CA2039756A1 - Stator having selectively applied thermal conductivity coating - Google Patents

Abstract

Patent 13DV-9640 Abstract of the Disclosure A gas turbine engine stator and method are disclosed which are effective for controlling non-axisymmetric radial running clearance between a stator flowpath surface and rotor blade tips. A coating of preselected thermal conductivity is predeterminedly disposed along the circumference of the stator to control circumferential thermally induced distortion and thereby control non-axisymmetric radial running clearance. In preferred embodiments of the invention, a high thermal conductivity coating may be applied to an annular flange for reducing temperature gradients therein. In another embodiment of the present invention, low thermal conductivity coatings may be applied to a stator to preferentially insulate the stator for reducing circumferential thermal distortion.

Description

Patent 13DV-9640 20397~, STATOR ~AVING S~L~CTIVELY APPLIED T~ERMAL
COND~CTIVITY COATING

The present invention relates generally to 5 gas turbine engine rotor and stator assemblies, and, more particularly, to a means and method for reducing radial clearance between the stator and rotor due to circumferential distortions, A conventional gas turbine engine includes a 10 rotor having a plurality of circumferentially spaced - rotor blades extending from a rotor disc and a stator assembly having a flowpath ~urface disposed adjacen~ to tips of the rotor blades which define a radial running clearance therebetween. The running clearance should be 15 maintained as small as poRslble to ensure that all ~luid flow is channeled through and not around the rotor blades for maximizing ener~y ~:ransfer between the fluid and the rotor blades.
The running clearance in a gas turbine engine 20 changes both during transient operation of the engine and at varying s~eady state operations of the engine.
~his is due primarily to di~ferential thermal movement, including expansion and contraction of structures in the engin~ as the engine is increased or decraased in 25 power. The gas turbine engine includes numerous structures, and airflow and combustion gases of varying temperature, ~he application of different temperature to struc ures results in differen~ial thermal movement during tran~ient and, or steady state operation of the 30 engine. And, the application of a constant temperature to a generally symmetric structure having struc~ural variation~ therein, due for example by machining tolerances, can also result in differential thermal movement.

P~tent 13DV-9640 2~39756 Unle 5 differential thermal movement in the ~as turbine engine structures is accommodated, thermal distortion and stress therefrom are generated. ~or example, upon increasing power of a gas turbine engine, rotor blades therein typically are heated and expand faster than a surrounding stator. This can result in the rotor blade tips rubbing against the stator unless the running clearance is initially set relatively large to avoid such rubs. Such large running clearance is undesirable since it decreases the efficiency of the engine.
The stator and rotor will continue to heat and expand until steady-state heat flux conditions are achievPd. The stator will typically reach steady-state quicker than the rotor due to its lower thermal mass.
Thus the running clearance m~y vary significantly during transient operation.
Alternativaly, the blades may be allowed to intentionally rub the ~tator by using ~lades with abrasive tips to cut the staltor surface round at a condition of maximum overlap between the stator and the rotor blades. However, a relatively large running clearance will then occur for all conditions of operation of the engine other than the one at which the ~5 maximum overlap occurred.
Conventional active clearance control structures are known for predetermlnedly channeling a cooling fluid to a stator for minimizing the running clearance between the stator and the rotor blades.
~owever, such systems are relatively complex and do not correct the cause of the varying running clearance.
Furthermore, the running clearance may be non-axl~ymmetric which requires more complex means for attempting to accommodate such non-axisymmetric running clearances as compared to axisymmetric running Patent 13DV-9640 2~397~i6 clearances. Even in situations where the non-axiqymmetric running clearance i~ predeterminable,the means for accommoda~ing such non-axisymmetric clearances, for example by using ac~ive clearance control, is relatively complex and does not necessarily effectively accommodate for such non-axisymmetric clearances, nor does it attempt to correct the cause thereof.
Furthermore, some non-axisymmetric running clearances are due to random occurrences such as a leak of a cooling fluid or a hot combustion gas on an adjacent ~tructure which results in circumferential distortion and a non-axisymmetric running clearance.
Once conventional means for accommodating such an occurrence would be to design for a worst case scenario and have a relatively large running clearance which is undesirable.

Q~ Qf ~he Invçn~iQn Accordingly, one object of the present invention i~ to provide a new and improved gas turbine engine stator.
Another object of the present invention is to provide a stator effective for accommodating non-axisymmetric runnlng clearance.
Another object of the present invention is to provide a ~tator e~fective for accommodating non-axisymmetric running clearance due to random factors.
Another object of the present invention i~ to provlde a stator effectiYe for controlling non-axisymmetric radial running clearance by controlling circumferential thermal di~tortion~ of a stator.
Another object of t~e present invention is to provide a stator effective for reducing non~axisymmetric running clearance.

