Preliminary analysis of effects of air cooling turbine blades on turbojet-engine performance

RESEARCH MEMORANDUM

PRELIMINARY ANALYSIS OF EFFECTS OF AIR COOLTNG TURBINE .

BLADES ON TURBOJET-ENGINE PERFORMANCE By Wilson B. Schramm, Alfred J. Nachtigall and Vernon L. Arne Lewis Flight Propulsion Laboratory Cleveland, Ohio Aufhurify c-m

NATIONAL ADVISORY COMMITTEE

FOR AERONAUTICS

BlAms OI? TlBBomr-mm -a By Wileon B. S&ram, Alfred J. Hachtigall and Bernon L. Arne The effects of turbfne-blade cooling on engine performam were analytically investigated for a turbojet angine in which cooling air !Fhe amlysis was based ti the ia bled from the engine compressor.

measured performance aud 0pePating cmditiona of a -ical nacooled axial-flow turbojet engine in order to obtain a fixed basis for corn-- parison between operation with two cooled-blade canfigurations and the performme of the uucooled engine.

The calculations weremadefora perature, engine speed, and flight Mach number over a range of altitudes. The analpis considered the effects of heat transfer on coolest requirements, blade-coolant-passage pressure drop, and tur- bine performance as well as the effects of heat loss, compressor bleed, coolant pumping requiremente, andother factors onengine perfortmnce. As a part of the,analysie, the 7nlnimm cooltng require- ments permitting stzbatituticm of ncrrrstrategic metals in turbiae bladlng and the desirable characteristics of hi@-tezqperature turbo- .

jet engines were also considered.

The results indicated that, fora constantturbine-inlettem- perature and en&e speed, air cooling of the turbine blades increased the specific fuel consumption and decreased the thrust of the engine.

For a given coolant flow, the percentage increase in specific fuel consumption was less +&an the percer?tage decrease in thrust, relative to the uncooled engine. The reqnfred coolant-flov ratio, deftied as the ratio of coolant weight flow to compreseor*ir weight flow, increased with altitude. Because of the high coolant-supply pressure required, it was generally necessary to bleed the cooling air from the compressor.

The highest possible cooling effectiveness was desirable to minimize the coolant weight.flow aad its effects cm engine performance.

mc4RME5m22 The analysis indicated that appreciable reduction in blade+loy content laay be feasible with some sacrifice In over-all engine performance. The application of cooled turbines to permit operation at high turbinebinlet temperature w&th appropriate modifications in desigu of other components offers possibilities of improvement in performance at both mximnm-power and maximum-range conditions and improved flexibility In engine and aircraft operation.

IWJSODUCTION Theoretical studies of turbojet-engine cycles indicate that substantialgaFns inaircraftperforms.uce canbe obtainedvith engine operationatgastemperatures thatare beyond the temperature limi- tations of current uncooled turbines. It is evident that the high- temperature turbine requires cooling to permit operation with currentlyavailsble materials,elthou& the engine performance is diminished by the effects of cooling losees. Another use of cooled turbines is to provide satisfactory endumuce at current gas tem- pwetures, u3thlow-temperature materiels havinga minimum content of strategic alloying elements that could become inaccessible in an emergency.

The previous analyses of turbine cooling, such es the blade- cooling studies reported in reference 1, were made for the purpose of determining which blade configurations had the necessary cooling capacity to offer possibility of substantial increases in gas tern- perature or reduction in strategic-metals usage. Iu reference 1 it is recognized that the cooling losses might alter the relative ~O~W~ison of cohfigurations snd the erteut of desirable increase in gas temperature or reduction in use of strategic metals.

Au analysis of the effects of air cooling on engine performance atcurrentgas temperaturewas msdes.ttheNACA Lewis laboratory. The purpose of this report 16 to present results illustrating the relative mgaitudes of the effects of air cooling for two blade configurations and several possible blade metals. The future potential performsme snd desigc requirements of improved air-cooled turbojet engines are also considered.

III order to make e detailed analysis of theee effects, know- ledge of the turbine geometry and operating couditions is essential in evaluating heat-transfer charscteristics under various flight conditions. Consequently; most of theanalysispresentedin this report cousiders only applicaticm of an air-cooling system to a HACA.RME3aE22 r typical turbojet engiae, for which the uncooled performance and the The use of a given operating conditione are known from engine tests.

engine as a Isis for the analysis has distinct advantages in com- parison with an ordinary cycle analysis because it permits study of altitude effects and the effects of coolAng on compressor and turbine operation. This method also provides a fired basis for comparison between the cooled and the uncooled engine and is generally more realistic than the simplified cycle analysis.

Some assumptions, which affect the siguificance of the final results,are required in the cooli.nganalysisalthough the methods With any specified allowable blade- Include most of the variables.

metal temperature distribution, the analysis provides good c~tive evaluation of air-cooling systems. The greatest uncertainties are found in the specification of the allowable metal-temperature dis- tribution in the blade as a functicol of the alloy content sud physical properties of the metal, and the behavior of the metal under the uncertain temperature and stress conditions prevailing in the actual engine.

The proposed turbine coufiguration thathas beenusedas a basis for the present analysis is IJLustrated in figure 1. The figure shows a half section of the turbine rotor snd casing, and The combustiou gases are in*- part of the jet-nozzle essembly.

duced from the left side of the figure into the turbine stator and rotor bladiag snd subsequently pass downstream into the jet nozzle- The rotor-blade cooling air, bled from the compressor, is intro- duced through a supply pipe to the stationary diffusing section at theturbinehubandpasses intoavaned shroud (on thedownstream face of the wheel), which carries the air to the blade base. The cooling air then passes through the hollow rotor blade and is dis- charged at the tip to mix with the r~~~i.u flow of working fluid. In this analysis, the stator blades are also considered to be cooled with compressor air introduced through sn external lpanifold. The air passes down theleadzIngportionof the statorbladesnd is discharged from small holes distributed over the trailing portion of the blade to achieve a film-coolfng effect.

Tvotypes ofrotorbladeare considered inthe analysis,both of which incorporate spdcial modifications to improve control of blade temperature. The two blade-coolant-passage configurations are illus-trated in figure 2.

4 XACA RM ESOE22 outline of AnalyBiB The outline of the snalysie logical falls into three nmL3.n steps: First, the coolant flow required at any given set of con- ditiona is established; second, the flow cmditions of pressure and temperature through the bladepassageandrotor supplyshroudare determined in order to establish the required supply pressure and ten@erature of the air at the hub of the turbine rotor; and finally, the ultimate effect of the coolAng on Jet-nozzle conditiona, engine thrust, and specific fuel conslnQtion are oopputed~ These steps are considered in greayter detail in the following sections.

Determination of coolant-flow requirementa. - The blade-metal temperature distribution 18 pri?mwCl.y dependent on the radial gas- temperature distrlbutlm, thi oool&M&t temperature at the blade root, the blade-profile heat-transfer coefficient, the coolant- passage heat-transfer coefficient, and the configuration of the coolant passage.

These variables can be expressed in equations within limitations that are briefly discussed later in the report.

In thie analysis, the coolant +.emperature at the blade root is arbitrarily specified at a constant value for all calculatims inamnuch aa the air teqpsrature at the point of compressor bleed is unknown in the first stage of the analysis. It is therefore assumed that the temperature of the coolingair Buppliedattheturbine hub can be contmlled by a heat exchanger so that the specified tem- .

perature at the blade root is obtae- The praotical necessity of providing thieair-Waperature con4zol is subsequently discussed.

