toward improved durability in advanced combustors and ... · reduce material temperatures, (3)...

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NASA Technical Memorandum 88932

Toward Improved Durability in Advanced Combustors and Turbines-Progress in the Prediction of Thermomechanical Loads

Daniel E. Sokolowski and C. Robert Ensign Lewis Research Center Cleveland, Ohio

4 HAS&-38-88933) IOYdRE IdEGCVEO DURABILITY ~ 8 7 - 2 m 1 IIY ADVANCED CCflIELSTOBS ANE IOEEXBES: P B G G B E S S 111 Tl iE P B E D I C l l G l J CF

IBEBMCdECHAIICAL LCADS ( U A s A ) 3 1 p Ava i l : Unclas & Z I S trC A03/b€ A01 CSCL 21E G3/07 0097622

Prepared for the 3 1 st International Gas Turbine Conference and Exhibition sponsored by the American Society of Mechanical Engineers Dusseldorf, West Germany, June 8-12, 1986

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TOWARD IMPROVED DURABILITY I N ADVANCED COMBUSTORS AND TURBINES

PROGRESS I N THE PREDICTION OF THERMOMECHANICAL LOADS

NASA I s t t o addre

Daniel E. Sokolowski and C. Robert Ensign Nat iona l Aeronautics and Space Admin is t ra t ion

Lewis Research Center Cleveland, Ohio 44135

SUMMARY

sponsoring the Turbine Engine Hot Sect ion Technology s the need f o r improved d u r a b i l i t y i n advanced combu

(HOST) P r o j - t o r s and

tu rb ines . A n a l y t i c a l and experimental a c t i v i t i e s aimed a t more accurate pre- d i c t i o n o f the aerothermal environment, the thermomechanical loads, the mate- r i a l behavior and s t r u c t u r a l responses t o such loading, and l i f e p r e d i c t i o n s f o r h igh temperature c y c l i c operat ion have been underway f o r severa l years and a re showing promising r e s u l t s . Progress i s repor ted i n the development o f advanced ins t rumenta t ion and i n the improvement o f combustor aerothermal and t u r b i n e heat t rans fe r models t h a t w i l l l ead t o more accurate p r e d i c t i o n o f thermomechanical loads.

INTRODUCTION

Since i n t r o d u c t i o n o f the gas tu rb ine engine t o a l r c r a f t propuls ion, the quest f o r g rea ter performance has resu l ted i n a con t inu ing upward t rend i n o v e r a l l pressure r a t i o f o r the engine core. Associated w i t h t h i s t rend a re i nc reas ing temperatures o f gases f low ing f rom the compressor and combustor and through the tu rb ine . For commercial a i r c r a f t engines i n the foreseeable f u t u r e , compressor discharge temperature w i l l exceed 922 K (1200 O F ) , w h i l e t u r b i n e i n l e t temperature w i l l be approximately 1755 K (2700 O F ) . M i l i t a r y a i r c r a f t engines w i l l s i g n i f i c a n t l y exceed these values.

Since 1973 increas ing f u e l p r ices have created the demand f o r energy con- se rva t i on and more f u e l e f f i c i e n t a i r c r a f t engines. I n response t o t h i s

. demand, engine manufacturers con t inua l l y increased the performance o f the cur- r e n t generat ion o f gas t u r b i n e engines. Soon af terward, the a i r l i n e i n d u s t r y began t o experience a no tab le decrease i n d u r a b i l i t y o r use fu l l i f e o f c r i t i c a l p a r t s i n the engine ho t sec t ion- - the combustor and tu rb ine . This was due p r i - m a r i l y t o c rack ing I n the combustor l i n e r s , t u r b i n e vanes, and t u r b i n e blades. I n add i t i on , s p a l l l n g o f thermal b a r r i e r coat ings t h a t p r o t e c t combustor l i n e r s has occurred.

For the a i r l i n e s , reduced d u r a b i l i t y f o r i n -se rv i ce engines was measured by a dramat ic increase i n maintenance costs , p r i m a r i l y f o r h igh bypass r a t i o engines. Higher maintenance costs were e s p e c i a l l y ev ident i n the h o t sec t lon . As shown i n re ference 1, ho t sec t ion maintenance costs account f o r almost 60 percent o f the engine t o t a l .

D u r a b i l i t y can be improved i n hot sec t i on components by us ing a s i n g l e approach o r a combination o f f ou r approaches. They a re the use o f (1) mate- r i a l s having h igher use temperatures, (2 ) more e f f e c t i v e coo l i ng techniques t o

reduce m a t e r i a l temperatures, ( 3 ) advanced s t r u c t u r a l des ign concepts t o reduce stresses, and ( 4 ) more accurate a n a l y t i c a l models and computer codes i n the design ana lys i s process t o i d e n t i f y ho t spots, h igh stresses, e t c .

based supera l loys. Cer ta in elements of these a l l o y s , such as coba l t , a re i n sho r t supply and a r e expensive. Recently researchers completed a study o f ways t o reduce t h e i r usage ( r e f . 2 ) . Advanced h igh temperature supera l loy components a l s o inc lude d i r e c t i o n a l l y s o l i d i f i e d , s i n g l e c r y s t a l , and oxide- dispersion-strengthened mate r ia l s . For such ma te r ia l s , the development t ime i s lengthy, f a b r i c a t i o n i s sometimes d i f f i c u l t , and again costs a re h igh. Thus, successful use o f these ma te r ia l s requ i res a balance among design requirements, f a b r i c a t i o n p o s s i b i l i t i e s , and t o t a l costs .

High temperature m e t a l l i c ma te r ia l s c u r r e n t l y i nc lude n i c k e l - and coba l t -

Current cool ing techniques tend t o be soph is t i ca ted ; f a b r i c a t i o n i s moder- a t e l y d i f f i c u l t . I n h igher performance engines, coo l i ng c a p a b i l i t y may be improved by increas ing the amount o f coolant . There i s a pena l ty f o r doing t h i s , however, i n the reduc t ion o f thermodynamic cyc le performance o f the engine system. I n add i t i on , the coolant temperature o f such advanced engines i s h igher than tha t f o r cu r ren t i n -se rv i ce engines. Consequently, more e f f e c - t i v e coo l i ng techniques are being inves t iga ted . General ly, they are more com- p lex i n design; demand new f a b r i c a t i o n methods; and may r e q u i r e a m u l t i t u d e o f smal l coo l i ng holes, each o f which in t roduces p o t e n t i a l l i f e - l i m i t i n g h igh s t ress concentrat ions. Acceptable use o f the advanced coo l i ng techniques w i l l r e q u i r e accurate models f o r design ana lys is .

The i n t r o d u c t i o n o f advanced s t r u c t u r a l design concepts u s u a l l y begins w i th a p re l im ina ry concept t h a t then must be proven, must be developed, and-- most c r i t i c a l l y - - m u s t be f a r super io r t o entrenched standard designs. Accept- ance c e r t a i n l y i s t i m e consuming, and b e n e f i t s must be s i g n i f i c a n t . For improved d u r a b i l i t y i n h igh performance combustors, an e x c e l l e n t example o f an advanced s t r u c t u r a l design concept i s t he segmented l i n e r ( r e f . 3 ) . The l i f e - l i m i t i n g problems associated w i t h h igh hoop st resses were e l im ina ted by d i v i d - i n g the standard fu l l -hoop l i n e r s i n t o segments. A t the same t ime, designers r e a l i z e d increased f l e x i b i l i t y i n the choice o f advanced coo l i ng techniques and ma te r ia l s , i nc lud ing composite ceramics.

F i n a l l y , the design ana lys is o f h o t sec t i on component pa r t s , such as the combustor l i n e r s o r t u r b i n e vanes and blades, invo lves the use o f a n a l y t i c a l o r emp i r i ca l models. Such models o f t e n i n v o l v e computer codes f o r analyz ing the aerothermal environment, the thermo-mechanical loads, heat t r a n s f e r , and ma te r ia l and s t r u c t u r a l responses t o such loading. When the p a r t s a re exposed t o h igh temperature c y c l i c opera t ion as i n a t u r b i n e engine, the r e p e t i t i v e s t r a i n i n g o f the mater ia ls i n v a r i a b l y leads t o crack i n i t i a t i o n and propagat ion u n t i l f a i l u r e o r break-away occurs. The usefu l l i f e o r d u r a b i l i t y o f a p a r t i s u s u a l l y def ined as the number o f miss ion cyc les t h a t can be accumulated be fore i n i t i a t i o n o f s i g n i f i c a n t cracks. Thus, designers need t o p r e d i c t use- f u l I 1 l i f e l 1 so they can design a p a r t t o meet requirements.

