US2660858A - Air-cooling gas turbine blade - Google Patents

Description

Dec. 1, 1953 L. c. LICHTY AIR-COOLING GAS TURBINE BLADE 2 Sheets-Sheet 2 Filed May 5, 1948' INVENTOR.

Patented Dec. 1, 1953 UNITED STATES PATENT OFFICE AIR-COOLING GAS TURBINE BLADE Lester C. Lichty, New Haven,

Socony-Vacuum Oil This invention is directed to improvement in the operation of power plants and is particularly adapted to use in the gas turbine or jet propulsion engine. 1-

An object of this invention is to provide apparatus for increasing the power output from and efiiciency of turbines by making it possible to use higher than normal driving-fluid temperatures.

A further object of this invention is to provide apparatus for using higher than normal drivingfiuid temperatures in turbines by supplying cooling air at substantially lower temperatures than that of the driving fluid to the turbine.

A further object of'this invention is to provide apparatus for preventing the overheating of the turbine blades of a turbine when using higher than normal driving-fluid temperatures by a novel means of admitting coolant fluid at a temperature substantially lower than the driving fluid to the turbine.

The invention will be described by reference to the attached drawings, in which:

Figure l is a diagrammatic section view of a jet propulsion engine incorporating the instant invention;

Figure 2 is a diagrammatic sketch of a portion of a turbine of a jet propulsion engine incorporating the instant invention; and I Figure 3 is a diagrammatic section view of a manifold, as seen through section 33 of Figure 1, adapted to apply the cooling gases to the turbine of a, jet propulsion engine.

In Figure 1, a gas turbine set is shown diagrammatically which includes a two-stage air compressor, a turbine, and related operating parts. Air is inducted into the first stagecompressor 4 by the motion of the rotor of the compressor. A portion of the air discharged from the first stage compressor is collected in collector ring 5 and supplied to burner 6 through conduit 1. The remainder of the air is conducted through the second stage compressor 8. The compressed air, discharged from the second stage compressor 8, is conducted to the coolant manifold 9 through the space defined by the concentric cylindrical walls I and II. Conduits I2, welded to the coolant manifold 9, conduct the air to the selected nozzle on the nozzle ring I3.

Fuel from a source not shown, is introduced continuously through pipe l4 into the center of burner tube IS. A selected amount of air is directed continuously from the air passing through the burner 42 through the burner tube l and mixes therein with the fuel. The fuel-air mixture is burned continuously, and the hot exhaust gases,

Conn., assignor to Company, Incorporated, New York, N. Y., a corporation of New York Application May 3, 1948, Serial No. 24,721 6 Claims. (01. 60-3919) 2 discharged from the burner tube l5, are mixe with the remainder of the air, which is passed around the outside of the burner tube [5. The mixture of air and exhaust gases is fed through conduit IE to manifold l1, and thereafter fed through conduits l8, welded to the manifold H and connecting with the selected nozzles on the nozzle ring [3. The turbine 19 is adapted to receive through its turbine passages alternately, during rotation, the hot driving fluid, and the coolant air, at a substantially lower temperature than the driving fluid, from their respective noz-' zles on the nozzle ring l3.

The second stage compressor is designed such that the pressure of the air admitted to the nozzles on nozzle ring l3 will be sufiicient to produce a gas velocity of the air as it leaves the nozzle substantially equal to the velocity of the driving fluid as it leaves its nozzle, thereby preventing shock loading of the turbine blades, assuming the 5 same nozzle design.

Suflicient power may be taken from the turbine H! to operate one or both stages of the compressor, or the compressor may be driven by a separate source of power.

Referring to Figure 2, showing a portion of the turbine I9 and the related nozzle ring l3, the relationship of the combustion gas fiow to the coolant air flow through the turbine is shown. The coolant air is admitted to alternate nozzles in the nozzle ring I3, whereby the nozzle ring is cooled. The rotor blades 20 are contacted alter-' nately. with relatively cool air and with hot combustion gases, enabling the temperature ofv the blades to be maintainedbelow thepre-existing maximum temperature limit, while using hottercombustion gases.

Figure 3 is a, diagrammatic section through the coolant manifold 9, taken at a location approximately that designated by line 3-3 in Figure 1, showing concentric cylindrical walls") and, further showing how partial sectors of this annular flow space may be blanked off by end plates 2|, leaving selected passages 22, defined by plate walls 23, 24, 25, whereby the coolant air may be directed to selected portions of the nozzle ring. These walls also define spaces 26, external to the air manifold, through which conduits 21 (see Figure 1), may lead combustion gases from gas manifold l1 (see Figure 1), to the remainder of the nozzle ring.

It is of course understood that the arrangement may vary from one in which alternate nozzles are on coolant and hot gases to arrangements in which several nozzles handle one fluid, followed by several nozzles handling the other. The exact mechanical details for either arrangement may be worked out by those skilled in the art.