In detsrminiq the required coolant flow, the known engine operating conditionsforany selectedflightocmditionare used to evaluate the blade Reynolds nutliber and the outside blade-profile heat-transfer coefficient. For a specified coolant-passage configuration snd allowable n&al te~rature distribution, it is possible to determine the weightflowof coolcLntrequAred,atthe specified turbine-inlet gas temgeratureand the specified inletcoolantte~erature at the blade root, with'ccnsiderstion of the effects of heat trsnefer' and oentrifugSlcompre8Blonontl2e cooling-air temperatureas it passes throughtheblade. Essentially the process is a trail-and-error Boluticn fm the passage heat-transfer coefficient that yields the desired metal-temperature distribution. me rotor coolant weight flow for any set of cmditions can then be nondinenBiomOly ergreseed by dividing by the total compceesor-edr weight flow that uas Fnlthllyumd inevalu&ingthebladeReynoldsnumberandheat- -&ulsfer coefficient* This ratio, deeiphed the coolant-flow ratio, can then be evaluated for a range of flight conditions inasmuch as the variatims of engine-*8 weight flow and compressor pressure ratio, which affect the heat-transfer coefficients, are kaown from tests.

KAcARME5m22 5 I The stator-blade coolant-flow requirements for this partioular analysis were estimated frcm limited data for film-cooled blades presented in reference 2. A fixed stator ooolant-flow ratio of 0.01 appeared adequate for all conditions. No allowance w&a made for cooling of such components as the turbine casing and. jet-nozzle assembly.

AIIB~JB~B of flow conditions in cooling passage- -After the coolant-flow requirements for the rotor blade have been evaluated, theflo~c~~~diti~8 inthe ooolLngpaasage ca~~beani%lyZSdtO deter- mine the Kach number distribution in the passage and, BubBgq~tly~ the required pressure at the blade root. The analy818 of Mxch mm- her distribution is mde according to methods given in reference 3, and takes into aCCOUnt the effects of centrifu& compreBBiozl, friction ~OBS~B, momentum change due to heat transfer, and changes in flow Bxea.. Either a pressure rise or a pressure drop may occur from blade root to tip depending op which of these effects predominates. The rise in pressure and tempemture of the oooling air from rotor hub to blade rOOt is then used to evaluate the external supply pressure requiredtopuzlqp the coolingairand the supply-air temperature per- miBBible at the turbine-rotor hub to rmintain the specified cooling- air temperature at the blade root.

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Cooled-turbine operating conditions and performance. - Up to this point in the agaly818, the Imown operating conditions of the uncooled engine are us& ae a first approximation to determine COOhILt-flOW and COOkGIt-SUpply ConditimB . New turbine Operating conditions are then computed for the cooled englue and the effect on engine thrust aud specific fuel COnBtrmIptiOn iB thereby estimated.

The cOOli.ng-&fr-Supply pressure that has been determined fixes the bleed point on the engine compressor for the rotor coolant. The stator coolant is bled from the oolqpressor discharge.

The power requiredof the turbtie expansion, including the extema lPtlm37iw through thelzlaincorftpressorand the internalpmp3ngthrough the shroud and blade passage, can then be determined. This power must be extracted f'rcmthereducedwei~tflowofworkingfluidavailable to the turbine; therefore the specific turbineworkie increetsed.

This increased specific turbine work and the heat transferred from theworkingfluid to the coolantareused indeterminingthenew turbine-dischazge cmditions. The mixingof therotor coolantwith them3lnxorkLngfluidfurther reduces the temperature downstream of the turbine. The net effect of cooling is a reduction in the jet-nozzle total temperetture and pressure ratio. The effect of bleeding the cou@ressor to obtain a--for cabin pressurization end IIACA RM ESCIE22 oonditiontig, auxiUaries, or deicing is analyzed in reference 4.

iS not applicable to this The method of reference 4, however, etnalpfs because it does not consider reintroducticm of the bleed air into the jet nozzle.

ABBUU@dOIlS Mxny aasunrptions are neceesary in order to make this analysis, although it is believed that the significant vsziables are .cou- The assumptions may be ditided into two classes, those sidered.

used in the development of the analytical methods, aud those used in this particular application of the mSthOd8.

ABBUiZptiOIlB in deYelOpIEtXIt Of S.nd.yBiS. - The most Bi@ifiCEUlt assumptions made in the developmeut of the methods are as fOuoW8: A one-Sdimensional blade-temperature distiibution is used in which blade temperature at any radial DTOBB section is constant.

Experimentsbave shown thatthetemperatures of the leadingand trailing edges of cooled blades sre considerably higher than the temper'ature of themainportion of cooled blades similar to the blades of figure 2, which do not have special provision for leading- and~trailing-edge cooling. The leading- and trailing-edge tem- perature gradients are not considered in determining the required .

coolant flow, and the blade temperatures obtained are therefore average temperatures over the midchord section of the blade. The analy818 makes use of a Biqplified form of the one-dimensiom~ equation, the derivation of which requires the additlonal assump- tions that the @a-temperature profile is uniform, that the blade- geometry and inside and outside heat-transfer coefficients have a meanvalueover the span of the bl.ade,and that the radiationand conduction effects can be neglected.

AS shown in reference 3, t&e one-dimensional ~1~~1~ can be used to consider the effects of the radial variation of inBIde and outeide heat-transfer coefficients, bladegeometry,andradial turbine-inlet g48-temperature prOfile, which is the strongest influence in the blade-temperature distribution. !Phe solution with radial variations, however, requires the use of tedious numerioal- integration procedures and for the comparative results desired in this &II&~~B~B the additional accuracy that could be obtained was unjuBtified.

AS shown in reference 5, neglect of radial conduction to the blade root generally introduces a negligible error in computing NAc%RMEX!E22 7 blade-temperature distributim and coolant-flow requirements for blade dimensions such as used b these caloulaticms. Neglect of radiation to the blade introduces some deviation in blade temper- atures but has little effect cn comparative results. The SIEd.~BeB in refftrf3ILCe 5 indicate that at a &as tmrperature of lSoO" F, neglect of the pre+do&mn t nozzle radiation msy result In a uniform error of*2!Zi0 F in the spanwise blade-temperature profile when the turbine nozzles are unoooled- (See fig. 8, reference 5.)

The effect of assuming constant geometry, heat-transfer COeffiCieZY&, and gas properties iB illustretted in figure 3 for the finned-blade passage shown in figure 2. A typical radial terqper- ature profile of the gas at the turbfne inlet sud it8 effect on the me-dimensional radial blade-temperature distribution computed for different assuzqtions is presented in figure 3. These results are taken from reference 3. The dot-dash curve represents the case inwtichradialvariationof internaland extermlgeomsizyt relative gzs velocity, internal and external heat-&?&Bfer coeffi- cients, and physical gas properties are considered. AssmqItion of constant mean values for all VSrtib~s except gas temperature results in the blade-temperature profile @Ten by the dashed line. The com- puted tezaperaturesare conSiderably higher near the tip of the blade but are comparable for radial blade stations between 0.2 end 0.6.

Roth temperature diatributionS are for a rotor coolant-flow ratio of 0.02, and in both cases the blade-metal temperature distribution * resembles the ~a-t8Ecpsrature profile. The effects of changes in shape in gEla-temperature profile are diSCUSBed later inasmuch as it fs a design condition that may be imposed.