E f f o r t s t o p red ic t the l i f e o f a p a r t genera l l y f o l l o w the f low o f analy- t i c a l models portrayed i n f i g u r e 1. Thus, des ign jng o f a p a r t such as a t u r - b ine blade t o meet a spec i f i ed l i f e goal may r e q u i r e a number o f i t e r a t i o n s through the " L i f e P red ic t i on Systemll of f i g u r e 1, va ry ing the blade geometry, ma te r ia l , o r cool ing ef fect iveness i n each pass, u n t i l a s a t i s f a c t o r y l i f e goal i s p red ic ted .

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Present models or codes frequently predict physical behavior qualitatively but exhibit unacceptable quantitative accuracy. In general, researchers need to improve predictive capability (1) to understand and model more accurately the basic physics of the phenomena related to durability, (2) to emphasize local as well as global conditions and responses, (3) to accommodate nonlinear and inelastic behavior, and (4) to expand some models from two to three dimen- sions. solution techniques, computer memory, and computer computational speed were increasing dramatically. These technological advances continue.

Fortunately, these needs were identified at a time when mathematical

OVERVIEW OF THE HOST PROJECT

To meet the needs for improved analytical design and life prediction tools, especially those used for high temperature cyclic operation in advanced combus-. tors and turbines, NASA is sponsoring the Turbine Engine Hot Section Technology (HOST) Project (refs. 4 to 8). The project was initiated in October 1980 and presently is planned to continue through 1989.

Objec ti ve

The HOST Project will develop improved analytical models for the aerothermal environment, the thermomechanical loads, material behavior, structural response, and life prediction, along with more sophisticated computer codes, which can be used in design analyses of critical parts i n advanced turbine engine combustors and turbines. More accurate analytical tools will better ensure--during the design process--improved durability of future hot section engines components.

Approach

The complex durability problem in high temperature, cyclically operated turbine engine components requires the involvement of numerous research disciplines. This Involvement must include not only focused research but sometimes interdisciplinary and integrated efforts.

Most disciplines in the HOST Project follow a common approach. First, phenomena related to durability are investigated, often using benchmark quality experiments. With known boundary conditions and proper instrumentation, these experiments result i n a characterization and better understanding of such phenomena as the aerothermal environment, the material and structural behavior during thermomechanical loading, and crack initiation and propagation. Second, state-of-the-art analytical models are identified, evaluated, and then improved by more inclusive physical considerations and/or more advanced computer code development. When no state-of-the-art models exist, researchers develop new models. Finally, predictions using the improved analytical tools are validated by comparison to experimental results, especially the benchmark quality data.

Programs

Fulfillment of the HOST Project objective is being accomplished through numerous research and technology programs. Numerous activities have been

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res a c t the and

r e c e n t l y completed o r a re i n progress. f o r 39 separate a c t i v i t i e s w i t h p r i v a t e i ndus t r y , most o f which a re mul t i -year and multi-phased. I n severa l a c t i v i t i e s , more than one con t rac to r i s invo lved because o f t he nature o f t he research and each c o n t r a c t o r ' s unique q u a l i f i c a - t i ons . w i t h u n i v e r s i t i e s . F i n a l l y , a t the NASA Lewis Research Center, 1 7 major

HOST management has issued con t rac ts

Th i r teen more separate a c t i v i t i e s a re being conducted through grants

arch e f f o r t s are supported by the p r o j e c t . v i t i e s being conducted i n the HOST Pro jec t . p r o j e c t ' s completion date, t he schedule f o r each a c t i v i t y i s being rev ised

Table I l i s t s a l l t he techn ica l Because o f recent changes i n

consequently, i s n o t presented i n t h i s repo r t .

RESULTS AND DISCUSSION

Engine manufacturers i n t e r e s t e d i n p r e d i c t i n g the l i f e o f a p a r t o f t e n use a " L i f e P red ic t i on System," as por t rayed i n f i g u r e 1. General ly, such a system def ines the engine opera t ing requirements and charac ter izes the h o t sec t i on environment; the system a l s o p r e d i c t s the behavior o f ma te r ia l s and s t ruc tu res w i t h i n t h a t environment, e s p e c i a l l y du r ing h igh temperature c y c l i c operat ion.

Development o f such a system requ i res work i n v o l v i n g s i x d i s c i p l i n e s : ins t rumenta t ion , combustion, t u r b i n e heat t r a n s f e r , s t r u c t u r a l ana lys is , f a t i g u e and f rac tu re , and sur face p ro tec t i on . HOST P ro jec t a c t i v i t i e s a re organized along those d i s c i p l i n e l i n e s . However, t h i s r e p o r t f o l l ows the " L i f e P red ic t i on System" f lowchar t t o rev iew progress on HOST a c t i v i t i e s . gress repo r t focuses on on ly two steps i n the p r e d i c t i o n system: charac ter iza- t i o n o f the aerothermal (ho t sec t ion) environment and o f thermomechanical loads. The repor t on each a c t i v i t y conta ins th ree elements: the need, the product, and h i g h l i g h t s o f the techn ica l r e s u l t s . References d i r e c t readers t o more d e t a i l .

The pro-

Character1 za t i on o f the Aerothermal Environment

Wi th in a gas t u r b i n e engine, the combustor i s o f pr imary importance t o ho t sec t ion d u r a b i l i t y . The l i f e o f c r i t i c a l ho t sec t i on pa r t s , such as the combustor l i n e r s and t u r b i n e blades and vanes, i s a f f e c t e d by the temperature l e v e l and un i fo rm i t y o f gases f l o w i n g over t h e i r surfaces.

The gas temperature l e v e l i s es tab l i shed by the engine performance requirements. Generally, h igher performance i n commercial a i r c r a f t engines 'IS obtained by increas ing the o v e r a l l pressure r a t i o and the component e f f i c i e n c y . One r e s u l t o f increased pressure r a t i o i s a p ropor t i ona te increase i n compressor discharge a i r temperature. combustor l i n e r and t u r b i n e a i r f o i l cool ing, t he impact i s h igher coo lan t temperatures. burn ing more f u e l and thereby i nc reas ing the combustor temperature r i s e . impact I s n o t only a g rea ter thermal load on the l i n e r s bu t a l s o a h igher average t u r b i n e i n l e t temperature.

Since p a r t o f the compressor a i r i s used f o r

M i l i t a r y a i r c r a f t engines ob ta in h igher performance p r i m a r i l y by The

Obta in ing gas temperature u n i f o r m i t y i s p a r t i c u l a r l y cha l leng ing where complex aerodynamics, chemical k i n e t i c s , and t u r b u l e n t heat t r a n s f e r take place. de l i ve red t o i t by the compressor. Compressor discharge a i r , f o r instance, can e x h i b i t d i f f e r e n t f l ow v e l o c i t y r a d i a l p r o f i l e s as the engine power s e t t i n g i s

The combustor must accept a range of c h a r a c t e r i s t i c s i n a i r f l o w

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changed and as the engine ages. This can a f f e c t t h e f l o w s p l i t s f o r combustion a i r and f o r shroud a i r , which I s used f o r d i l u t i o n and l i n e r coo l ing . form compressor a i r f l o w c i r c u m f e r e n t l a l l y can impact l i n e r backside tempera- tu res and cause l i n e r temperature a x i a l s t reak ing . ' I t a l s o has been known t o a f f e c t combustor e x i t temperature p r o f i l e s and p a t t e r n f a c t o r , which d i r e c t l y i n f l u e n c e thermal load ing p r o f i l e s on t u r b i n e f i r s t stage vanes and blades.

