CA2184821A1 - Turbine combustor cooling system - Google Patents

Abstract

A gas turbine engine has a low cost combustor liner cooling system combining the benefits of high internal heat removal with improved film cooling by employing a large number of strategically positioned, laser-drilled cooling passages (28). Cooling air flows through these specially tailored passages (28) to absorb heat from the liner (20) prior to injection as a protective film on the interior surface. The passages (28) are set in staggered rows on a thickened portion of the liner (20) and have a rough internal heat transfer surface and an exit with a steep injection angle (34B) to evenly distribute the cooling film along the interior surfac of the liner.

Description

21 84~1 WO 9~/25932 PCT/US94103138 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 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 ~

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 ~

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 wo 9~/2~932 PCTIUS94/03138 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 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.

Claims (7)

WHAT IS CLAIMED IS:

1. A gas turbine engine combustor of the type disposed in an air supply plenum and adapted to receive air and fuel at one upstream end and discharge hot motive gases from the other downstream end comprising:
a generally cylindrical thin metallic liner having a smooth interior surface exposed to a flow of hot motive gases and a contoured exterior surface exposed to a flow of cooling air in said air supply plenum;
said contoured exterior surface having an axially spaced series of thickened ribs, each extending outwardly therefrom and substantially completely around the circumference of said generally cylindrical liner, the distance between the ribs of said axially spaced series increasing in the direction of flow of said hot motive gases; and a number of small diameter cooling passages drilled through said ribs at a shallow angle to the flow of hot motive gases said number decreasing in each rib downstream from the first of said series of ribs whereby the circumferential spacing between passages in each rib is about equal to the axial spacing between adjacent ribs.

2. The combustor of Claim 1 wherein each of said ribs has an upstream edge intersecting said exterior surface at a sharp angle and said cooling passages are drilled substantially perpendicularly therethrough.

3. The combustor of Claim 2 wherein said sharp angle is between about 95° to 110° and said cooling passages have a length to diameter ratio of more than five.

4. The combustor of Claim 1 wherein said circumferential spacing is about four to twenty times the diameter of said cooling passages at least over a portion of said liner.

5. The combustor of Claim 1 wherein each of said cooling passages has a length to diameter ratio of at least about five.

6. The combustor of Claim 1 wherein said cooling passages are arranged to direct cooling air into the combustor at an angle of less than about 20° to the interior surface.

7. The combustion liner of Claim 1 wherein the cooling passages in one rib are staggered from the passages in each adjacent rib.