.

The physical significance of the aBBwr@tion of me-dfmeneional blade-temperature distribution is that the blade configu~~tionwould hgLve to be modified to aUOW, hBOf&Tl? a8 iB ROBBible, UUifOY'm COOlfng of the profile, at the 8ame time retaining the essential performauce and flow characteristics of the Original uncooled turbines.

Analysis of the effects of heat transfer, friction, CentiifUgal co~ression, and geometry in the blade passage is based on one- dimensional equations given in reference 3. It is believed that with the B~u. PaSSageS that OCcuZ’ h Cooled blades eBBe&ially OIle- dimensional flow tiu. &St. Reference 3 &so BhOWBthat Bohtian of the equation8 for Mach nuaiber and preBB~e.diS~ibUtiOn is practi- cally unaffected by the use of constant heat-tramfer coefficients and unirorm turbtie-inlet gas-temperature profile, as was aBBUIEed in this aIla&yBiB. In determining the pasSage-flow chsracteristics, the c static pressure that prevails Inside the passage at -the blade-tip section is assumed equal to the turbine-discharge total pressure and NAcARMFmE22 .

the entire velocity head of the cooling air as it emerges from the tip i6 ~BBlRWd lost- In reference 6, it is shown that a h??ge pm?t # of the kinetic energy of the cooling air can be recovered as reaction if the blade tips are especially designed for recovery.

III the WiS Of cool- lOB8eB ernd their effect oil angfne The statorand operation, several additional ass~tions are mde.

rotor cooling air are not.considered as part of the turbine working I-J fluid but are considered as part of the Jet-nozzle gss weight flow.

It iS &SO assumed that the colnpreeBor operating point iB UIIaffeCted by the interstage bleed for coolingairand that the statorand rotor blading for the cooled engine csn be redesigned in order to maintain matched compressor operation and to extract the higher specific work frcm the reduced wei@A flow of workkg fluid available.

AssunQtiona in application of analytical methods. - In the appliCatiOn of the analytical ?.nethods previotmly outlined for deter- mining the coolant-flow requirements, additional assumptions were mde that have no bearing on the melAmd8 or equations but affect the numerical results. It is currently necessaxy to use blade- profile heat-transfer correlatious that have been obtained with ststic CELB~~~~B of blades and coolant-passage heat-transfer corre- &3tiCXlB that amu for hIbeE- In both cases the effect of cen- trifugal forces on the heat-transfer coefficient is neglected, and for this reason it is expected that the numerical results may be .

Scnnewhat OptimfBtiC. Re&q3XdhBB of absolute ma&tude, however, the use of such correlations introduces Repold number effects into the cycle c-a~ctIhtionB and lpakes the computed engine variables BeIIBitiVe to the main parameter8 that affect heat tX%IIBfer. It is believed that the trends that are indicated in the results will not be seriously affected when more applicable he&t-transfer data become available.

Except where otherwise noted, a flight M%h nm&er of 0.788, a unifQln turbine-inlet telggerature of 1500° F, an engine speed of 7600 rpmr snd g relative total inlet cooling-air temperature of 300° F are used throughout the ~~18. An allowance ~a8 made for heating of the cooling air as it passed radially outward along the face of the turbine wheel and the required amount of intercooling between compressor and turbine to control the cooling-air teqer- aturewas computed. No extra ~oBseB due to pressure drop in the inb3rCOOhW were considered and it W&B aB8tUUed that the heat x&a rejected to the fuel; thus no calculation of momentum losses due to supplying ram air for the heat exchange were necessary in the analysie of over-all engine perforusmce.

NxARME5cE22 9 c The control of blade-temperature distribution is dependent ou a nuuiber of other factors, which influence the initial. design of the engineandthe blading. Some of these factors are the geometric the radial temperature profile configurationof the coolant passage, of the gss, and the existence of local temperature gradients tithe profile. The namer in which these factors enter the setup and the evalUaticm of the analytic is indicated in the foXLoMng secticm: Significance of blade-cooling-passage configmation~ - The coolant-passage configuration influences the coolant-flow require- ments and the pressure drop or pressure rise of the cooling-air flow as it paBseB thrOU& the blade. Ram air is generally not at sufficient pressure to supply the required cooling air flow; the engine colnpressor must therefore be bled, tiich fn turn complicates the design and affects engine operatim.

Three types of air-cooled blade in order of increasing effec- tiveness axe: the-&FnhoIlcmblade, thehollowbladewithinsert, The effec- and thehollowbladetithftie inthe coolantpamage.

tiveness is defined 88 the difference between effective &%a km- perature andblade-metal temperature ditided by the difference between effective @LB temperature and inlet coolant temperature at the blade root. Thus the effectiveness of a cooled blade is a c measure of the metal-temperature reductfon achieved by the cooled blade for a given coolant flow, relative to the temperature reduction 4 ideally poEBible for given gas and COOht teITQeratLU%EI- The effectiveness is limited by the magnitude of the heat-transfer coefficients and the length of the conductive path through the metal.

It can be seen that the effeCtiVeIMB8 ratio for a given blade-cooling-petssage confY.guratiou also expresses the amomt by which the effective -8 temperature can exceed the rue-1 temperature for a given coolant flow and coolant temperature. Analyses have shown that the effectiveness of the plain hollow blade is very poor, which limits the deBi@ to more effective configurations such as the insert and finned blades illustzated in figure 2.

In the insert blade, the effectiveness is augmented by restricting the coolant to the amulsr passage immediately adjacent to the inside surface of the blade. The improved effectiveness of the finned blade is largely achieved by the increased surface area exposed to the coolant and the conduction of heat to these areas by means of the fins. When the finnedbladeand the insert blade are coqpared, ina given case essantislly the same amountofheat- transfer and metal-temperature reduction can be accom&ished. but NA,cA RME5m22 with coneidembly less weight flow of cooling air with the finned blade. In the finned blade, however, the air-temperature rise is mmemtum pressure 10~8 of the more rapid snd, consequently, the cooling air is greater. In most cases the friction 108808 are also higher in a finned blade, and thus the higher effectiveness of the finned blade is offset to some extent by the gJ?eatir coo&mt supply pressure that must be maintained. The net effect can be evaluated cmly by the complete engine w~L~JB~B, which reflects the effects of compresBor bleed snd ping power. A further cmsideratiou is theneed to congrromise betweenthe co&@erityofthebladeshaving higher effectiveness and their producibility. In nuny cases, however, L l.tlUit OtiBtS OBL the SlLlOIRlt Of tiEBe that CBP be SUB- Wed, and because of losBes the amount of cool.antflowmuBtbe mipimized- It is therefore necessary to develop fabrication methods for the More complex blade types so that the marimum effectiveness posaib~ can be attained.

Effect of gas-temperature profile- -The effectofradial turbine-inlet-se-tewerature profile on blade-metal-temperature distrilwticm is shown in figure 4.

The typicalggm-temperature profile is taken from figure 3. !f!he uniform @3B--&m.!p3ratU.re profile shown in figure 4 iB an integrated UleEul of the typical @LB- temperature profile. Anallowable blade-metal-temerature distri- butian based on stress-rupture data for a Cr-Mo-Va steel blade is also given.

The blade-temperature distribution6 for the typical gas- teaprature profile and the uuiformgas-teqp3rature profile are matahed to the allowable curve so that the blade-metal temperature does not exceed the allowable blade-metal temperature at any point.