Nonuni-

F i n a l l y , the combustion system must accept a range i n the q u a l i t y o f f u e l s t o be burned. I n recent t imes, t he aromatic conten t o f a i r c r a f t gas t u r b i n e f u e l s has n o t on ly increased b u t var ies f rom one l o c a t i o n t o another i n the wor ld. The most no tab le e f f e c t o f lower q u a l i t y f u e l i s i n coke format ion. I f the coke bui ld-up occurs on the fue l i n j e c t o r , t h e change i n spray p a t t e r n may be s u f f i c i e n t t o cause temperature s t reak ing on the l i n e r s . Coke deposi ts on the i n j e c t o r s o r l i n e r s may reach a c e r t a i n s ize, break away, and f l o w i n t o the t u r b i n e where they cause "shot-peening" o f a i r f o i l surfaces--some o f which may have thermal b a r r i e r coat ings- or become trapped between a r o t a t i n g b lade and t u r b i n e case, w i t h subsequent damage t o the seal . Another e f f e c t o f h igher aromat ic f u e l s i s an increase i n rad ian t heat load ing on ho t sec t ion p a r t s t h a t 'I s e e II s 11 r h rad i 5 t 1 nn

With the above considerat ions i n mind, i t i s o f utmost importance t o char- a c t e r i z e and understand the aerothermal environment around ho t sec t i on pa r t s . This i s t he i n i t i a l s tep i n developing more accurate models t o a n a l y t i c a l l y p r e d i c t such phenomena. Such cha rac te r i za t i on i s prov ided by appropr ia te advanced ins t rumenta t ion t h a t i s used i n c a r e f u l l y designed and conducted experiments. To t h i s end, several a c t i v i t i e s supported by the HOST P r o j e c t were undertaken and have now been completed. Such a c t i v i t i e s under the inst ru- mentat ion d i s c i p l i n e a l l o w one t o view t h e i n t e r i o r o f an opera t ing combustion chamber, t o measure h igh frequency gas temperature f l u c t u a t i o n s , and t o measure instantaneous and average gas f l o w v e l o c i t i e s . Under t h e combustlon d i s c i - p l i n e , a broad data base has been obtained f o r d i l u t i o n j e t m ix ing and has been used t o develop emp i r i ca l r e l a t i o n s tha t a l l o w designers t o p r e d i c t combustor e x i t temperature p r o f i l e s . I n add i t ion , s t a t e - o f - t h e - a r t aerothermal models o f the combustion system have been assessed and recommendations made f o r requ i red improvements. Summaries o f a c t l v i t i e s t h a t have improved the researcher 's or des igner 's a b i l i t y t o charac ter ize the aerothermal environment a re presented below.

i n s i d e an opera t ing t u r b i n e engine hot sec t i on component i s impor tant f o r de tec t i ng abnormal opera t ion and f a i l u r e modes. For example, l ook ing i n s i d e a combustor w i l l he lp determine the l oca t i on and shape o f the flame and the oper- a t i n g cond i t ions under which the flame impinges on the combustor w a l l o r l i n e r . The c o n d i t i o n o f t he l i n e r ( s ) , f u e l I n jec to rs , bulkhead, and t u r b i n e i n l e t guide vanes cou ld be q u a l i t a t i v e l y assessed du r ing combustor operat ion. Observed ho t spots can be monitored f o r p rogress ive l i n e r damage, i n c l u d i n g crack i n i t i a t i o n and propagat ion. accumulation o f coke deposi ts and f u e l spray p a t t e r n and t o measure r e l a t i v e o r absolute flame temperatures. Flame I n s t a b i l i t y and turbulence could be observed and/or measured and poss ib l y r e l a t e d t o combustor acoust ic no ise.

The purpose f o r t h i s a c t i v i t y , which was conducted by Uni ted Technologies Research Center (UTRC), was t o develop an o p t i c a l system f o r v iewlng the i n t e r - i o r of a combustor ( o r perhaps tu rb ine) du r ing h igh pressure, h igh temperature opera t ion and t o exp lo re methods f o r improving the i n fo rma t jon content o f

Hot sec t i on v iewins system. - Development o f a d iagnos t ic t o o l t o look

I t a l s o may be poss ib le t o observe the

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recorded images. The final system was planned for use inhouse at NASA Lewis as a research tool and for loan to qualified users.

The completed viewing system consists of a water-cooled probe, a probe actuator, controls, an optics/light source interface, and interchangeable viewing optics. view and resolution options. while figure 3 shows the probe.

The interchangeable viewing optics provide a variety of field-of- Figure 2 is a schematic of the viewing system,

Development of the system has included testing of a pulsed laser illumination source and computer image enhancement. Followjng delivery of the system to NASA Lewis, it has been used in the High Pressure Facility to obtain high speed movies of burning around fuel injectors. Using the results of this research, UTRC i s constructing several more probes for its own corporate use, as well as for an electric power company to use on a gas turbine generating system. tor test rig. Reference 9 provides more detailed information on this viewing system.

The U.S. Navy soon will be using the NASA viewing system in a combus-

Researchers believe that future testing with the system using pulsed laser illumination, coupled with a suitable filter and a short duration shutter that allows suppression of the intense flame illumination, will provide a ltview'l through a luminous flame. Image enhancement then can be used to expand the range of contrast of images that have been recorded.

Dynamic gas temperature measurement system. - In gas turbine engines, it is generally accepted that large scale gas temperature and velocity fluctuations occur throughout the combustion chamber and i n the combustor exhaust gases because o f incomplete mixing o f the combustion gases and air used for dilution and liner cooling. The fluctuating temperatures and velocities of these exhaust gases have been related to damage to turbine vanes and blades. That is, temperature variations in the gas stream cause measurable temperature variations on the surface of such parts. Thus, a probe that can accurately measure the amplitude and frequency o f the gas stream temperature fluctuations is necessary in order to improve the prediction of turbine blade and vane life In addition, the understanding of turbulent combustion requires a knowledge o f the temperature variation with time.

In the past, high-response measurements have been attempted by using a passive electrical compensation network based on an estimate of the temperature probe time constant. The intent of an effort performed by Pratt and Whitney was to develop a more rigorous method for measuring gas temperature fluctuations at frequencies up to 1000 Hz with peak temperatures of up to 1922 K (3000 O F ) and a temperature fluctuation range of up to 2755" (1360 O F ) . The system had to be capable of measurements in a combustor exhaust at pressures up to 20 atmospheres.

The approach chosen for development is shown in figure 4. It features a two element thermocouple probe. A compensation function is derived using the two-thermocouple signal amplitude ratio in response to the gas temperature fluctuation. Frequency compensation for the recorded signals can then be accomplished for high frequency fluctuations up to a limit imposed by signal- to-noise ratio.

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The measurement and compensation system have been developed and evaluated in equipment ranging from laboratory burners to full-size commercial and military engines. atmospheric combustor are presented in figure 5, while similar results using a F-100 engine combustor are presented in figure 6. In each figure, the instantaneous temperature versus time is shown: first, for the uncompensated thermocouple output and, second, for the compensated thermocouple output. The dynamic temperature range for the compensated output not only shows the importance of compensation but also the significant fluctuations in temperature levels on the order of 2755 K (1360 O F ) occurring in a turbine engine combustor. See reference 10 for details on the probe concept selection, development, and testing.

Test results with the probe located at the exit plane of an

Further efforts i n progress are (1) experimentally verifying the accuracy of the probe using a specially designed rig to produce known temperature fluctuations, (2) reducing the time for data processing, and (3) reprogramming the compensation system to run on a general purpose computer.

Laser anemcmetry f e r hc t sertic:: appllcatlcns. - The gases flcwing thrcugh a turbine engine hot section exhibit complex behavior, including turbulence. In the turbine component with its successive stages of flow turning vanes and rotating blades, the behavior is difficult to characterize, let alone model, for performance and heat transfer. Understanding of such flow behavior requires a detailed knowledge of the flow stream velocity components in the axial, circumferential, and radial directions. Such information is needed within the vane passages, between vanes and blades, and within the blade passages; it must also include secondary flows over the tips of the blades. To understand the flow behavior and to measure flow velocity components, researchers at the NASA Lewis Research Center used the following approach: 1 ) conduct experiments of increasing complexity from a single bench type to possibly real engine rigs and 2) develop a nonintrusive measuring probe that will not be damaged by the hostile environment, will not change the flow behavior being measured, and will provide measurements within rotating blade passages. The only means to do this i s with a laser anemometer system. Such a system has been developed; it currently measures a single velocity component in these hostile environments.

The primary optical design for the laser anemometer is a conventional single component dual-beam, fringe-type configuration, as illustrated in figure 7(a). System components include the optics, a traversing stage for three-coordinate positioning of the system relative to the test rig, a particle generator for seeding the gas stream with 1.0 pm alumina, and fused silica win- dows machined to match the inside curvature of the turbine case. nents of the system, including a minicomputer for traverse control and data acquisition, have been developed and tested in an atmospheric jet combustor facility. bustor rig. An identical system is being installed In the Warm Turbine Rig at Lewis. Test measurements will provide data for mean velocity, flow angle, turbulence intensity, and turbulence scale. A further discussion of the system development and preliminary test results may be found in references 5 and 6. Analytical efforts performed during optimization of the laser anemometer system are reported i n reference 11.

All compo-

Figure 7(b) shows the laser anemometer system adjacent to the com-

Combustor aerothermal modeling. - To enable the design of a combustor that alleviates high temperature gradients and produces a desired exit temperature

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distribution, researchers must develop a thorough and accurate characterization and understanding of the convective and radiative heat fluxes in the combustor. This is the driver for the combustor aerothermal modeling program.

The objectives of this three phase program are: first, identify state-of- the-art combustor aerothermal models, assess their predictive capability, and recommend improvements; second, develop beneficial improvements to the aerody- namic aspects of the more promising models; and finally, incorporate chemical reactions into the models that will then allow local heat flux and temperature predictions.