~atent 13DV-9640 2~39~56 Another object of the present lnvention i8 to provide a stator effectlve for reducing a cause of non ax~symmetric running clearance.
Another object of the pre~ent invention is to provide a qtator having relatively simple means for controlling non-axisymmetric running clearance.

a.~--~

~ gas turbine engine s~ator effective for surrounding a rotor having a plurality of circumferentially spaced blades to define a radial clearance between the stator and blade tips is disclo~ed. A method and means for controlling non-axisymmetric radial runnlng clearance between the ætator and the blade tips includes a coating of preselected thermal conducl:ivity predeterminedly disposed along a clrcumferenc~3 of the stator. In an exemplary embodiment of the invention, the coating ha~
eithex low or h~gh thermal conducti~ity for reducing thermally induced circumferent:Lal distortions of the ~tator.

~çf ~ ri~tlon~of ~ Dr~win~

The novel features believed characteristic of th~ invention are set forth and differentiated in the claim~. The invention, together with additional objects and advantages thereof, is more particularly described in conjunction with ~he accompanying drawing in whichs Figu~e 1 i6 a schemati~, sectional view of an axisymmetric gas turbine turbo~an engine.
~ igure 2 i5 a perspective, cut away view of a portion of a high pressure compressor in the engine illustrated in Figure 1.

P~tent 13DV-9640 2~3975~

Figure 3 i8 a transverse, partly ~chematic sectlonal view of the high pressure compres30r illustrated in Figure 2.
F~gure 4 is a sectional view of a high pres~ure turbine and adjacent structures in the ngine illu~trated in Figure l.
Figure 5 is a partly ~ectional, perspective view of a portion of a flange used in a ~tator adjacent to the high pressure turbine illustrated in Figure 4.
Figure 6 is a schematic representation of a portion of tha circumference of the stator casing and adjacent structure illustrated in F~gure 4.
Figure 7 is a schematic representation of a transverse sectional view of the flange illustrated in F~gure S showing nominal and distorted positions thereo~
due to diferential temperature~3 therein.
Figure 8 is a schemiltic representation of a transverse sectional view o~ the high pressure compre~sor illustrat~d in Figure 2 showing nomimal and distorted positions thereof due to differential temperatura~ therein.
Figure 9 is a schematic, transverse sectional view of the high pressure turbine stator ca~ing illustra~ed in Figure 4 showing nominal and distorted positions due to differential temperatures therein.
Figure 10 is a schematic, sectional view o~
the high presqure turbine stator casing illustrated in Figure 4 showing a single inlet air tube and a thermal conductivity coating adjacent there~o.

~~3L~

Illustrated in ~igure l is a schematic representation o~ an exemplary high bypass turbofan gas turbine engine lOo The enginQ 10 includes in serial Patent 13DV-9640 ~)39756 flow communication about a longitudin~l centerline axis 12 conventional structure3 including a fan 14, a low press ur e co~pressor (LPC), or booster compres~or 16, a high pressure compressor (~PC) 18, a combustor 20, a high pres~ure nozzle 22, a high pre~sure urbine (HPT~
24, and a low pressure turbine (LPT) 26. The low pres~ure turbine 26 i~ joined to both the fan 14 and the LPC 16 by a conventional first rotor shaft 27~ and the HP~ 24 is joined to the HPC 18 by a conven~ional second rotor shaft 28 for independent rotation relative to the first shaft.
In operation ambient air 29 is channeled into the fan 14 of the engine 10, and a ~irst portion 30 is channeled into the LPC 16 for compression, and a second lS portion 32 bypasses the LPC 16 for providing thrust from the engine 10. The air first portion 30 i8 compressPd in the LPC 16 and further compressed in the HPC 18, channeled to the combustor '20, mi~ed wlth fuel to undergo combustion ~or generating relatlvely hot combustion discharge ga3es 34 which are channeled through the HP nozzle 22 for driving the ~PT 24 and the hPT 26.
The engine 10 operates from low to high power setting3 for powering an aircraft during various modes o~ operation including idle, take-off, cruise and descent. The engine 10, therefore, operates under tran~ient condltions upon acceleration or deceleration of the first and second rotor shafts 27, 28 as the engine is either powered up or powered down during operation. ~he engine 10 also operates at steady state conditions, such as, for example at aircraft cruise wherein the power of the engine 10 remains at an intermediate fixed amoun~ and the speeds of the firQt and second rotor shafts 27, 28 are relatively constant.