The coolant-flow ratio required for the uniform gas-temperature profile is 0.0145, whereas that required for the typical profile is 0.020.

The uniform profile appears to be more desirable because the coolant flow required is lower and the blade is not overuooled, over a large part of the span, to the extentthatoccurs when the gas-temperature profile varies widely.

meet of temperature gradients in blade profile. - The possible Ipa@tude of temperature gradients in the trailingedge of the finned- blade profile is shown in figure 5. The BpEUIW%Be distribution Of trailing-edge temperature is comparedwiththe gas-temperature profile and midchard-aectim-temperature dis-&ibutim from figure 3. For the typical gas-temperature profile, the tmiling-edge temperature reaches 1560°Pas oompared#Ithamarimumtemperature of1230° Fat other pZ%B of the profile where cooling is more directly applied. The u NACA RM EsoE22 of the lmg conduction fs a result pathfromthe trailingedge to the cooling-airpas~age~ llhe desimbility Of blades with BhOrt trailing -8 iS apPareI& CarefxiL deal@ is required to r&i&X&e hot spots; and ti the Cm- promiaebetweenblade-profile aerodynadc~aad heatkaasfer, reduction 3.n required coolant-flow and trailTng-edge temperature by means of a blade shape that favors coolzIng is belleyed to justify 801118 loss In turbine efficiency i&at will probably results The most aerodynamicallyefficientturbine doesnotnecesssxilyresultinthe highest over-all engine efficiency.

Determlnatiou of Allowable Blade-!E!mqmrature Dietributian Another iqortantvariable,the dete3?mJnatimofwhichdependB Qp & number Of ~BBU3@tiCElB, is the allowable blade-metal-temperature Typioal factma that should distribution for a desired blade life- enter this determina t&mare the combined stresses due to centrifugal forces, gas bending forces, vibratory excitation, and the stmngth uader these COnibfned StrsBBeB.

properties of the blade materials The influence of them& cycles and corroeion effects should also be correlated,andalltheaevariables shouldbe systematical&related to thealloy~ontentof the blade~terials,asweUas to thelIfe of the blade. No satisfactory method has as yet been devised whereby all these factors can be taken into acoouut.

Specification of allouable tempaxature. - It is Cllrr@L~ necessarytorelateblade ufe to operating B~sSSandteI.Up8?XLttEe by means of the BtresB-rupture charsC+eriBtiCB of the zrraterials, whichvarywidelywiththealloyoontent. The sim@estcentrifugal B~eSB iS Used. as the Cdf23riOZL in determung an allowable temper- ature distribution. The asBu7DgtionB used to determine the allowable blade-metal temperature hsve no effect an the methods of WB~B used to determine the reqafred coolant flaa far any specified allowable temperature. The effect of a given coolant flow on engkne per- formance is comlpletely independent of the factme that enter the determ3nationofauallowable temperature. The use of rupture criterion specifies the upper limLt of blade lffe and the lower limit of required coolant flow for any selected ccmdition. Actual blade lifewillproberblybe~~ehorterbeoetuse ofothm recognized influences suuh as fatigue, therm1 effects, aud corrosion phenomena, which at present must be evaluated from experience.

The rupture properties of representative alloyat which are used todefine theallowable blade temperature,areiUuatitedin fig- ure 6, which presents allowable stress a@nst temperature for a 12 NACA Rbl Emi22 .

rupture life of 1000 hOmE and the cooling requirements for a repre- sentative engine condition. Of the alloys shown, the highestpro- .

perties are developed by S-816, a hi&-tenrperature &Lloy that contains approrimtely 96-percent strategic metals and about 4-percent iron, which is suitable for uncooled operation at 1500° F at the stress level of about 16,500 pounds per sqmre Inch shown by the dot-dash K line. This BiLFeBS level iB arbitrarily Bet in figure 6 for the i+ purpose of comparing other nraterials with S-816 operating uncooled at1500°F on the basis oflOOO-hourrupture life. Atthelowend Of the a=Oy S&e iB %g 1015, which iB BUbBkaIItially loo-perCeIIt iron; for this stress level, SAE 1015 has an allowable temperature of about 970' F, which Is more thsn 500’ lower than the uncooled blade. As a first Step in redUCtiOnOf strategic-met&Content, an intermediate group of alloys containing 50 percent orsmore iron and eliminating the IllOat B-tsategiC metals , columbium and cobalt, may be oonsidered. Typical of these alloys is the 16-25-6 aJ.loy, which, for the specified BtreBB level of 16,508 pounds per square inch, has an allowable temperature of about 1380' F +nd requires lees than 125°Freductianintemperature belowthe uncooled blade of S-816.

In the range betweenSAElOl.5 and the16-25-6alloy, thereare many ferrous alloys that develop remarkable properties with very smalladdition~ of strategic alloyLug elements and thus offer possibilities of substantial savings in st2categic elements. A tygical alloy in this rsnge is Cr-Mo-Va steel having lea8 than 5-percentalloy ccmtentandanallowable temperatureaxouud llOO°F at the stress level of 16,500 pounds per square inch. Thepro- pWt%38 Of this Steel are superior to lpazly highly alloyed BtaiIihSB steelsattemperatures up to llOO°F. New developments such as the onepointshowntifigure 6 for the Ti-Bo (titanium-bnron)ateel Indicate that furtheriqprovements maybe expected. This particular steel has 2.25-peroent chromium and l-percent mo4bdenum, with additions of 0.30-peraent titanium and 0.0%percent boron.

Limitations in allowable blade temperature. - The significance of diminishing returns in the coolfng prooeBB is app?DeI& from the lower curve in figure 6, which iUustrrrtes the rotor coolant-flow rrrtios required for the arbitxary stzesB level starting with boo F as the uncooled-blade temperature. The coolant-flow ratios are plotted at the midspan of a seven- finned blade assuming a constant value of thermal conductivity for allmetalsshoun. The vertiCa1 dotted lines indicate the allowable tenrperatures for loOO-hour rupture life for the various blade InStSrhd.8 at a BtreBB Of 16,500 POMdS per square inch. !l!he COO&r&- .

flow requirement increases rapid4 with temperature reduction below l3 NAcARMEsOE22 For example, the 100' F decrement from that of the uncooled blade.

l 1500° to1400°Fis obtainedwithan incrementofapprorimately 0.0025 in coolant-flow ratfo; whereas the 100° F decrement from .

1000° to 900°Brequires a corresponding increment of 0.0135 2n coolant-flow ratio. The cools&flow requiredtoachieve the 4U0°F' reducticm in metal te~erature for the Or-MO-V& steel is a little id more than one-h&f that required for the SAX 1015 steel. 'Ilhus, elimUating the remaining 8mUamount of alloying elements results in approximately tuice the cools&-flow requfremen~s for a fixed turbine-inlet temperature of 1500° F,anditis evident thata com- promise is necessary in sstablishi~g a perm%ssible alloy content for the nonstrategic-alloy blade- The low coolant-flow ratios that are illustzated in the lower curveapplyto8nactualengTne oulyfordesigo conditions that fulfill those of the calculation. These low coolant-flow ratios could be obtained in prsctice ouly for ideal conditions of es- teqerature distribution and blade stresses and for a modified blade profile end petssage configumtion that provide a favorable blade- temperature distribution approachiagthe one4.imeusional caseD In t&Is senses each petit on the ewe represents & design point and is the minimum coolant-flow ratio tit could occur under the specified operating conditions of the turbine. SiWlurly, the desigu conditions that result in these &nLmm coolant flows represent eb goal or objective that should be considered in the AnItAal design of a cooled turbine.