The first phase of the effort, model assessment, was completed by Garrett Turbine Engine Company, Pratt and Whitney, and General Electric Company and is reported in references 12 to 14, respectively. In brief, current models can qualitatlvely predict the complex three-dimensional flow field within a combustor. Quantltatlve modeling of these flows, however, requires the reduction o f both computer execution times and removal of numerical diffusion in the calcu- lations. Researchers believe that as more efficient numerical solutions are developed, effective investigations will be possible in areas such as modeling of turbulence and scalar transport in the interaction of various flow streams and the development of the fuel spray and its interaction with the surrounding airstream. Finally, the assessment efforts discovered a serious deficiency of available and suitable benchmark quality experiments.

Based on findings and recommendations of the first phase effort, the second phase efforts, currently in progress, are necessarily focused on improving analytical models and computer codes. ments also are being conducted to aid model development and validation. activities include efforts to improve numerics, flow Interaction experiments, and fuel swirl characterization experiments. Allison Gas Turbine Division is involved in all three general areas, while AVCO-Everett, United Technologies Research Center, and the University of Minnesota are each addressing a specific area. Additional efforts at Purdue University and the University of California at Irvine directly support the effort at Allison. Table I indicates the contract or grant number associated with each contractor/grantee and the work being addressed.

Important benchmark quality experi- The

Dilution let mixing studies. - Mixing combustion gases with cooler dilution air inside the combustor is important not only to reduce the bulk temperature level of gases flowing into the turbine but also to provide acceptable circumferential and radial temperature profiles at the combustor exit. Since these profiles determine the local thermal loads on turbine first stage vanes and blades, they directly influence durability and cooling requirements o f such parts. Previous NASA-sponsored efforts dealt primarily with obtaining experimental data and developing empirical correlations for a single row of jets mixing with an isothermal normal flow in a constant cross-sectional area duct. A more inclusive data base was needed to develop such correlations further and improve temperature prediction accuracy for combustors.

The purpose for this activity, which was conducted by the Garrett Turbine Engine Company, is twofold. First, extend the data base for jet mixing to include realistic effects of combustion chamber (1) flow area convergence, (2) nonisothermal mainstream flow, (3) opposed (two-sided) in-line and staggered injection, (4) orifice geometry, and (5) double (axial) rows of holes. Second,

8

extend emp i r i ca l co r re la t i ons . An i n t e r a c t i v e computer code i s under development a t NASA Lewis f o r mix ing analyses and three-dimensional p i c t o r i a l repre- sen ta t i on o f the temperature f i e l d a t t he combustor e x i t .

A schematic o f t he t e s t sec t ion and o r i f i c e con f igu ra t i ons i s shown i n f i g u r e 8. ranged from 477 K ( 4 0 0 O F ) t o 811 K (1000 O F ) , and normal ly i n j e c t e d a i r j e t s having a nominal value o f 294 K (70 O F ) . Changes i n temperature d i s t r i b u t i o n a t t he t e s t s e c t i o n e x i t due t o va r ia t i ons i n mix ing were noted p r i m a r i l y as a f u n c t i o n o f jet- to-mainstream momentum f l u x r a t i o , o r i f i c e c h a r a c t e r i s t i c s , and spacing .

Tests were conducted using heated a i r f o r t h e mainstream, which

An example o f a t y p i c a l d i l u t i o n j e t f l o w f i e l d i s shown i n f i g u r e 9 w i t h a p i c t o r i a l representa t ion o f the temperature p r o f i l e s shown f o r exper imental data and empi r i ca l model p red ic t i ons . Several pub l i ca t i ons t h a t cover the th ree work phases a re i nd i ca ted by references 15 t o 17, respec t i ve l y . Resul ts o f v e l o c i t y measurements and co r re la t i ons , a long w i t h temperatures and v e l o c i - t i e s ca l cu la ted us ing a three-dimensional numerical code, a re for thcoming i n a repnr t Under rrrntrart NAS2-22!!0.

The exper imental data and empir ica l model developed i n t h i s program have been p u t i n t o use by engine manufacturers as a design t o o l . an emp i r i ca l model ou ts ide the measurement range, however, i s unce r ta in a t best . The a p p l i c a b i l i t y o f these empir ica l models i s l i m i t e d t o the geometry and f l o w cond i t ions and t o parameters t h a t have been co r re la ted . Numerical models do n o t have these l i m i t a t i o n s , and f o r more gener ic app l i ca t i ons a three-dimensional numerical model w i l l be needed as a p r e d i c t i v e t o o l . Current numerical r e s u l t s compared w i th the experiments a re very promising (see f i g . 9) and i n d i c a t e t h a t they can be used now i n t h e i r present form as a d iagnos t ic t o o l f o r e s t a b l i s h i n g f l o w trends. The numerical model c o r r e c t l y descr ibes the e f f e c t s o f t he p r i n c i p a l f l o w and geometric var iab les .

The usefulness o f

Charac ter iza t ion of t h e Thermomechanical Loads

Fo l low ing the cha rac te r i za t i on o f t he aerothermal environment, the next impor tan t step i n ho t sec t ion analyses i s the c h a r a c t e r i z a t i o n and understand- i n g o f thermal and mechanical loads imposed on the p a r t (see f i g . 1) . HOST P ro jec t a c t i v i t i e s i n t h i s area range f r o m advanced ins t rumenta t ion measuring heat f l u x , t o extens ive heat t rans fe r experiments, and t o improvements o f models and computer codes. I n the HOST P r o j e c t emphasis has been on aerothermal loads. A number o f these a c t i v i t i e s a re discussed below.

Turb ine heat f l u x sensors. - For a i r cooled t u r b i n e blades and vanes, a c r i t i c a l design parameter i s the hot s ide heat t r a n s f e r c o e f f i c i e n t . Current techniques f o r es t imat ing t h i s c o e f f i c i e n t do no t p rov ide the accuracy needed t o p r e d i c t ac tua l a i r f o i l temperatures. To prov ide more accurate heat t r a n s f e r c o e f f i c i e n t values, researchers must know the heat f l u x t o an a i r f o i l surface, p lus the f l o w stream and sur face temperatures. Very l i t t l e data i s a v a i l a b l e on heat f l u x i n r e a l i s t i c gas tu rb ine environments. Researchers know t h a t con- v e c t i v e heat t r a n s f e r i s predominant a t lower t u r b i n e pressures and tempera- tu res . performance gains, a long w i t h the p o s s i b i l i t y o f burn ing more luminous broad- ened s p e c i f i c a t i o n fue l s , s c i e n t i s t s b e l i e v e t h a t r a d i a t i o n heat t r a n s f e r i s

But as t u r b i n e engine pressure and temperatures are increased f o r

9

of greater importance. total heat flux is important.

Thus, development of a sensor that accurately measures

The objectives of this work are, first, to deveTop and test total heat flux sensors that are suitable for use on turbine airfoils and, second, to compare several methods of measuring heat flux in an atmospheric combustor rig.

The effort, which was conducted by Pratt and Whitney (P&W), built on the experience gained in a pre-HOST program that developed total heat flux sensors for use in combustor liners (see ref. 18). For airfoils, two sensor designs were identified as meeting the following criteria: sensors must be compatible with blade and vane materials (nickeland cobalt-based alloys), must have good spatial resolution (sensing dlameter less than 0.15 cm or 0.06 in.), mujt not interfere significantly with local temperature distribution and heat flow, and must not change the aerodynamics of the airfoils. criteria were the embedded thermocouple design (fig. lO(a)) and the Gardon gage design (fig. 10(b)).

The sensors meeting these

Development of the blade and vane heat flux sensors has been completed and

Results from this work indicate that these

reported in reference 19. other heat flux measuring techniques has been completed recently and will be reported under contract NAS3-23529. sensors are sensitive to transverse temperature or heat flux gradients. Fur- ther work to investigate this sensivity and to devise ways to minimize it is planned. In addition, demonstration tests o f combustor liner total heat flux sensors have been reported in reference 20. This work compares measurements made with a total heat flux sensor and a commercial optical radiometer. Typi- cal results fromtwo types of heat flux sensors are presented in figure 1 1 .

Experimental work to compare these sensors with

Turbine airfoil external heat transfer. - The trend toward higher turbine inlet temperatures to achieve advanced turbine engine performance imposes a more severe thermal load on turbine airfoils--that is, the nozzle guide vanes and rotating blades. At the same time, lack of the ability to predict local gas-to-airfoil heat transfer rates with acceptable accuracy is a principal obstacle toward timely and cost-effective development of high temperature turbine hardware. Improvements in predictive capability in this area can have broad and significant benefits in terms of enhanced turbine life, reduced development and maintenance costs, and improved engine performance.