Patent 13DV-9640 ;~39~6 Since the HPC 18 compresses the ambient air 29 for generating the compressed air 30, the alr 3 undergoes heating, which can typically reach up to about 1,100F. The combustion discharge gases 34 are a~
temperatures up to about 2,000~. Both the compr~ssed a~r 30 and the combustion gases 34 heat adjacent struc~ures in the HPC 18 and the ~PT 24, respectively, thus providing temperature gradients therein which must be accommodated for reducing thermally induced stress and thermally induced distortion.
More specifically, and for example, the ~PC
18 includes a stator in the form of an annular casing 36 as illustrated in Figures 1 and ~ which surrounds blade rows each including a plurality of circumferentially spaced blades 38 extending radially outwardly from the rotor 28. Referring also to Figure 3f which illustrates a cross section of the HPC 18 showing a single blade 38 for clarity, each of the blades 18 includes a blade tip 42 at a radially outer end thereof which is spaced from and faces an annular stator flowpath surface 44~ The flowpath surface 44 is the radially inner surface of the ~PC casing 36 which is positioned around the blade tips 42 to define a radial running clearance Cr~
Accordingly, as the air 30 is compressed in the ~PC 18, it is heated and therefore heats the blades 38. The blades 38 expand upon heating and the running clearance Cr is thereby affected. ~ypically, the casing 36 does not heat as fast as the blades 38 and therefore does not expand as quickly as the blades 38. The running clParance Cr~ thereore, must be sufficiently large to avoid rubbing of the blade tips 42 against the flowpath surface 44 which iq conventionally known.
A similar, conventionally known running clearance Cr is also found in the MPr ~4. More Pa~ent 13DV-9640 197~6 i specifically, and referring to Flgures 1 and 4, the ~PT
24 includes a stator in a convent~onal form of an HPT
annular caæing 48. The casing 48 includes fir~t and ~econd axially spaced annular flanges 50 and 52, respectively, formed integrally therein, Each flange 50 and 52 includes a radially outer portion 50a and 52a extending radially outwardly from a radially outer ~urface 54 of casing 48. ~he flanges 50 and 52 also include radially inner portions 50b and 52b, respectively, which extend radially inwardly from a radially inner surface 56 of the c sing 48. A
conventional intermediate turbine nozzle 58 is ~paced radially inwardly from the casing 48 and includes a plurallty of circumferentially spaced hollow no~zle vane~ 60 sultably joined to a radially outer nozzle casing 62. The nozzle casing 62 includes integral, axially spaced first and second annular ~langes 64 and 66, respectively, which ara conventionally joined to the casing flrst and second flanges 50 and 52, respectively~
The HPT 24, in this exemplary embodiment, includes a first rotor stage 68 and a second rotor stage 70 both joined to the second rotor 28 which provides power to the ~PC 18 from ~PT 24. The first stage 68 include~ a plurality of circumferentially spaced blades 72 disposed between the ~PT nozzle 22 and the intermediate nozzle 58. Each of the blades 72 has a blade tip 74 which is spaced from and faces a conventional shroud 76. More speclfically, the ~hroud 76 includes an inner flowpath surface 78 which is positloned around the blade tips 74 to de~ine the radial running clearance Cr. The shroud 78 has a downstream end 30 which is conventionally ~ecured to and ~uppor~ed by the flrst flange S0 at the flange inner portion 50b, and an upstream end 82 conventionally secured to and supported by the casing 48.