Ih the results subsequently given, varfous ewes are identified bythealloythatwas used to specifytheallowable mtalteqmrature; it is necessary to recall that practical realization of a turbine design using the indicated coolant-flow ratios and metals is subject to the limitations discussed throu&out the section AEALYEIIS.

The methods and the cclnsideratfcms previous& outlined have been applied to c-0 the effectiveness of a fkmd and an insert blade, similar to those shom In figure 2, using the substitution of nonstrategic blade alloys as a means of selectJug the allowable blade- t.emDerature distribution. The typicalresults givenv to thispotit for the fimmd blade show effects that occur at some constant spec-' ified flight condition. Over a range of flight conditims, altitude andflightmchnumber effectsare encounteredbecause the bladeheat- tzansfer coefffcients are dependent cm the variations in engine-n~,&&s weight flow and pressure ratio that occur with altitude and flight epeed.

14 mLcARMR5OR22 .

Coolant-Flow Requirementa The coolant-flowratios reauiredfor the finnedand insert blade of nasstrategic Cr-Mo-Va steel&e presented in figure 7(a)- The coolant-flow ratio for both finned and insert blades increases vith altitude, for example from 0.0105 at sen level to 0.018 at the 40,000-foot altitude for the finned blade. The lower effectiveness of the insert blade is shown by the substantially larger coolant- flow ratios, almost twice that of the finned blade, throughout the &ltitu&9 range. Finnedbladeswith smiU.er finthichsessand spacing than-the blade used In thisaualysis havebeenmde for experimental purposes and have greater theoretical cooling effectiveness than the blade considered in this report. Thebhschnumberthrough the coolant passage also increases with coolant-flow ratio aud altitude and both of these factors teud to increase the coolant-supply pressure required for the coolLug air. A critical altitude might be defined far a given .blade where the required coolant supply pressure reaches the ccqpressor- discharge pressurej although this l&.aitation could be relieved with an auxiliary compressor. Bladqs with high effectiveness and with convergent cooling passages tsnd to have a higher pressure drop t&m blades of low heat-transfer characteristics and constaut-srea pet~B&~S. It is therefore unwise to esteLblish the mechanlcal end aerodynamic design of the blading without considering passage con- figuration inasmuchas pressure limitations maybe encountered later when attemptiq to provide adequate cooling. Ho detailed analysis has been made of the potentialities of air cooling at altitudes above 40,000 feet. In general a cooling-air pressure ratio equal to the turbine pressure ratio is always available by bleed from the comEpressor discharge.

EnginePerformnce A comparison of blade coufigurations cn the basis of required coolant-flow ratio alone does not include the adverse effects of the higher pressure &.op in the finnedblade. Inorder tomake this com- parison, it is necessary to complete the engine calculations to deter- mine the thrust and the specific fuelconsuqtion. The comgesiscmof the two configurations on the basis of percentage 1088 in specific engiue thrust at the required coolant-flow ratio is presented in fig- ure 7(b)for the samfli tconditionaaudfor the fixedturbine- P inlet temperature of 1500 F. The finnedbladehas the lower thrust 1088 throughout the altitude range, and at a mximtm altitude of 40,000 feet shows a thrust 1088 of approxim3tel.y 3.5 percent as com- .

pared with 5.5 percent for the insert blade. like percentage loss in Il.5 HAcARME5cE22 d specific thrust does not completely reflect the change In over-all a@ne efficiaacy brought about by the cooling losses inasmuch a.8 the air bled from the c-essor for cooUng is not burned and thus less fuel is used in'the cooled engine. It is, however, an iqportant performanoe pariable that is considered in evaluating cooling- A truer measure of the cooling losses is the percentage increase in specific fuel consuqpticm, which bears a direct relation to the over-allefficiencyof the engine. Thepercentqe incre8ge infuel cmsmxptio.uisp&t4x3dagain8taltitudefclr the two cooling-passage configuratitms in figure 7(c). The specific fuelcmstmrptionis increased significantly with both blades and the net losses uith thefinnedbladeareless thanwith the insertblade.

Theaaaner inwhichthe cooUngvaziable8 influence the thrust and fuel consumption is shown in table I, which presents the sigpificantitems in the coqpmisonemcagthe finnedblade, t&e insertblade, and the uncooledengine.

Although the coolant-pasage pressure ratiot which represents the ga?essure drop of the coolingair, is largerfor thefinnedblade thanfor the insert blade, the external coolantcompreasorhorse- gower is less for the finned blade. This conditionresults framthe emallerweightflowof coolantrequired. IiLkeuise, the internal coolamtpumpingpowerfor thefinnedblade islowerthanfor the insert blade. Lnaddition, theweightflowofwmHngflnidavail- able in the tmbine is higher with the finned blade. The work required per pound of turbine worHn.g fluid conseqmrtly is less with the finned blade, which results in higher jet-nozzle total- temperature and pressure ratios.

Similar comparisonsha.ve beenlilade overawiderange of cca- ditions, and at higher coolant flows more marked differences occur betveenthefinuedandinsertblades. A corqarison with the results of table I is given in the following table for higher required coolant flows, which represent a reduction of approxirately 1000 F in the allowable blade-metal temperature relative to that of table I- The other engine operating conditions are the same&s those of table I.

HAcARMEmE22 Required cool- Righerrequired ELntflcwof cool&Ilt flow (lower allowable table I blade temper- ature) Finned Insert Finned Insert blade blade blade blade Rotor coolant-flow ratio 0.018 0.034 0.030 0.063 Decrease in thrust, percent 3.5 5.5 6.3 Il.0 Increase in specific fuel consuqpticn, percent 0.70 1.l5 2.47 4.07 Fcr either the finned or insert blade, the increase in required coolaut- flowratio results ina substantialincrementin specific fuelccm- suInption. At the higher coolant-flow ratios, the advantage of the finned blade is indicated by the much larger differences in perforrasnce .

As previously mentioned, these results are the minimum coolant- flow ratio and cooling losses that could occur and other factors euch as gas-temperature profile, large local temperature gradients, aud the combined stresses, vhich occur in actual engine operation, would result in substantial increases in coolant flow and cooling losses. This fact reemphasizes the necessity for attempting to r approach the design conditions of gas-Wuperature profile and favorable blade-temperature distributim previously mentioned.

Heat-$Ichanger Requirements The desirability of a heat exchanger between the compressor bleed point and the turbine to control the coolkg-air temperature The emphasia~has previouely been placed on was previously mentioned- coolant-flow ratio; however, the cooling+tti temperature is also of In addition to the saving in cooling-air-flow requirements importance .

permitted by lower inlet cooling-air temperatures, the required coclant- supply pressure is also reduced due tc the decreased cooling-air Mach number in the passage.

The required coolant-flow ratios given herein are all for an assumed relative total cooling-air .temperature of 300° F in the cooli: ' passage at the blade root. The compressor-air temperature at tht bleed point is, in most cases , higher than that allowed by the specified cooling-air temperature of 300° F at the blade root and NACARME3oE22 17 a heat exchanger would be required. For the two configurations con- - sidered herein, the maxim heat dissipehtim is required for the insert blade at sea level and is approzilpgbtely 63,000 Btu per hour.

This dissipation is less than the probable oil-cooling requirement i-2 of the e-0 so that the problem of heat dispo&t. is not critical in supplying cooling air to the blade root at 300° F. Further reduction in cooling-air temperature would be desirable for the reasons previously given.