All modern high pressure turbine airfoils contain complex cooling techniques. It is convenient to separate these techniques into those that become part of the external flow field (i.e., film cooling) and those that do not. Two efforts have been completed recently that experimentally and analytically address the needs of both types of airfoil cooling. In each effort, the objectives included these: (1) assess the capability of currently available modeling techniques for predicting airfoil surface heat transfer distributions i n a two-dimensional flow field, (2) acquire experimental data as required for model improvement and validation, and (3) further develop analytical methods.

The first effort, which is reported in detail i n reference 21, focused on the external surface heat transfer of a convectively cooled airfoil. data base covering a wide range of airfoil profiles and operating conditions was provided by two sources. Three data sets were selected from the litera- ture. Two additional airfoils (Identified as C3X and Mark 11). representative of highly loaded, low solidity airfoils, were tested in a three vane cascade

A broad

10

a t s imulated engine cond i t ions . f i g u r e 12.

The two a i r f o i l p r o f i l e s tes ted a re shown i n

The a n a l y t i c a l a c t i v i t y was conducted because the re i s no standard f o r c o n s i s t e n t l y and accura te ly p r e d i c t i n g ex te rna l convect ive heat t r a n s f e r t o s o l i d sur face a i r f o i l s opera t ing i n a gas t u r b i n e environment. ensure e a r l y and wide user acceptance o f any improvements i n p r e d i c t i v e capa- b i l i t i e s , t h e I n i t i a l task i n t h i s a c t i v i t y i nvo l ved i d e n t i f y i n g and assessing the c u r r e n t s ta te -o f - the-ar t p r e d i c t i v e methods. Three boundary l aye r methods were selected. The major d i f fe rences among the methods were i n the a n a l y t i c a l form o f t he governing equations t h a t were solved and the complexi ty o f the t u r - bulence model assumed. A f t e r documenting the inadequacies o f the methods over the broad range o f a i r f o i l p r o f i l e s and opera t ing cond i t ions , researchers chan- neled f u r t h e r examinations and development i n t o one promising method. This i s the d i f f e r e n t i a l boundary l aye r method, a v a i l a b l e I n the computer code STAN5 ( r e f . 22), which uses a mix ing length hypothesis (MLH) or zero-equation turbulence model fo rmula t ion . To b e t t e r p r e d i c t a i r f o i l heat t r a n s f e r , s c i e n t i s t s f u r t h e r developed the turbulence model i n STAN5. D e t a i l s o f t h i s development and the resu!t!ng mnrl!f!od STAN5 cndo are presented !n reference 21. ?c! ! ! I & s t r a t e the improved p r e d i c t i o n c a p a b i l i t y , f i g u r e 13 presents a comparison o f the C3X a i r f o i l exper imental heat t rans fe r c o e f f i c i e n t da ta w i t h the unmodif ied and mod i f ied STAN5 p red ic t i ons .

To b e t t e r

A second e f f o r t , I n v o l v i n g leading edge f i l m cooled a i r f o i l s , addresses d u r a b i l i t y requirements i n some advanced t u r b i n e engines. I t has been demon- s t r a t e d t h a t m u l t i p l e ho le ( i .e. , shower head) f i l m coo l i ng o f t he c r i t i c a l lead ing edge reg ion can s i g n i f i c a n t l y improve d u r a b i l i t y f o r a i r f o i l s t h a t o therwise a r e cooled i n t e r n a l l y by a combination o f convect ion and j e t imping- ment techniques. Minimal systemat ic e f f o r t has been d i rec ted a t cha rac te r i z ing l o c a l heat t r a n s f e r downstream o f leading edge i n j e c t i o n f o r h i g h l y loaded a i r - f o i l surfaces opera t ing a t r e a l i s t i c l e v e l s o f Mach number, Reynolds number, and wal l - to-gas and coolant-to-gas temperature r a t i o s .

Researchers have completed the second e f f o r t and repor ted on i t i n d e t a i l i n re ference 23. These researchers extended the i n i t i a l data base by us ing the same two-dimensional l i n e a r cascade and C3X a i r f o i l p r o f i l e , having a lead ing edge showerhead a r ray c o n s i s t i n g o f f i v e rows o f c o o l i n g holes fed by a common

. plenum. The f i l m cooled a i r f o i l i s shown i n f i g u r e 14, along w i t h the convec- t i v e l y cooled design f o r comparison. Recovery reg lon heat t r a n s f e r measurements were taken a t two t ransonic e x i t Mach number cond i t ions w i t h t r u e chord Reynolds numbers o f order l o 6 . I n add i t ion , both b lowing s t rength and coo l - an t temperature were var ied t o quant i f y j e t turbulence product ion and thermal d i l u t i o n mechanisms.

The researchers then used the extended data base t o guide development of a mathematical model f o r desc r ib ing the h i g h l y complex, three-dimensional, coo lan t jet /mainstream f l o w i n t e r a c t i o n process I n terms o f a two-dimensional boundary l a y e r ana lys is framework. The o b j e c t i v e was t o c rea te a method f o r p r e d i c t i n g recovery reg ion ex te rna l convect ive heat t r a n s f e r phenomena associ - a ted w i t h a lead ing edge f i l m coo l i ng process. Indeed, a d i r e c t extens ion o f a nonf i lm cooled two-dimensional boundary l aye r ana lys i s fo rmula t ion t h a t i s repor ted I n d e t a i l i n re ference 23. In b r i e f , two parameters (FTU f o r free-stream turbulence i n t e n s i t y f a c t o r and FTG f o r f ree - stream t o t a l gas temperature f a c t o r ) are de f ined t o model turbulence produc t ion and thermal d i l u t i o n phenomena. Computationally, these two parameters a re used

The method developed i s ,

11

to alter the boundary layer outer edge free-stream turbulence intensity and total gas temperature (enthalpy) boundary conditions. A comparison of results from model prediction and experimental data is presented in figure 15.

Finally, as computer technology evolves, boundary layer analysis design procedures are likely to be replaced by Reynolds and/or full Navier-Stokes (N-S) equation analyses. standard. Its continued development, however, will improve future capability and reduce empirfcism. In some areas, as shown above, this is an absolute necessity. by the HOST Project. The status of this analytical effort to apply time- dependent ensemble-averaged Navier-Stokes equations to transonic turbine cascade flow fields is reported In references 7 and 24. Results to date indicate the ability of the N-S analysis to predict, in reasonable amounts of computation time, the surface pressure distribution, heat transfer rates, and viscous flow development for turbine cascades operatlng at realistic flow conditions. This code's abilfty to predict the C3X data is shown In figure 16. that the code seems to be picking the chordwise fluctuations in the data is particularly encouraging.

Currently, the N-S framework is not an industry

To this end, a N-S method development program is being supported

The fact

Three-dimensional boundary layer code assessment. - It is widely recog- nized that turbomachinery flows are three-dimensional. Two-dimensional boundary layer methods, such as STAN5, are at best only applicable to midspan regions, such as described above. Some important insight into three- dimensional effects can be gained by the use of three-dimensional boundary layer analyses. As part of the HOST Project, an effort was directed toward the assessment, improvement, and documentation of the three-dimensional boundary layer code descrlbed in reference 25. dimensional boundary layer analyses is the definition of the boundary conditions. They, too, are frequently three-dimensional. The work described i n reference 25 pald particular attention to this problem and developed some gen- eralized techniques for defining the boundary or edge conditions.

A particular problem with three-

The results of the assessment of a three-dimensional boundary layer analysis are demonstrated in figure 17, which compares experiment and analysis for end-wall heat transfer. The results show both the strength and the weakness of a three-dimensional boundary layer analysis. Qualitatively the agreement is good. The overall levels compare quite well, and the Stanton number contours in the trailing edge region show very similar trends. edge region near the so-called saddle point, however, the experimental Stanton number contours are not captured by the analysis. This is the region where the horseshoe vortex is formed that has been shown to have a strong heat transfer effect. The horseshoe vortex cannot be represented in this boundary layer analysis .

In the leading

Gas-side heat transfer with rotation. - All of the turbine heat transfer work described so far was conducted in the stationary frame of reference. Rotation, of course, is an important variable in turbines. Accordingly, the project includes several efforts conducted in the rotating frame of reference. These efforts are just beginning to produce results. For now the program can be described, but only a qualitative preview of the results can be presented.