Patent 13DV-9640 ~:~39~$~
_g_ The HPT running clearance Cr must al80 accommodate differential thermal movement between the blades 72 and the shroud 76 during operation in a manner similar to that de~crlbed above for the HPC running clearance Cr. The combustion gases 34 are relatively hot and heat the blades 72, thereby caus~ng expan~ion thereof. The casing 48 and the shroud 76 are cooler than the blades 72 and therefore do not expand as quickly.
Furthermore, the running clearance Cr of both the HPC 18 and the HPT 24 may not only be axisymmetric relative to the engine longltudinal centerline 12, but may be non-axisymmetric, which i8 considerably more complex to accommodate. A significant contributor to non-ax~symmetric radial running clearance Cr variation~
i9 circumferential thermal distortion of the flowpath structure. ~his distortion includes at least two types.
The first type of thermal distortion can be shown by examining Fi~ures 4 and 5. Flgure 5 ~ illustrates in more particularity, the portion of the ~PT first flange 50 showing a nominal radius thereof r and a th~ckness t of the fir~t flange outer p~rtion 50a which extends above the casing outer surface 54. Sinc~
the flange 50 supports in part the shroud 76, thermal disto~t~on therein affacts the radial position o~ the shroud 76 and ~herefore the amount of the H~T running clearance Cr. Due to manufacturing tolerances, the thickness t and radius r may vary around the circumference of th~ flange 50. Such variation is a random variation which causes a variation in thermal mass around the circumference of the flanga ~0. This can re~ult in both a transient and steady state circumferential thermal gradient in the flange 50 whlch can distort the casing 48 out o round and thus create non-axisymmetric running clearances ~r. Since thls is a P~tent 13DV-9640 2~)39756 random occurrence, it i~ therefore not predictable and difficult to accommodate.
The ~econd type of distortion 1~ predic~able and is due to de~ign features located at di~crete circumferential location~ relative to the longitudinal centerline axi~ 12 of the engine which have thermal response characteristics different from the rest of the ~tructures. One example is the horizontal 3plit line flange common to compressor casings.
More specifically, and referring to Flgures 2 and 3, the ~PC casing 3~ includes an arcuate upper portion 36a extend~ng 180 degrees, and arcuate lower portion 36b extending 180 degree~. A pair of coplanar, horizontally extending fir~t and second flanges 84 and 86 respectively, are joined -intagrally to each of the casing upper and lower portions 3ba and 36b, re~pectively, for joining together the upper and lower port~ons 36a and 36b by conventional means such as bolts. The added thermal mass of the flanges 84 and 86 causas them to thermally lag the thermal respon~e of the casing 36 and thereby creates a thermal distortion in the casing 36 in both transient and steady state operation.
Another example of a di~crete design feature having dif~erent thermal response characteristics includes local air ports around a casing used to supply secondary airflow. More specifically, and referring to Figures 1 and 4, the ~PT 24 further inaludes a plurality of circumferentially spaced air inlet tubes 88 joined in 10w communication to the ~PT casing 48. The tubes 88 are conventionally joined ln flow communicatlon to bleed air tubes 90 join~d to the ~P~ ca~ing 3fi for bleeding a portlon of the compressed air 30. Referrlng to both Figures 4 and 6, the turbine nozzle 58 further include3 a plurality o~ circum~erentially spaced inlet holes 92 Patent 13DV-9640 ;Z ~39~56 dispo3ed in the noz~le ca~ing 62. The ~PT ca~ing 48 is spaced from the no~zle caslng 62 and defines a plenum 94 which receives the compressed air 30 from the inlet tube~ 88 for channeling the alr 30 into the nozzle inlet holes 92 and into the hollow nozzle vanes S0 for cooling thereof as i8 conventionally knownO
Durlng operation of the engine 10, compressed alr 30 is channeled through the inlet tubes 88 into the plenum 94 and genQrates a temperature gradient in the casing 48. This temperature gradient generates circumferential distortion in the casing 48 during bo~h transient and steady state operation, and inasmuch as the casing 48 supports the shroud 76 through the first flange 50 and the .shroud downstream end 80, the HPT
.15 running clearance Cr is af~ected.
The above types of non-axisymmetric running clearance variation may be add.Ltionally appreciated from an examination of the schematic representations illustrated in Figures 7, 8 and 9. More specifically, Flgure 7 illustrates the schematic representation o~ a nominal, or average radial po~ltion 94 of a flange, such as the ~lange S0 lllustrated in Figures 4 and 5. The nominal po~ition 94 can either be at a steady state condition, or at a ~articular tranQient condition. At the top of Figure 7, a thermal lag, or distortion 96 is illustrated which represents for example, a local section o~ the flange 50 ha~ing either a relatively larger thickness t or a relatively larger radius r which would result in increased thermal mass and therefore a decrease in thermal response upon heating of the flange S0. As a result, such portion of the flange 50 experiencas a circumferen~ial distortion which in this exemplary case is a local distortion in the radial direction due to relative expan~ion less than the 3S ad~acent portions of the flange 50. The circumferential Patent 13DV-9640 ~39~S6 dl~tortion 96 illustrated in Figure 7 can al~o occur, for example, upon a leak of relatively cool airflow upon the flange 50.
Figure 4 illu~trates a conventional clearance control manifold 98 surrounding the casing 48 which receives compressor discharge air 100 from the ~PC 18 through a conventional fluid conduit 102, as shown in Figure 1. If a portion of the air 100 should leak from the manifold 98 at a discrete point against the flange 50, the circumferential distortion such as is shown in Figure 7 can re ult~
Figure 8 illu~trates schematically the ~PC
ca~ing 36 and horizontal flange~ 84 and 8S. The nominal, or average position of the casing 36 and the flanges 84 and 86 during either a steady ~tate condit~on or at a particular tran~ient condition is represented by the nominal position 104 representing an average radiu~
of the casing 36. A3 the air 30 i8 compressed in the HPC 18, the ca~ng 36 will, for example9 expand ~aster than the horizontal ~langes 84 and 86 since the casing 36 i8 relatively thin and has a relativley low thermal mass as compared to the relatiYely ~hick and high thermal mass flanges 84 and 86. ~ccordlngly~ the resultant relative rad~al position of the casing 36 and flanges 84 and 86 is represented by distor~ed position 106~
The distorted position 106 intersects the nominal position 104 at four nodes 108 wherein the radius of the distorted position 106 is equal to the radius of the corresponding nominal position 104. The distorted position includes two antinodes 110 of maximum radial di~placement relatlve to the nominal position 104 disposed at a 12 o'clock and 6 o'clock po~i~ion~, or symmetrically about a vertical centerline axi~ 112 of the engine 10. Two antinodes 113 of minimum radial Patent 13DV-9640 displacement relative to the nomlnal position 104 are ~ymmetrically d~sposed at 3 o'clock and 9 o'clock positionæ symmetrlcally about a horizontal centerline axls 114 of the engine 10, which is disposed perpendicular to the vertical centerline axis 112.
Figure 8, in conjunction with Flgure 3, clearly indicates that a~ the air 30 i3 compressed in the HPC
18, the ca~ing upper portion 36a and lower portion 3bb expand more than the nominal position 104 while the casing 36 adjacent to and with the horizontal flanges 84 and 8~ thermally lags such expan~ion and results in a negative radial distortion relative to the nominal po~ition 104~
Figure 9 illu~trates a nominal position 116 of the ~PT casing 48 illustrated in Figure 4 at a circumference through transverse centerlines of the inlet tubes 88 at either a steady state condition or at a particular tran~ient condition. As the combustion dlscharge ga~es 34 heat the turbine nozzle 58, heat is conducted and radiated to the casing 48. The compressed air 30 i8 channel through each of the inlet tubes 88 and heats, or cools as the case may be, the casing 4~
adjacent to the tubes 88 differently than the casing ~8 between adjacent tubes 88. In an example where the compre~gor air 30 is effective for cooling t~ casing 48 a~ it enters each of the tubes 88, the distorted position 118 re~ults as shown in Figure 9.
The distorted position 118 includes sixteen c~rcumerentially spaced nodes 120 representing no di~ference in relative radial position between the distorted position 118 and the nominal position 116.
The eight circumferentially 3pa~ed inlet tube 88 are represented in Figure 9 by respective centerline positions at which positions the casing 48 has an antinoda 1~ o~ minimum relati~e radial displacement Patent 13DV-9640 2~39t7S6 from the nominal position 116. Equidiata~tly ~pac~d between adjacent inlet tubes 88 is an antinode 124 of maximum relative radlal d~splacement from the nominal position 116.
Figure 9 clearly indicates that the casing ~8 has thermally induced circumferential di~tortions having maximum radial position at the maximum antinodes 124, since the casing 48 between adjacent inlet tubes 88 i~
relatively hot and therefore expands radially more than the casing 48 adjacent to and at the inlet tubes 88.
Since the inlet tubes 88 provide relatively cool air 30, the casing 48 adjacent to the inlet tubes 88 thermally lag~ in radial expansion and therefore has a relatively ~maller radial po~ition than the adjacent casing 48 between adjacent inlet tubes 88. Since the casing 48 experiences the distorted po~ition 118, the shroud 76, a~ illustrated ~n ~igure 4, which i5 supported by the first flange 50 of the casing 48, will also experience a corresponding circumferent~al dis~orted position re~ulting in a non-axisymmetric runnlng clearance Cr.
In accordance with a preferred embodiment of th~ present invention, me~ns for controlling the non-axi~ymmetric radial clearance Cr between the ~tator flowpath surfaces and blade tips described above is ~5 provided~ The controlling means includes a coating of preselected thermal conducti~ity predeterminedly dispo~ed along the cireumferenca of the stators.
~ ore specifically, and for example, the invention ma~ be practiced in accordance with the flange illustrated in Figures 4 and 5. One embodiment of the invention may include the flange 50 having random variations in the thickness t and the radiu~ r, or, the flange 50 being subjected to a local ~emperature difference~ such as ~or example caused by leaking of a cooling fluid such as the compressor di~charge air 100 Patent 13DV-9640 2~397~i~