Effect of Increased Turbine-Inlet TeZQpeZXtur8 an Required Coolant-Flow Ratio and pklglne Performance As turbine-inlet temperature is lacreased in an uncooled turbo- jet engine, both thrust and specific fuel constmptinn increase. The results previouslygivenhave shown thatair cooling of the turbine blades adversely affects the eugine performance- The calculaticms tiere therefare extended for the 4O,OCO-foot-altitudeflightcon- dition to determine coolant-flow requirements and engine perfor- name at turbine-inlet tempekatures of 200C" and 2500° F. The general engine cxmfiguration, engiue speed, compressor-air weight flow, and compressor pressure ratio were held constant0 It was also assumed that the turbine bleLding could be modiffed,to match the compressor for all coolant-flow ratios and turbine-inlet temper- atures. In evalwting the engine performance&t-the 2COO"and 2300°F turbfne-inlet temperature , ituasaasumedtbattiC, coolingairwas bled off at the compressor discharge. Otherwbe, the same assuarp- tiona that were made in the analysis for a turbine-inlet temperature of 1500°Fwtie made for the high-temperature calculatfons. At , the neglect of radiation effects is not as reasonable au assumption as at the gas temperature of X0@ F andwouldmake the calculated c~lantflowslower thanrequiredif radiation were considered; however, this asszmrption has no effect on the perforarancevalues obtainedforagiven coolant flow.

The values of thrust and specific fuel consumption obtained in this snalysis are plotted in figure 8. The thrust is plotted in figure 8(a) as percentage of that of the uucooled engine at a turbine- inlet temperature of 1500° F a&n& total coolant-flow ratio for three turbine-inlet temperatures. Fm eachturbdne-inlettemper- ature, the thrust decreases tith increasing coolant-flow ratio- Turbine-inlet temperature can be increased to overccme this loss; but as a result of this me-temperatxze increase, the coolant-flow requirement for a givenblademetalalso dncreaaes causingaddi- tional cooling losses. This fact is illustrated by the two ewes , NACA RM ESClE22 for finned blades of Cr-Mo-Va steel and SAYS lOl.5. For the Cr-Mo-Va steelbladeata turbtie-inlettenrperature of1500°F, the total, or rotor-plus-statcrr, coolant-flow ratio requirement is 0.028 and the thrust decreases to 96.5 percent of rated t&rust.

To restore the engine to rated thrust requires an increase in tUrbim+iI&t ~U@eELtUre that r88IdtS in an IINrWLSe in required coolant-flow ratio to approximately 0.034, as indicated by the point of intersection of the Cr-MO-V& steel ewe end the ewe of lOO-percent rated thrust. For the blade of SAE 1015, a larger inCre&Se in turbtie-inlet temperature is necessary to restore the thrust and is accompanied by a still larger increase in required coolant-flov ratio.

Lncomlpletelyeliminatingtihe etrategic-aUoycontentinthe blades, for e-10, by substituting Skis 1015 for 8 low-alloy steel, alloyrequirements of other engine parts that are al80 exposed to the hot gases and represent a larger percentage of the total englue weight may be increased as a result of the higher turbine-inlet tenQerature. ~anengine usingncnstrategic~~~~terials, the rotor blades should have the highest alloy ccpltent of any of the parts exposed to the hot @see in order to achieve maximum saving of strategic nIet&ls.

Ata turbine-inlet temperature of 2000°F, anenginewith Cr-Mo-Va steel blades requires a total coolant-flow ratio of 0.075 - and has a net thrust 23 percent greater than the uncooled engine at lSOO" F.

Ataturbine-inlettemature of 2500°Fwith the Cr-Mc-Va steel blade, a total coolant-flow ratio of 0.130 is required and a net-&rust43 percent greater than that of autancooled engine at1500°Fis obtained.

A.plot of specific fuel consumption agafnst total coolant-flow ratio for three gas temperatures is shown in figure 8(b). For an uucooled engine, the specific fuel consumption increases as turbine- inlet-gas temperature is increased. Also, far all three gas tem- peratures, the trend of specific fuel cons~tion is to increase ae coolant-flow ratio is increased. At low gas temperatures, the specific fuel consumption is sensitive to low coolant-flow ratios; whereas at hi&er gas temperatures larger coolant-flow ratios ten be used with- out excessive increaees ia specific fuel cousumption relative to a coolant-flow ratio of 0. Thisrelationprobablyexists because for a given coolant flow the cooling losses are a fmdler percentige of the higher total energy available at higher temperatures.

NAcARME5oE22 The following table gives some results of the c&Lculstions made of e@.ne ~rfOl?73PUZCeat&t~bine-inlettelu.pe~~e of %00'Fwith uncooled turbine blades and with flnaed and insert blades. The cool- ant flows usedwiththefinnedandinsertbladeaare the calculated required coolant flows for Cr-Mo-Va steel blades with lOOO-hour X'tIptIlr8 life. The calculations were made for a flight Mach nu&er of 0.788, an tititude of 40,000 feet, and an engine speed of

-r Tme of blade

-- ~cooled coo3 d .

IcclEiert Required rotor coolant- flow ratio 0 0.120 0.235 Specific turbine work, 72.9 90.5 Btu/lb U-8 Turbine pressure ratio 1.61 1.83 2.14 Jet-nozzle pressure ratio 3.72 3.27 2.80 Jet-nozzle toix~l temperature, OF 2267 2029 1802.

3575 1359 ll.37 Thrust, lb specific fuel ccmuI@ian, lb/lb&r 1.648 1.662 1.724 Comparison with table I shows that the specific turbine work is higher for the cooled engine at a tmbine-inlet temperature of 2500° F thsnsta tempemture of l!XO°Fas a ccsmequence of thehigherpm@.ng power andreduced workingfluidavailable to the tmbdne. Because the turbine-inletter@erature ishigher, however, therequiredturbine pressure ratio is lower and the jet-nozzle temperature 8nd pressure ratio 8x0 hi&r. The net result is a thrust iacrease of 48.2 per- cent far the finned blade and 26.6 percent for the insert blade.

The fuelcczmtqption fs aubsteLnti&lyincreasedover thatobt&nedat a turbFzle-inlet tewerature of 1500° F, but co@mison of the cooled withthe uncooled engine at2500°F shows thatthefuel consumption of the f3mxed blade is slightly greater tMn that obtained for an tmcooled engine at 2500 F at this compressor pressure ratio. The engine with insertblades shows abont4.6 percantincxease in fuel conslnnpeionxhenc~edwi~theuncooledangineatthe same turbine- inlet teraperature.

Intheprevious calculationsforhig#mrturbine-inletterqper- atures, it was assumed that, as turbine-inlet temperatures we increased, the comgressor-airweightflowandpressureratioare held constant and fhattheturbineblading couldbemodified tomatch 20 NACA RM E5oJ522 .

the comgressorand the turbine. For an 8ngine having a fixed turbine- stator Configuration, a certain percentege-of compreisor-air weight .

flcwmustbebledoff inorder toaatchthe turbiue to the coqressor abcve the desieppoint. tlhenever the iii required bleed for cooling purposes is less than that required for F etching, it is desirable to find weans by Which this difference in weight flow can be passed through the turbine. There are two possible solutions. The cwessor pressure ratio could be allowed to increaseasteqperatureie increasedandthe resultlngrecmery ingas densitywouldaid inpassingthe~8ightflowthrough the choked stator.