One effort continues to focus on the external flow over the turbine airfoils. The work is being done in a large low speed turbine that has already

12

been used to obtain a significant body of turbine data, as for example rotor- stator interactions described in reference 26. includes heat transfer coefficients along the midspan o f the pressure and suction surfaces on the vanes and the rotating blades. turbulence are being studied. The rig is to be configured as both a single stage, shown in figure 18, and a one and one-half stage turbine. The slngle- stage data are complete. The turbulence effects are quite significant on the first vane. The overall heat transfer is highest on the blades, but for the most part the effect of turbulence is submerged in the vane wake effects. Although rotational forces exist on the external airfoil surfaces in a rotating machine, they are not thought to be large. greater.

The data being acquired

Both high and low inlet

The vane wake effects are much

Coolant-side heat transfer with rotation. - The findings above with gas side heat transfer do not apply to the internal coolant passages. ancy and Coriolis forces are expected to be quite strong. passages themselves are very complex. This complexity is illustrated in figure 19. s!mu!atPd !r! ;rn exper!ment t h a t rntztos r t e s t sect!o!! w!t!? cutf!eu, !nf!ou, and back out again. Both smooth and turbulated surfaces are being studied. The turbulators are at 90" and 45" to the flow. Researchers are using dimensional analysis on the data. The smooth wall data are also being compared to computations using the TEACH code (ref. 27), which has been modified to include rotatjonal terms. The smooth wall experiments are complete. The trends in the data are shown in figure 20. The pronounced effect of rotation is obvious. Early indications suggest this data set can be successfully correlated by appropriate nondimensional groupings of buoyancy and rotational parameters. Further information on this activity can be found i n references 5 to 7.

Here buoy- In addition the

The serpentine passages with turbulators on the surfaces are being

CONCLUDING REMARKS

In this paper, progress has been reported on 16 of the 69 research activities that have been initiated and supported by the HOST Project. The focus has been on more accurate prediction o f the aerothermal environment and thermal loads. Progress in the remaining efforts is also being made, especially for actlvities related to more accurate prediction of material behavior, structural response, and life prediction.

The cooperative efforts among contractors, academia, and government con- nected with the HOST Project have been outstanding. While the project justifi- cation was based on civil aircraft propulsion needs, the technical results will have both civil and military applications.

For the improved or new analytical models and codes that are being developed in the HOST Project, experimental validation is limited primarily to the benchmark quality data. The burden is on the user community to further vali- date the HOST products, where required, with real engine-like data and, i n turn, to develop confidence for broader usage of the models and codes.

During the past year, a national policy statement was made by G . A . Keyworth, I1 (ref. 28) that addresses specific goals for the United States in three areas--subsonlcs, supersonics, and transatmospherics. With propulsion as the pacing technology to achieve these goals, it seems certain that the trends toward increasing hot section pressure and temperatures will continue.

13

The higher temperatures w i l l cont inue t o chal lenge t h e d u r a b i l i t y requirements o f advanced combustors and tu rb ines and perhaps even h ighe r pressure r a t i o com- pressors. I n the fu tu re , nonmeta l l ic h igh temperature ma te r ia l s , such as po ly - mer m a t r i x composites, ceramic m a t r i x composites, arid m o n o l i t h i c ceramics, w i l l even tua l l y be introduced t o hot sec t i on a p p l i c a t i o n s . Successful and cost - e f f e c t i v e app l i ca t i ons o f such m a t e r i a l s w i l l r e q u i r e s p e c i f i c a n a l y t i c a l models--for which few e x i s t . However, they may evolve f rom t h e c u r r e n t models and codes i n t h e HOST P ro jec t . F i n a l l y , r e v o l u t i o n a r y s t r u c t u r a l design concepts may be required, p r i m a r i l y f o r a p p l i c a t i o n o f t h e nonmeta l l ic m a t e r i a l s . To ensure e a r l y success i n such app l i ca t i ons , gener ic models and codes w i l l be requi red f o r design analyses.

REFERENCES

1. Dennis, A.J., and Cruse, T.A., "Cost Bene f i t s f rom Improved Hot Sect ion L i f e P red ic t i on Technology," A I A A Paper 79-1154, June 1979.

2. Stephans, J.R., "COSAM Program Overview," COSAM (Conservat ion o f S t r a t e g i c Aerospace Ma te r ia l s ) Program Overview, NASA TM-83006, 1982, pp. 1-11.

3. Tanr ikut , S., Marshal l , R.L., and Sokolowski, D.E., "Improved Combustor D u r a b i l i t y - Segmented Approach w i t h Advanced Cool ing Techniques," A I A A Paper 81-1354, J u l y 1981.

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8. Sokolowski, D., and Ensign, C . , "Improved D u r a b i l i t y i n Advanced Combustors and Turbines through Enhanced A n a l y t i c a l Design Capab i l i t y , " A I A A Paper 85-1417, J u l y 1985.

9. Morey, W.W., "Hot Sect ion Viewing System," R84-925830-33, Uni ted Technologies Research Corp., East Har t ford, CT, 1984. (NASA CR-174773)

10. Elmore, D.L., Robinson, W . W . , and Watkins, W.B., "Dynamic Gas Temperature Measurement System," PWA/GPD-FR-17145-VOL-l, P r a t t and Whitney A i r c r a f t , West Palm Beach, FL, 1983. (NASA CR-168267)

11. Seasholtz, R.G., Oberle, L.G., and Weikle, D.H., "Opt imizat ion o f Fr inge- Type Laser Anemometers f o r Turbine Engine Component Testing," A I A A Paper 84-1459, June 1984.

12. Sr in ivasan, R., e t a l . , "Aerothermal Modeling Program, Phase I," GARRETT- 21-4742-1;-2, G a r r e t t Turbine Engine Co., Phoenix, AZ, 1983. (NASA CR-168243- VOL-1; -VOL-2)

13. Sturgess, G.J., "Aerothermal Modeling Program - Phase I , " [WA-5907-19, P r a t t and Whitney A i r c r a f t , East Har t ford, CT, 1983. (NASA CR-168202)

I 1 4

14. Kenworthy, M.J., Correa, S;M., and Burrus, D.L., "Aerothermal Modeling,

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Phase 1,' 1983. (NASA CR-168296-VOL-l,-VOL-2)

Program - Phase I," GARRETT-21-4302, G a r r e t t Turbine Engine Co., Phoenix, AZ, 1982. (NASA CR-168031)

16. Sr in ivasan, R., Coleman, E., and Johnson, K., ' D i l u t i o n J e t Mix ing Program," GARRETT-21-4804, Garret t Turbine Engine Co., Phoenix, AZ, 1984. (NASA CR-174624)

17 . Sr in ivasan, R., Coleman, E., Meyers, G., and White, C., H D i l u t i o n J e t M ix ing Program," GARRETT-21-5418, G a r r e t t Turbine Engine Co., Phoenix, AZ, 1985. (NASA CR-174884)

18. Atkinson, W.H., and Strange, R.R., "Development o f Advanced High- Temperature Heat F lux Sensors," PWA-5723-27, P r a t t and Whitney A i r c r a f t , East Har t ford, CT, 1982. (NASA CR-165618)

19. Atkinson, W.H., Cyr, M.A., and Strange, R.R. , "Turbine Blade and Vane Heat F l u x Sensor Development Phase I," PWA-5914-21, P r a t t and Whitney A i r c r a f t , East Har t ford, CT, 1984. (NASA CR-168297)

20. Atkinson, W.H., Cyr, M.A., and Strange, R.R., "Development o f Advanced High Temperature Heat F lux Sensors, Phase 11," PWA-5914-39, P r a t t and Whitney A i r c r a f t , East Har t fo rd , CT, 1985. (NASA CR-174973)

21. Hyl ton, L.D., e t a l . , " A n a l y t i c a l and Experimental Evaluat ion o f t h e Heat Transfer D i s t r i b u t i o n Over the Surface o f Turbine Vanes," EDR-11209, D e t r o i t D iesel A l l i s o n , Ind ianapol is , I N , 1983. (NASA CR-768015)

22. Crawford, M.E., and Kays, W.M., "STAN5 - A Program f o r Numerical Computation o f Two-Dimensional I n t e r n a l and External Boundary Layer Flows," NASA CR-2742, 1976.