against the flange 50. To reduce the circum~erential distortion 96 caused thereby as illustrated in Figure 7, the first flange 50 is preferably provided with a coating 126 having a high thermal conducti~ity.
A~ used in the present description, high thermal conductivity means increased ability for conducting heat as compared to the thermal conductiv$ty of the underlying surface, and low thermal conductivity means a reduced ability for conducting heat as compared to th~
underlying surface. Low thermal conduc~ivity is synonymous with a good heat ln~ulator, whereas a high thermal conductivity is synonymous with a good heat conductor.
Slnce the flange 50 i~ de~cribed above as lS subject to random variations which result in thermal distortion, a high thermal conductivity coatlng 126 is preferred and should co~er the flange 50 as much as possible. For example, the flange 50 includes two axially spaced side sùrfaces 128 join~d to a top surface 130 as illustrated in Figure 5. The coating 126 is applied wi~h a generally conl3tant thickness over the entirety of he side surfaces 128 and the top surface 130~ The high conductivity coating 126, which for example may be relatively pure nickel plated on the surface~ 128, 130, i8 e~fective or transferring heat from the hotter regions of the flange 50 to cooler regions ~or obtaining a more uniform temperature of the flange 50. The flange 50 may comprise conventional Inconel 718 ~IN718~, and nickel has a thermal conductivity approximately five times greater than the thermal conductivity of IN718~ By providlng the coating 125 on the flange 50, di~erential temperatures on the flange 50 are reduced resulting in a reduced di~torted po~ition 132 as illustrated in Figure 7. The reduced distorted position 132 includes an antinoda 13~a having Patent 13DV-9640 2~)3~75~