It is essential, in order to prevent shock loading of the rotor blades, that the coolant air be admitted to the rotor blades at approximately the same velocity as the combustion gases. As is well known in the; agrt, ,the proportioninggofithe nozzles, in=reiatiomto=entry pressure thereto; determine exit velocity. In order that the nozzle design for coolant air be generally similar to. that for gases, which is obviouslydesirablefon mechanical reasons, it is preferable to raisetiie' pressure of the coolant air toiaslevehapp rpximat -t ing that of the combustion gases: Thentilization; of the second stage compressowservesz-thisipurmpose.

I claim:

1. In a gas turbine possessing a nozzl'eiringt. conduit means for conducting hot gases to alternate nozzles on said nozzle.- ring; atlleastlone -fuel. adapted to{continuously-receive fuel and. air 3T1QFj$0 :blllll;th8;-;I111X13L11B, therein, to produce.. hot gases; agcornpresser. adapted to induct! lldfgtgqdl charge it; under; p6SSL11,-.3 Ild 0QIl:-. du tg mea or conducting said, compressed air: to saighburnen;,thedmproyementrwhichscomprisesr ammond-pornpressor adaptedto compressapore tion oi',the;air;talen-frornithe discharge .of saidv first com ressor,-;means fon'condueting-said pore tiop ofs a -fr.om. said:v compressor tosaid-- second-compressor, conduit meansv forconducting said air from said seconndcompressor, a manie dz nn cted o the-enact. idz nd tv n aa wtareae' tri a d tv iw ompress d r d: direct-said: air-r ten alternate; nozzles on said: noz-z ie m not fifish l i -pa sae eei dri. i ZZQROtOLTIQQHD-tfidfffifl rota-tioxr upon-,san a s wrh ne nas aecsaa ptsdi o on n ou y e eaiternatelzz a d-twelan a irnd said hot? from said nozzles on said nozzle;ring.;;, i

2%: Imae as5turbine;sett mrisinse i wo s ag aimoommessorlat least? one: ,fuehburimm and: a: turbinesequippedzwitha aenozzle; ring; the improye ment which comprises means fonconductin aaz partioneofithe ainfmm thesfirststaaeaofzthezcom pressor: to tha fuel:.-burner;-, saidr-burnen adaptedl tozburna'fnel-ir-rzsaid; air ;to:produce; a hot-Working 1 gas;v means-lion.- conducting. saidshot: working; gas: toe. alternate nozzles ion; means: for conducting: the remaining from: theafirst: stage? oi--theacompressor to: the second: stage; and means-riorxconducting theicompresseda air-fromzthe:seoondzstage toaiternatea-nezzlessonz theinozzleudngcnotaused? Working:- gas;- theasecond stage oft/the; compressor: being designed to impart to theacompressediiair snfficientzenergy-: to; produce a dischargevelocity; from the" nozzles substantially: equal to i the. discharge velocity from: thenozzles of? the. hot work-ingrgas.

3: 111a gas turbine system, agas-turbine includin'g a rotorhavingradial peripheral blades, and adjacent circumferential stationary nozzles; a-- combustor; afirst-compressor-for supplying com tor the tiansfer: hot: 0

the; turbine nozzle;- ring-,- to 1 1 forssupplyjng heatedzrgasesjo selectedg-sta-iib mlfy tor compresseel lieated gases; and means including a-second compressor for supplying a coolant, at; a p ressure, excess of said compressed air,

to another sector of said stationary nozzles, said 'pressure and nozzle capacities being such as to secure equal velocities at the nozzles of both nozzle sectors; 7 v

:.IInaieas turbin sls i a es urb nenclud ins. airotoit; radial-pe ipheral blad s andadlace. circumferential stationary nozzles, a combust nozzles; a; first; compressor for; supplyingazw pressed fair. to; s aid: combustor; and meansinellld ing. a. second; compressor for suppl inacoolant; under;v pressureto; other selected; stationary; nozzles the: pressure; ctr-the; coolant; being; such; and;

the coolant nozzle capacities; being; sucmasc 01 secure .equalxgas welocities; at. exitfrom -thejnozzles-g of both gioupss. Y

6:: In; a. gas: turbine system, aig-as turbine ineluding; a: rotor, radial; peripheral), blades: omsaidz;

rotor andra. plurality of? circumferentiafllispacedi stationary; nozzles:- adjacent said 1 blade on the; upstream side-thereof; a combustor; forrsupplyingg heated: gases; to selectedrsttttionam; nozzles; t

compressor for supplying compressed airrto saigii;

combustor, and means; including a sec r d cornpressor for supplying; coolant ,under1=pressure to: othen electedz ationaryinozzle ES' RTG IQH -I I References Cited. in; the file-0f.- thiS patents UNITED STATESwBATENTS-i N m er Name. t

, 9 9,895 Eemale Apr. 27;,1909

926,157" Weiss June-2,9; 1909: 1,375,931 Rateau Apr; 2151921 1,42 987 hn onmifigfln une izflsz 1;7 ,957 d m ir c. 1.1 29. 39 89. 9; Wh t e u 13; 14 I 2, 39,683" Stalker ,Nov. 29%,1949- as entane: n 2, 95

seamen PATENTS Number Country Date 1959 20 Great Britain; Octi afloat- 4845289 Great Britain Ma s; 1-938 346,611 France Dec. 3; 1964 371,230 France; Jan. region-"1 641,739 France Apr. 2.321928

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