23. Turner, E., Wilson, M., Hyl ton, L., and Kaufman, R., ' 'Turbine Vane External Heat Transfer. Volume I - A n a l y t i c a l and Experimental Evaluat ion o f Surface Heat Transfer D i s t r i b u t i o n s w i t h Leading Edge Showerhead F i l m Cooling," EDR-11984-VOL-1, D e t r o i t D iesel A l l i son , I nd ianapo l i s , I N , 1985. (NASA CR-174827)

24. Yang, R.J., e t a l . , "Turbine Vane External Heat Transfer. Volume II- Numerical Solut ions o f t he Navier-Stokes Equations f o r Two- and Three- Dimensional Turbine Cascades w i t h Heat Transfer , ' EDR-11984-VOL-2, D e t r o i t D iesel A l l i s o n , I nd ianapo l i s , I N , 1985. (NASA CR-174828)

25. Anderson, O.L., "Assessment o f a 3-0 Boundary Layer Analysis t o P r e d i c t Heat Transfer and Flow F i e l d i n a Turbine Passage,ll R85-956834, Uni ted Technologies Research Center, East Har t fo rd , CT, 1985. (NASA CR-174894)

26. Dring, R.P., Joslyn, H.D., Hardln, L.W. and Wagner, J.H., IITurbine Rotor- S t a t o r In teract lon, l ' Journal o f Engineerinq f o r Power, Vol. 104, No. 4, O C t . 1982, pp. 729-742.

1 5

27. Patankar, S.V. and Spaulding, D.B., 'A Ca lcu la t i on Procedure f o r Heat, Mass and Momentum Transfer i n Three-Dimensional Parabo l ic Flows," I n t e r n a t i o n a l Journal o f Heat and Mass Transfer , Vol. 15, No. 10, Oct. 1972, pp. 1787-1 806.

28. Keyworth, 11, G.A., "Nat ional Aeronaut ica l R&D Goals, Technology f o r America's Futures,' Execut ive O f f i c e o f t he President, O f f i c e o f Science and Technology Po l icy , Washington, D.C., Mar. 1985. (PB86-209772)

16

TABLE I . . HOST Pro ject A c t i v i t i e s

Instrumentation Hot Section Yrewing system . . . . . . . . .

*Dynamic Gas Temperature Measurement System . A *Oynamic Gas Temperature Measurement System . B Turbine S t a t i c S t r a i n Gage . A Turbine S t a t i c S t ra in Gage . B *Turbine Heat Flux Sensors . . . . . . . . . . Laser Speckle S t ra in Measurement . . . . . . . High Temperature S t ra in Gage Mater ia ls Hot Section Sensors . . . . . . . . . . . . . *Laser Anemometry f o r Hot Section Applications HOST Instrument Applications . . . . . . . . .

. . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Contract (C). Grant (6). or

C C

Combusti on '*Assessment o f Combustor Aerothermal Models . I . . . . . . . . . . . . . . . . C *Assessment o f Combustor Aerothermal Models . I 1 . . . . . . . . . . . . . . . . C *Assessment o f Combustor Aerothermal Models . 111 . . . . . . . . . . . . . . . C Improved Numerical Methods . I . . . . . . . . . . . . . . . . . . . . . . . . . C Improved Numerical Methods . I 1 . . . . . . . . . . . . . . . . . . . . . . . . C Improved Numerical Methods . 111 . . . . . . . . . . . . . . . . . . . . . . . . G

Fuel Swir l Characterization . I . . . . . . . . . . . . . . . . . . . . . . . . C Fuel Swir l Characterization . I 1 . . . . . . . . . . . . . . . . . . . . . . . . C

Flow In te rac t i on Experiment . . . . . . . . . . . . . . . . . . . . . . . . . . C

Mass and Momenta Transfer . . . . . . . . . . . . . . . . . . . . . . . . . . . C Oiffuser/Combustor I n te rac t i on . . . . . . . . . . . . . . . . . . . . . . . . . C * D i l u t i o n J e t Mixing Studies . . . . . . . . . . . . . . . . . . . . . . . . . . C Latera l J e t I n jec t i on i n t o Typical Combustor Flowfields . . . . . . . . . . . . G Flame Radiat ion Studies . . . . . . . . . . . . . . . . . . . . . . . . . . . . N

Tiwhine b a t Transfer Mainstream Turbulence Influence on Flow i n a Turning Duct . A . . . . . . . . . *2-0 Heat Transfer wi thout F i lm Cooling . . . . . . . . . . . . . . . . . . . . *2-0 k a t Transfer w i t h Leading Edge Fi lm Cooling . . . . . . . . . . . . . . . 2-0 Heat Transfer w i t h Downstream F i lm Cooling . . . . . . . . . . . . . . . . . measurement o f Blade and Vane Heat Transfer Coef f ic ient i n a Turbine Rotor . . *Assessment o f 3-D Boundary Layer Code . . . . . . . . . . . . . . . . . . . . . *Coolant Side Heat Transfer w i t h Rotation . . . . . . . . . . . . . . . . . . . *Analy t ic Flow and Heat Transfer . . . . . . . . . . . . . . . . . . . . . . . . Ef fec ts o f Turbulence on Heat Transfer . . . . . . . . . . . . . . . . . . . . . Tip Region Heat Transfer . . . . . . . . . . . . . . . . . . . . . . . . . . . . Impingement Cooling . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Computation o f Turbine Blade Heat Transfer . . . . . . . . . . . . . . . . . . . Advanced Instrumentation Development . . . . . . . . . . . . . . . . . . . . . . Warm Turbine Flow Mapping w i t h Laser Anemometry . . . . . . . . . . . . . . . . Real Engine-Type Turbine k r o t h e n n a l Testing . . . . . . . . . . . . . . . . . .

Mainstream Turbulence In f luence on Flow i n a Turning Duct . B . . . . . . . . .

St ruc tu ra l Analysis Thermal /St ructura l l o a d Transfer Code . . . . . . . . . . . . . . . . . . . . . 3-0 I n e l a s t i c Analysis Methods . I 3-0 Component Speci f ic Modeling . . . . . . . . . . . . . . . . . . . . . . . . . . L ine r Cyc l ic L i f e Determination . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . I n e l a s t i c Analysis Methods . I 1

. . . . . . . . St ruc tu ra l Components Response Program . . . . . . . . . . . . . . . . . . . . . High Temperature Structures Research Laboratory . . . . . . . . . . . . . . . . Cons t i t u t i ve hodel Development . . . . . . . . . . . . . . . . . . . . . . . . . Cons t i t u t i ve Modeling f o r I so t rop i c Mater ia ls . I Cons t i t u t i ve Modeling f o r I so t rop i c Mater ia ls . I 1 Theoretical Const i tu t ive Models f o r S ing le Crystal Al loys

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Biax ia l Const i tu t ive Equation Development for Single Crysta ls and O i rec t i ona l l y

Fatigue and Fracture Creep-Fatigue- L i r e Predic t ion for I so t rop i c Materials . . . . . . . . . . . . . Elevated Temperature Crack Propagation . . . . . . . . . . . . . . . . . . . . . L i f e Predic t ion and Mater ia l Cons t i t u t i ve Behavior for Anisotropic Mater ia ls . . Pnalys is o f Fatigue Crack Growth Mechanism V i t a l i z a t i o n o f High Temperature Fatigue and Structures Laboratory

S o l i d i f i e d Al loys . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . Surface Protection

E f fec ts o f Surface Chemistry on Hot Corrosion . . . . . . . . . . . . . . . . . Thermal Ba r r i e r Coating L i f e Predic t ion . I . . . . . . . . . . . . . . . . . . Thermal E a r r i e r Coating L i f e Predic t ion . I 1 . . . . . . . . . . . . . . . . . . Thermal B a r r i e r Coating L i f e Predic t ion . 111 . . . . . . . . . . . . . . . . . A i r f o i l Deposit ion Model . . . . . . . . . . . . . . . . . . . . . . . . . . . . Mechanical Behavior of Thermal B a r r i e r Coatings . . . . . . . . . . . . . . . . Coating Oxidation/Oiffusion P red ic t i on . . . . . . . . . . . . . . . . . . . . . Deposit ion Model Ve r i f i ca t i on . . . . . . . . . . . . . . . . . . . . . . . . . Dual Cycle Attack . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Rig/Engine Corre la t ion . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Burner Rig Modernization . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Notes: * A c t i v i t i e s discussed i n t h i s repo r t A c t i v i t i e s i n se r ies A, B

I. I 1 . I11 A c t i v i t i e s i n p a r a l l e l

C

N N N N

NAS3-23154 NAS3-24228

....... NAG3-501 2510 2520/2530 2510

NAS3-23523 NAS3-23524 NAS3-23525 NAS3-24351 NAS3-24350 NAG3-596 NAS3-24350 h'AS3-24350 NAS3-24352 NAS3-22771

F33615-84-C-2427 NAS3-22110 NAG3-549 2780

NAS3-23278 NAG3-617 NAS3-22761 NAS3-23695 NAS3-24619 NAS3-23717 NAS3-23716 NAS3-23691

NAG3-522 NAG3-623 NSG3-075 NAG3-579 2640 2620 2640

NAS~-24358

NAS3-23272 NAS3-23697 NAS3-23698

5210 ~ ~ 5 3 - 2 3 6 8 7

5210 5210 5210 NAS3-23925 NAS3-23927 NAG3-511 NAG3-512

NAS3-23288 NAS3-23940 NAS3-23939 NAG3-348 5220

NAS3-23926 NAS3-23943 NAS3-23944 NAS3-23945 NAG3-201 NCC3-27 5120 5140 n 2 0 5120 5160

17

MATERIALS

c

EN VIR 0 N ME N CHARACTER-

CONSTITUTIVE FAILURE

STRUCTURAL

SUMMATION

Figure 1. - Integration of analytical models leads to l ife prediction of hot section parts.