The limiting value of compressor pessure ratio et a givenrotative speedwouldbereachedatthe point of coqressor surge. The other possible solution is to incorporate an edjustable- angleturbtie eta-tar, suchas proposed inreference 7. This variable- angle turbine stator couldbeadjustedelongwithincreasing turbine- inlet temperature tc maintain nratchlPg b8tW8an co5Qresaor end turbine.

Possibilities of Turbojet Engine The present justification of the turbojet engine in mi.lLtary air~tisprimarflythehigh thrust of the engine. Ipgeneral, further increase in engine size to achieve greater thrust does little to improve 8ngine flexibility. !t!he necessity for establishing the aircraftand engine desiguata fixedhI&-pcwerpointresults in hi& cruieingfuelccnstnqptionbecause theaircraftis forced to fly et an uneconcmicalhigh speed.

The p8rfornmhnce potentialities of uncooled turbojet engines et various desie points is illus-trated in figure 9 (deta from refer- ence l), which presents the specific thrust and specific fuel CCZI- sumptim ELSfunctions of compressor pressure ratio and turbine- inlet temperature- The relative ultimate range of a suitable high- speed-aircraft ccmfiguratiou is cross-plotted 011 the figure.

At any fixed compressor pressure ratio tithin the rauge of ccm- pressor pressure ratios considered, the longest cruising range at subsonic speeds is obtained With the lowest turbine-inlet temperature- However, the fact that there is little margin of excess power available for take-off, climb, end combat umneuvers With such an engine lpay make it iIup~ctical. Figure 9 also shows that much higher specific thrust canbe obta~dathi~turbine-inlet~~erature; butth8marimum .

range of the aircraft, for a given c~essor pressure ratio, is less.

A worthWhile objective would be to evolve a power plant for mdern .

rimA R M IE!mz22 2l high-speed aircraft that would have the desirable features of both high-andlow-temperature operation, therebyincreasingboththe mximum-powerperf ormanceandtheamtximum-age performmce.

This objective could be realized with an engine design that permits operation at cmxataatcompressor preemure ratio and weight flow regardless of turbine-inlet temperature and that embodies a turbine-cooling system that permits selection of any turbine-inlet teruperature within an adequate rsage. The essentisJ. design features are a compressor haviag high peak pressure ratio, for example 16, andanadjus~bleblade-angle turbiae stator t&tpermitsIllatching of the componentsatanyttzrbine-lnlette~e~~~thinthe desired range without difficulties from compressor-surge char&c- teristics. The adjustable tmblne stator provides a varfable flcu areabetweenthe cmqressorandthe turbineand inconjmctiohwith a variable-area tail pipe gives the turbojet engine many of the desirable cruise-ccmtiol chamcteristica of the reciprocating air- craft engine, dependingcmthe effectiveness of thevariable-area stator.

If such an engine were desigued for operation at a maximum turbine-inlet temperature of 2340° F and a compressor pressure ratio of 16, the performance at all design points shcwn by the dashed lfne in figure 9 forepressureratio of16wouldbe amilable by proper adjustment of the variable-angle turbine stator. The mximum specific thrustwouldbe in the order of 70 percent greater thanthatof currentengineswfth considerable inqnzbvemsnt~specific fuel-- Bl.UQtiCZL. W iththe use of thesdjustableblade-angleturbine stator, the-turbine-inlet-temperature could bereducedfor cruisingsta lower power with approxi~~~tely 23 percent further reduction in specific fuel consmption. The fael ccnsurqptionatthe cruisingcon- dition would be reduced epproxizmbely 33 percent relative to engines using a covqressor pressure ratio of 4 and a turbine-inlet teqemttme of 1540° F. The power regulation available in this case would have the additimal advantage of perm$tting &p appropriately d.esi@ed air- craft to operate at higher lift-drag ratios at reduced flight speeds, contr8.ryto the Umitations currentlyencounteredwithtucbojet~ This mode of operation would provide &n sdditianal powered aircraft- advantage in msximmirangewhileretaining thelergemarginia excess power available. This powerreductiolmfor cruisingcouldbeaccom- plished with an improvement in epecific fuel consu@Am, which is the opposite of the case far the nom&L engine matched at anlg one design point. Theadvantageofthis SRYmgementfcXmximumendurance, es in airport-izcaffic patterns, is also apparent in coqparism with HACh RM E5m22 the current fixed-et&or turbojet engine, which does not yet permit economical "stacki@ operation. The foregoing presentation is necessarily brief but serves to in&hate the nature of future developments that may lead to substantial iqrovements in over-all aircraft performance.

The application of the turbojet engine at very high flight speeds presents a different problem. At supersonic speed, any engine variable that affect8 the physical size of the fuselage or other body enclosing the engFneis of prime importance. The two most significant results of range analysis at supersonic speeds are that the compressor pressure ratio8 required are lower and that for any given compressor pressure ratio above 4 an optimum turbine-inlet texperature exists for greatestrange which la a higher temperature then that currently possible with uncooled turbines. This circum- stance results directly frost the high thrust per square foot of frontalarea that is obtainedwlthhighturbine-inlette~e~turee.

Typical results from an analysis of air-cooling the turbine blades of en axial-flow turbojet engine have been pr8sen-ted to illustrate theapplicationof theoryand experimsntaldataand to out- line the characterietics of the cooling process, the limitations that may occur> end some of the design considerations. These results ll?Ryb%SlXQUWized as follows: 1. For a fixed turbine-met tewerature, air cooling decreased The effects of air cooling on the specific thrust of the engine.

specific fuel consumption were considerably smaller.

2. The occurrence of diminishing returns with increase in degree of cooling was observed as a fundamental characteristic and suggested that complete elimInationof the strategic elements fromtheblade alloy is inadvisable on the basis of cooling requirements and losses.

3. The highest possible cooling effectiveness was desirable to ml.xhrLze losses in engine performance.

4. The required coolant-flow ratio increased with altitude, and some compressionof the coolingairwas generallyrequired for the COOlbIg4ir Btrpp4.

mcARME5cE22 5. The radial turbine-inlet-temperature distribution seemed to have a considerable effect on coolant-flow requirements, aSa the desirable distribution seemed to be unifoz% rather thau increasing radially with a peak occurring neeLT the blade tip.

6. Because of the high coolant-supply pressure required, it uaa generally nece8sary to bleed the cooling air from the compressor.

7. In selecting the blade profile , considerable attention must be given to the cooling of hot spots such es the leading and trailing edges, which otherwise may rtmmseveral hundred degrees hotter thau the midchord section of the blade.

8. The analyses indicated that substantial reduction of strategic- metal content in turbine-blade alloys cau be made with reasonable sacrifice in over-all engine performance.

9. Inmilitmyengines usingnaa&rategic materials, the rotor blades should still have the highest alloy content relative to the other parts exposed to the hot gases.

10. The application of cooled turbines with high pressure-ratio compressors and adjustable-blade-angle turbine stators offers improve- ment in flexibility, maximum range, aad endurmce.

LexisFlightPropulsionIaboratory, National Advisory CoPrmittee for Aeronautics, Cleveland, Ohio.

1. Ellerbrock, Herman H., Jr., and Schafer, Louis J., Jr.: Application of Blade Cooliug to Gae Turbines. NACA RM E5OAO4, 1950.