minimum clrcumferential dlstortion, or radial displacement from the nominal position 94 which i~
substantlally less than that associated with the minlmum antinode 96a of the uncorrected distortion 36.
In accordance with another embod~ment of the present invention, and as illustrated in F~gure 3 and 8, a coating 134 having a low thermal conductivity i8 dlsposed on the innar flowpath surface 44 of both the HPC caslng upper portion 36a and lower portion 36b and generally aligned coextensively with the maximum antinodes 110. More specif~cally, the coatings 134 are disposed along first and second arcs 134a and 134b on the casing upper and lower portions 36a and 36b respectively, symmetrically relative to the vertical centerline axis 112. The extent: of the arcs 134a and 134b, in degrees ~1~ and ~ is determined for each particular design appllcation, but the arc~ 134a and 134b generally ext~nd along the inner flowpath surface 44 under the maximum antinode8 110 up to tbe ad~acent nodes 108 to compensate for the maximum an~lnodes 110.
By introducing the low thermal conductivity coating 134 predeterminedly poæitioned relative ~o the maximum antinodes 110 in the casing 36, heat transfer into the ca~ing 36 at such loc~tions is reduced for better matching the thermal response o the casing 36 away from the flanges 84 and 86 to the tharmal response of tha casing 36 ad~acent to and at the flanges 84, 86. A more uniform temperature response of the casing 36 therePore reduceQ the max~mum and minimum antinodes 110 and 113 as shown as the reduced circumferential thermal distor~ion position 136 illustra ed in ~igure 8. The low thermal conductivity coating 134 may compris~ any aonventional thermal barrier coating such as f for e~ample, a ceramic based mixture which is ~uitably secured to the flowpath surface 44 by conven~ional means æuch as rapid Patent 13DV-9~40 ~:03975~i solidificatlon plasma deposit~onO ~he coating 136 preferably ex~ends for the full axial extent of the ~PC
casing 36~
Accordingly, by introduclng the low thermal conduc~i~ity coating 134 in the predetermined clrcumferential posi~on along the flowpa h ~urface 44 away from the horizontal ~langes 84 and 86, clrcumferential dl~tortion due to differential thermal expansion and contraction of the ca~ing 48 1~ reducedr Since the flowpath surface 44 faces the ~PC blade tips 42, the non-axisymmetric radial running clearance Cr therebetween is al80 reduced from the maximum associated with the distorted position 110 without the coating 134, to reduced value~ assoc~ated with the reduced dlstortion 136 obtalnable by using the coating 134.
In accordance with another embodiment of the present invention, and as illustrated in Figures 4, 6, and 9, a low thermal conductivity coatin~ 138 is disposed on the ~P~ aasing inner surface 56 around each o~ the inlet tubes 88 for reducing the di~ferential, radial thermal movements, iLncluding expansion and contraction, of the casing 48. ~ore speci~ically, the coating 138~ which preferably comprises a thermal barrier coating such as the above mentioned ceramic based mixture suitably secured to the inner surfaca 56, preerably extends from the fir~t ~lange 50 to the second flange 52 and around each inlet tube 88 for a portion of the distance between adjacent tubes 88.
Figure 6 illustrates that the compressed air 30 radially enters each inlet tube 88 at a max~mum velocity and i9 turned circumferentially generally parallel to the casing ~urfaca 56~ ~he air 30 is extra~ted ~or the purpose o~ cooling at inlet holes 92 a3 it flows circum~erentially~ Thus the velocity o~ the str2am of air 30 will decrease as it travels circum~erentlallyO

Patent 13DV-9640 X~39'756 Ad~acent streams of the air 30 will reach zero velocity at about a midplane 140 disposed equidistantly between adjacent inlet tube~ 8B. The alr 30 is effective for cooling the casing inner su~face 56, which coollng i~
S proportlonal to the velocity of the alr 30.
Accordingly, the air 30 most efficiently cool~ the lnner surface 5S adjacent to the inlet tubes 88, and cools the inner surface 56 away from the inlet tubes 88 ~ith a contlnuing decrease in effectivenes.
since the velocity of the air 30 is continually decreasing.
Accordingly, in a praferred embodiment of the invention, the low thermal conductivity coating 138 need only be applied ad~acent to each inlet tube 88 for effectlvene~s. The circumferential extent of the coating 138 between ad~acent in'Let tube 88 is determined for particular design appl~Lcations. Although in a pre~erred embodiment of the invention, the coating 138 has a constant thickness, in another embodiment of the invention as illu8trated in Figure 10, the coating 138 c~n have a varying thicknes~ of a maximum value at the ifltersection of the inner surface 56 and the inlet tube 88, continuously decreasing therefrom toward an adjacent lnlet tube 88. Such a ~arying thickne~s coating 138 ~an more efec~ively match the heat transfer ability of the a~r 30 by providing a large thickness of the coating 138 where the v~locity i8 greate3t with decreasing thicknesses of the coating 138 where the velocity is decreased, to more efectively and uniformly insulate ~he inner sur~ace 56.
Accordingly, by so predeterminedly thermally ln~ulating the inner surface 55 of the casing 48, a reduced circumferential distortion 142 as illus~ratad in Figure 9 may be obtained. The coating 138 is e~fectiva for reducing the amount o~ cooling o~ the casing 48 P~tent 13DV-9640 2~13~75~i --lg--adjacent to each of the inlet tubes ~8, thus resulting in a relatively higher temperature of the casing 148 adjacent to the inlet tubes 88 and reduced minimum and maxlmum antinodes 142a and 142b respectively, as compared to the minimum and maximum antinodes 12~ and 124, re~pectively, of the distorted po~ition 118 which would reau~t without the use of the coating 13~
In all o the above e~bodiments of the present invention, the preselected thermal conductivity coating ~e.g. 126, 134 and 138~ is predeterminedly applied along a circumference of a respective stator for reducing circumferential thermal distortion in the respective stator due to differential thermal mov2ment~, including expansion and contraction~ which correspondingly reduces the differential radial position along the stator circumfer~nce for thusly reducing the respectlve non-axi~ymmetric running clearance Cr. In random occurrences ~uch as those as~ociated with the flange 50, the high thermal coinductivity coating ensures a more uniform temperature o;f the flange 50, thus reducing non-axisymmetric runnlng clearance~ And in the HPC 13 having the horizontal flanges 84, 86 and in the ~PT 24 having the inlet tubes 88, a preferred appl~cation of the low thermal conductivity coating (134 and 138) is effective for reducing the non-axisymmet ic running clearance.
Accordingly9 in accordance with another embodiment of the invention, a method i~ provided for controlling the non~axisymmetric radial clearance between a stator and a ro~or which comprise~ the step o applying the coating having the preselected thermal conductivity along a circumference of the stator positioned for controlling circumferential thermal distortion of the stator which cause~ changes in the 35 radial running clearance between the stator and rotor~

