by
multiple pulses from a high energy beam, such as a laser beam. The
fommed liner iS F "' ned under the beam so a row of passages 28 may
10 be drilled in each rib by rotatins the liner about its longitudinal axis. Theliner is then moved axially to drill the other do~vr,,l.~:u,,, ribs. Preferably,the passa3es are drilled into the angled upstream edge 25 of the rib 26
as shown in FIG, 2A in order that the laser beam may be directed about
p~"~ i.ularly (i.e. 60 to 90) to the metallic surface, This reduces
S beam ,~rle~l;ui,s and increases the drilling efticiency of the first laser
pulse (by about 50%) at each location of a coolin3 passaae 28.
However, the an~le of beam incidence may be as low as about 20, as
shown in FIG. 2B, if some d~t~ in hole quallty can be tolerated.
While not preferred, the altemate configuration shown in FIG. 2B may
20 be made by drilling the passages 28 from the hot surface 21 so that the
beam entrance will have a sli3htly flared shape to cu" ",ensut~ for the
steeper injection angle 34 required. That is, it is not possible to use a
laser to drill holes at an31es of less than about 20 to the metallic
surface. By usina multiple pulses of the laser beam, the coolin~ pas-
sages may easily be formed with a slightly roughened intemal surface
to improve heat transfer to the cooling air. In addition, laser drillin3
provides high production rates at relatively low cost.
Another major advanta3e of the present invention is the ability to
tailor the cooling ~ ne" ~ccu,~i"3 to variations found in the hot
3as temperature during testin3 by simply drilling additional coolins
passages in the ribs of the hotter areas.