COMBUSTOR LINER -Xc \ \ / I FLAME

r CAMERA -%iEw iNG LEN s

/ FUSED FIBERS FOR ’ IMAGING

r SINGLE FIBER FOR

Figure 2, - Combustor viewing system schematic (ref. 9).

18

ILLUMINATING r VIEW IN G LENS FIBER 7

/ I / I

\

\ L 12.7 mm dia

r 0 . 5 0 8 m m (0.020in) / dia TYPE B WIRE I I 1- F0-31!:5'') i / r 0.254 mm (0.010 in)

0.076 mm (0.003 in) dia TYPE B WIRE -------e-

0.381 mm (0.015 in) dia TYPE B WIRE -.

0.478 cm (0.188 in) dia COORS AD-99 CERAMIC STICK WITH FOUR ,' 0.078 cm (0.031 in) dia HOLES CERAMA-DIP

538 CEMENT OR SERMETEL P-1

Figure 4 - Sensor t ip geometry for dynamic gas temperature measurement system (ref. 5 and 10).

19

7 6 m (3 m i l ) T I C OUTPUT COMPENSATED 76 prn (3 m i l ) T I C OUTPUT

64OF RMS 800 281'F RMS 400 1

0 20 40 60 80 100 Time (ms) 0 20 40 60 80 100

Time (ms)

FIGURE 5 . - COMBUSTOR EXIT TEMPERATURE FLUCTUATIONS FOR SUBSCALE COMBUSTOR TEST (REFS. 5 AND 10).

76pm (3 mil) TIC output Compensated 76pm (3 mil) TIC output WITH TIME SCALE EXPANDED

393" RMS 1600 400 73' RMS

200 800

O F 0 "F 0

-800 -200

~ 1600

0 100 200 300 400 500 225 235 245 255 265 275 Time (ms) Time (ms)

FIGURE 6 . - COMBUSTOR EXIT TEMPERATURE FLUCTUATIONS FOR F-100 COMBUSTOR TEST (REFS. 5 AND 10).

20

L PINHOLE

(A) FRINGE-TYPE LASER ANEMOMETER SCHEMATIC (REF. 6). FIGURE 7.

(B) LASER ANEMOMETER AND OPEN-JET COMBUSTOR (RtF. 6). FIGURE 7. - CONCLUDED.

21

I

// I

P O 0 “ 4 0 0

I I 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0

0--0 0 0 0 0 r 0 0 0 0 0 0 n w J

8 II

W J CJ z a

22

0

m .- c

i o A

..

.- U

O Y-

I- t

E I-

E I-

23

EMBEDDED SWAGE WIRE THERMOCOUPLES

COLD / SIDE-. /

ELECTRICAL SCHEMATIC EMBEDDED THERMOCOUPLE SENSOR

'HOT SIDE

ALUMEL -1

2

HOTSIDE \ 3 ALUMEL

1-2 REFERENCE TEMPERATURE 1-3 SENSOR OUTPUT

(a) Schematic of embedded thermocouple sensor (ref. 5).

Figure 10.

ALUMEL

1 2

HOT SIDE

1-2 SENSOR OUTPUT 1-3 REFERENCE TEMPERATURE

I

HOT SIDE

(b) Schematic of Gardon gage sensor (ref. 5).

Figure 10. - Concluded.

AIRFOIL WALL

HEAT FLUX

(MWIM~)

fl 0’

/ /

l * O L - . 8

0 MEDTHERM RADIOMETER 0 TOTAL HEAT FLUX SENSOR

COMBUSTOR PRESSURE (ATMOSPHERES 1

Figure 11. - Typical test results from two types of heat f lux sensors located in a combustor l iner (ref. 20).

Figure 12. - Two airfoil profiles tested along with body-centered coordinate system grids generated as part of an inviscid blade-to-blade analysis (ref. 21).

25

UNMODIFIED STAN5 RESULTS

1.1

.I

.E

.4

.Z

a

1 .0

.I

. e

. 4

.2

e

I

.2 .4 .6 .I SUCTION

SUSfACE DISTANCE SlARC

MODIFIED STAN5 RESULTS

H/M0

.I .6 . 4 .z e

W 1 4 S 4512 D RUvleQ 4412 t RVNI4Q 4312

. z . 4 - 6 .I PRESSURE SUCTION

SWfACE DISTANCE W A R C

1.e

.I

.E

. 4

. 2

e

I, e

- 8

.a

. 4

.z

e

Figure 13. - STAN5 solutions compared with C3X airfoil experimental heat transfer coefficient data il lustrating the effects of varying Reynolds number (ref. 21).

P r e d i c t i o n ID 4400 STAN5 m d i f i e d ) 4400 STAN5 [ u m d i f iec

-------- 1.0 r

0 h, = 1135 W/d/C (200 B t u / h r / f t 2 / F ) I

c W

W .-

0.6 c W 0 V

L

2 0.4 cn c 0 L

L = 0.2 I"

'100 80 60 40 20

Data ID Re = 1 . 9 9 ~ 1 0 ~ 1 .o

0 I \ L 0.8 L

c W

0 .-

0.6 E W 0 u L

0.4 2 cn c 0 L L

0.2 = z

n 0 20 40 60 80 100"

P e r c e n t sur face d i s t a n c e , S Figure 15. - Comparison of predicted suction and pressure surface heat transfer coefficient

distributions using run 4400 STAN5 input data streams for showerhead cooled C3X vane (ref. 23).

27

.2

Figure 17. - Comparison of measured and calculated Stanton Number distributions for end wall surface (ref. 25).

0 DATA ANALYSIS -

I I I I I I I I I

28

ROTATION

Figure 18. - Large low speed turbine stage (ref. 26).

Figure 19, - A typical cooled aircraft gas turbine blade.

LLEA#IG SURFACE WIlH ROTATIW 0.5 I I

GUARD OUTWARD FRST NWARD SECtMD STRAIGHT TURN STRAIGHT TMN

Figure 20. - Effect of rotation on coolant passage heat transfer,

29

1. Report No.

NASA TM-88932

Toward Improved Durabi lity in Advanced Combustors and Turbines - - Progress in the Prediction o f Thermo- mechanical Loads

2. Government Accession No.

7. Author@)

Daniel E. Sokolowski and C. Robert Ensign

19. Security Classif. (of this report)

Unclassi f ied

9. Performing Organization Name and Address

National Aeronautics and Space Administration Lewis Research Center Cleveland, Ohio 44135-3191

20. Security Classif. (of this page) 21. No of pages 22. Price'

Unclassified 30 A0 3

2. Sponsoring Agency Name and Address

National Aeronautics and Space Administration Washington, D.C. 20546-0001

5. Supplementary Notes

3. Recipient's Catalog No.

5. Report Date

6. Performing Organization Code

8. Performing Organization Report No.

E-3374 10. Work Unit No.

533-04-1 1 11. Contract or Grant No.

13. Type of Report and Period Covered

Technical Memorandum 14. Sponsoring Agency Code

Prepared for the 31st International Gas Turbine Conference and Exhibition, sponsored by the American Society of Mechanical Engineers, Dusseldorf, West Germany, June 8-12, 1986.

6. Abstract

NASA is sponsoring the Turbine Engine Hot Section Technology (HOST) Project to address the need for improved durability in advanced combustors and turbines. Analytical and experimental activities aimed at more accurate prediction of the aerothermal environment, the thermomechanical loads, the material behavior and structural responses to such loading, and life predictions for high temperature cyclic operation have been underway for several years and are showing promising results. Progress is reported in the development of advanced instrumentation and i n the improvement of combustor aerothermal and turbine heat transfer models that will lead to more accurate prediction of thermomechanical loads.

17. Key Words (Suggested by Author(s))

HOSl; Durability; Turbine engine; Aero- thermal loads; Thermomechanical loads; Instrumentation; Combustor; Turbine; Turbine heat transfer data

18. Distribution Statement

Unclassified - Unlimited Subject Category 07

*For sale by the National Technical Information Service, Springfield, Virginia 22161 NASA FORM 1626 OCT 86

toward improved durability in advanced combustors and ... · reduce material temperatures, (3)...

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