Patent 13DV-9640 39'756 ~n the flange 50 embodiment o~ the invention, the method includes applying the coating 126 having a high thermal conductivity to the flange S0 for reducing temperature gradient~ in the flange 50 and thereby reduclng S differentlal radial thermal movement ~n the flang~ 50~
In the ~PC 18 embod~ment of the invention including the horizontal flanges 84, 86, the method include~ applying the coating 134 having a low thermal conductivity along a circumference away from tha flanges 84 and 86 for reducing di~f~rential radial thermal movement between the casing 36 and the horizontal flange~ 84, 36.
~ n the ~PT 24 Pmbodiment of the invention including the inlet tubes 88, the method includes applying the coating 138 havlng a low thermal conductivity on the casing i~ner ~urface 56 adjacent to each of the tubes 88 for reducing the differential radial thermal movement of the casing 48.
In all three method examples, the reduction in circumferential thermal distortion correspondingly results ln a reduction in the non-axisymmetric running clearance Cr, as well as a reduction in thermally induced stress in the re~pective stator.
The em~odiments of the invention disclosed above may be used to accommodate c~rcumferential thermal di~tortion in the stators both under transient and steady-state operation. ~owevar, the optimum value of the ~hermal conductivity coating and its optimum position will be determined ~or each particular design, and should al80 be determined by evaluating operation at both tran~ient and steady-state operation to ensure that the resulting reduced radial running clearance occurs at desired transient and/or staady-state operations.
While there have been described hereln what are considared to be preferred embodiments of the Patent 13DV-9640 ~:~39~56 pre~ent inYentlon, other modifications o the inventlon ~hall be apparent to tho~e skilled in the art from the teachings hereln, and it i8~ therefore, de~ired to be secured in the appended claim~, all such mod~fications as fall within the true 3pirit and scope o~ the invention.
More specifically, and for example only, a high thermal conductivity coating may al~o be utilized with a flange or ring subject to rad~al thermal gradients from an inner diameter to an outer diameter~
Such temperature gradients will cause stress in the flange or ring due to differential thermal expansion and contraction as well as cause thermal distortion~ By utllizing a hi~h thermal conductivity coating over the ~lange or the ring, the thermal gradients can be reduced by conductlng heat to the colder sections o~ the flange or ring and thereby reduce thermal distortion and stress.
Furthermore, although the preferred embodiments o~ the invention include means ~or controlling non-axisymmetrlc radlal clearance by reduced clrcumferential thermal di~tortions, in other applications, it may be desirable to increase circumferential thermal distortion~ at preselected positions. For example, it may be desirable to increase circumferential thermal distortion in a situation where an increased interference fit between two concentric plloting ~urfaces i8 desired at elevated temperature but a les3er interference fit is desired at ambient conditions for ease of assembly.

Claims (21)

1. A gas turbine engine stator for surrounding a rotor including a plurality of circumferentially spaced blades, each having a tip, comprising:
a stator flowpath surface positionable around said rotor blade tips to define a radial clearance; and means for controlling non-axisymmetric radial clearance between said stator flowpath surface and said blade tips, said controlling means including a coating having preselected thermal conductivity predeterminedly disposed along a circumference of said stator.

2. A stator according to claim 1 further including an annular flange fixedly supporting said stator flowpath surface, and wherein said coating is disposed circumferentially on an outer surface of said flange and has a high thermal conductivity for reducing said non-axisymmetric radial clearance.