2. Euepper, K. H.: TepsperatureMeasurement ouTwoStationaryBucket Profiles for CasTurbines withBou~&sry-Iay~ Cooling. Trays.

NO- F-TS-1543-m, Air MaterielC ommd, U.S. Air Force, Jan. 1948.

18576, CADO.)

(AT1 No.

3. &-own, W. Bgron, and Rossbach, Richard 3.: Numerical Solution of Equation8 for One-Dimensional Gas Flaw in Rotating Coolant NACA RMESOE04, 1950.

Pas8eges.

24 HAcARME5OE22 4. Hensley, Reece V., Ram, Fkank E., and Emtz, Stanley L.: mf8Ct of Heat and Power Extraction on Turbojet-Engine Performance. .

ormutce Evaluation with Compressor- I -AnalyticalMethod 0fPerf Outlet Air Bleed. NACA TN 2053, 1950.

G 5. Pressman, Joeeph R., and Livingood, John R. B.: Cooling of Gas g Turbines. VII - Effectiveness of Air Cooling of Hollow Turbine Blades with Inserts. NAGA Rpd~7G30, 1947.

Reducing the Power Losses of Internally Cooled Turbines - 6. Kress, H.: F.B.1887, 10.23.43. 'Power Jets Rep. No. R-1185, Power Jets (Res. end Development), Ltd., Feb. 1946.

7. Silvern, David ,H., and Slivba, WilliamR.: Analytical Investi~tiou of Turbine8 with Adjustable Stator Blades and Effect of These NACA RM E5oEo5, 1950.

Turbines on Jet-Engine Perfmnance.

mCARME5oE22 .

COMPARISON OF smcANT ~CEVARUBZESFOR coom Am UN- TTJmomE ENGINE @ ltitude, 40,000 feet; flight &ch number, 0.788; turbine-inlet 1500' F; Cr-Ma-V& steel, lOOO-hour rupture life4 terqperature, ofblade Uncooled COC sd Lneert Fbned Compressor inlet weight flow, . .

lb/set . . . . . . . . . . - . 20.14 20.14 20.14‘ Required rotor-coolant . . - - - weight flow, lb/set . . . . . . 0.36 0.69 . . - e - Rotm coolant-flow ratio . . . .

0.018 0.034 Stator coolant-flowratio . . . . . me- 0.010 0.010 Rotor ccdant-temperature - e m rise,% . l l . . l . l l - l . .

624 379 . . - - - Heat loss to cooleaat, hp . . . . 91.6 80.6 - - - Coolant-paesage pressure ratio . . .

1.33 1.20 %Aernal coolant compressor we- power, hp.. . . . . . . . . . . .

46.3 65.9 Elterna1c00lantpuY&ping power,hp.. . . . . . . . . . . . - w - 26.0 49.6 Main compressor power plus auriliaries, hp . . . . . . .

2152 2092 2058 Totalturbine power,hp . . . . . 2152 2x4 2173 Turbine weight flow, lb/e,: . .

. . 20.50 19.92 19.60 . .

Specific turbine work* Btu/lb . 74.2 76.8 78.4 Turbine pressure ratio . . . . . . . 2.15 2.2l 2.25 Jet-nozzle pressure ratio . . . * . 2.79 2.72 2.67 . .

Jet-nozzle total temperature, % 1236 1215 us9 Thrust,lb . . . . . . . . . . . . .

950 917 898 _ Specific fuel consumption~ lb/lb-hr . . . . . . . . . . . . . 1.328 1.337 1.343 - - - Decrease in thrust, percent . .

. . 3.5 5.5 Increase in specific fuel consumption,percent . . . . . . .

0.70 1.15 .

NACA R M E5OE22 statm oooiins-air SuplDls -cl , NACA R M E!FOE22 (a) Insert blade.

(b) Finned blade.

Figure 2. - Cross sections of hollow air-cooled blades using two methods of increasing cooling effectiveness.

.

28 NACA RM E50E22 l lloo I Meets for computing blade-metal temperatur distribution /'/ 1 , // --- Varying geometry, relative 1ooo gas velocity, heat-transfer / coefficients, and gas properties . / --- Meti values of geometry, / relative gas velocity, heat-transfer coefficients, / - / and gae progerti- + / a % 8 r: . .

0 .2 .4 .6 .a 1.0 Radial blade station - Compzwisim of tuo m&hods for computing blade- Figure 3.

metal temperature distribution usUg gas-temperature profile shown. Finned blade; a&titu&, sea level; engine speed, 7600 rp; rotor-coolant-flow ratio, 0.02.

(Da- from reference 3.)

NACA R M E5OE22 -29 Rotor-coolant- s da H 8000S .

.2 . .a 1.0 R&al blade station Figure 4. - Comparison of coolant-flov requirement and blade- metal-temperature distribution for two gas-temperature profiles. Finned blade; altitude, set level; engine speed, 7600 rpm. Allovableblade-metal-temperature curve determined for Cr-Mo-Va steel with rupture life of approxtmately 4000 hours. ~~ypi06.1 gas-temperature profile and oorrssponding blade-metal dlatribution ourve from fig. 3.)

.

NACA R M E50E22 .

O R 1300 ii 01200 I3 1ooo ,a H acxl .4 .6 .2 .a 1, .O Radial blade .otation Figure 5. - Approxiraxte trailing-edge-temperature distribution for gas-temperature profile and midohord-seotion temperature ehown. Finned blade; ooolant-flow ratio, 0.02; thermal oonduotivity, 15 (Bt~/(hr)(i’ t)(~F)). (Typioal gaa- temperature refile and midohord-seotion-temperatura ourve flwn fig. 3.

NACA R M ~5oE22 Blade-x&al temperature, oF Figure 6. - Comparison of stress-rupture properties of various blade materials for 1000~hour blade-rupture life- Rotor-coolant-flow requirements against blade-metal temperature at m idspan for finned blade.

Altitude, hO,OCO feet; flight Mach number, 0.788.

NACA RM E50E22 NACA RM E50E22 (a) Required rotor-coolant-flow ratio. (a) Required rotor-coolant-flow ratio.

/ / -- - -- ---- (b) Percentage decrease in thrust.

I Blade .Insert ----____ ----m-----e-- 10 20 30 40x103 Altitude, ft (c) Percentage increase in specific fuel consumption.

Figure 7. - Required coolant flow, percentage decrease in thrust, and percentage increase in specific fuel consumption for finned and insert blades. Blade material, Cr-Mo-Va steel; blade-rupture life, loo0 hours; flight Mach number, 0.788; turbine-inlet temperature, lWO" F.

NACA RM E5oE22 Turbine-i

3T-

(a) Variation of thrust. Required coolant-flow ratios for finned blades of two alloys.

Total coolant-flow ratio (b) Variat$on of specific fuel consum&ion.

Figure 8. - Variation of thrust and specific fuel consumption with total coolant-flow ratio for three turbine-inlet temperatures.

Altitude, 40,000 feet; flight Mach number, 0.788; compressor pressure ratio, 4.42; engine speed, 7600 rpm; stator coolant-flow ratio, 0.01.

32 16 I Turbine-inlet 2 loo ---I-- P 596 .7 .8 .g 1 o 1.2 1.3 1.4 1 Specific fuel cokmptik~ lb/(hr)(lb thrust) Figure 9. - Perf’ ormance of turbojet-engine cycle. Flight Mach number, 0.738; altitude, 30,000 feet; compressor and turbine efficiencies, 0.90.

(Data from reference 1.)

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