3. A stator according to claim 1 further including:
an annular flange having an upper portion including a spaced pair of side surfaces joined by a top surface and a lower portion extending radially inwardly from said upper portion for supporting said stator flowpath surface;
said flange having an antinode of minimum thermal radial growth at a portion subjected to cooling from a cooling fluid channelled thereover; and said coating having a high thermal conductivity and being disposed on said flange side and top surfaces circumferentially around said flange.

Patent 13DV-9640

4. A stator according to claim 3 wherein said coating is nickel.

5. A stator according to claim 1 further including an annular casing having an inner surface defining said stator flowpath surface, and a horizontal flange joined to said casing; and wherein said coating is disposed circumferentially along a first arc on an upper inner surface of said casing positionable toward said blade tips and circumferentially along a second arc on a lower inner surface of said casing positionable toward said blade tips; and said coating has a low thermal conductivity for reducing said non-axisymmetric radial clearance.

6. A stator according to claim 1 further including:
an annular casing having an upper portion, a lower portion, and an inner surface defining said stator flowpath surface;
a pair of coplanar, horizontally extending flanges joined to each of said casing upper and lower portions for joining together said upper and lower portions;
said casing have antinodes of maximum thermal radial growth in said casing upper and lower portions and antinodes of minimal thermal radial growth at said flanges due to heating of aid casing; and said coating have a low thermal conductivity and being disposed on said casing upper portion inner surface and said casing lower portion inner surface to reduce said maximum and minimum antinodes.

Patent 13DV-9640

7, A stator according to claim 6 further including a vertical centerline axis and wherein said flanges are disposed perpendicularly thereto, and said coating is disposed on said casing upper portion inner surface over a first arc relative to said centerline axis, and on said casing lower portion inner surface over a second arc relative to said centerline axis.

8. A stator according to claim 7 wherein said first and second arcs are disposed symmetrically relative to said centerline axis.

9. A stator according to claim 8 wherein said coating is a ceramic based mixture.

10. A stator according to claim 1 further including a turbine outer casing having an inner surface, and a support joined to a shroud, said shroud including said stator flowpath surface; and a plurality of circumferentially spaced air inlet tubes joined to said outer casing for providing compressed air into said outer casing; and wherein said coating has a low thermal conductivity and is disposed around each of said plurality of inlet tubes on said casing inner surface for reducing said non-axisymmetric radial clearance.

11. A stator according to claim l further including:
a stator casing having first and second axially spaced annular flanges and a radially inner surface, each flange having a radially inner portion;
a plurality of circumferentially spaced air inlet tubes joined in flow communication with said stator casing;
a turbine nozzle including a nozzle casing Patent 13DV-9640 spaced from said stator casing and having first and second annular flanges joined to said stator casing first and second flange inner portions, respectively, and a plurality of circumferentially spaced hollow nozzle vanes extending from said nozzle casing, each nozzle vane being in flow communication with a respective inlet hole in said nozzle casing for receiving air from said inlet tube; and a coating disposed on said stator casing inner surface around each of said inlet tubes, said coating having a low thermal conductivity for reducing differential radial thermal movement of said casing.

12. A stator according to claim 11 further including an annular shroud for surrounding said rotor, said shroud being disposed upstream from said nozzle and supported by said stator casing first flange inner portion.

13. A stator according to claim 12 wherein air is flowable through said inlet tubes and generates in said stator casing antinodes of minimum thermal radial growth at each of said inlet tubes and antinodes of maximum thermal radial growth between each of said inlet tubes, said coating being effective for reducing said minimum and maximum antinodes.

14. A stator according to claim 13 wherein said coating comprises a thermal barrier coating of generally constant thickness extending around each of said inlet tubes partially between adjacent inlet tubes.

15. A stator according to claim 13 wherein said coating comprises a thermal barrier coating of varying thickness decreasing from each of said inlet tubes toward adjacent ones of said inlet tubes.

Patent 13DV-9640

16. A stator according to claim 13 wherein said thermal barrier coating is a ceramic based mixture.

17. A method for controlling non-axisymmetric radial clearance between a gas turbine engine stator and rotor comprising the step of applying a coating having a preselected thermal conductivity along a circumference of said stator positioned for controlling circumferential thermal distortion of said stator causing changes in radial clearance between said stator and said rotor.

18. A method according to claim 17 wherein said stator includes an annular flange and said coating applying step includes applying a coating of high thermal conductivity to said flange for reducing temperature gradients in said flange.

19. A method according to claim 17 wherein said stator includes a casing and a horizontal flange joined thereto, and said coating applying step includes applying a coating having low thermal conductivity on said casing away from aid flange for reducing differential radial thermal movement between said casing and said flange.

20. A method according to claim 17 wherein said stator includes an annular casing having an inner surface and a plurality of air inlet tubes joined in flow communication to said casing; and said coating applying step includes applying a coating having low thermal conductivity on said casing inner surface adjacent to said tubes for reducing differential radial thermal movement of said casing relative to said tubes.

21. The invention as defined in any of the preceding claims including any further features of novelty disclosed.