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Reports quoted are not necessarily available to members of the public or to commercialorganisations.

Ac6454a For

D u i t I Or

DEFENCE RESEARCH AGENCY Av1ilblity oAVail axd/o-r

Aerospace Division Dist

RAE Pyestock

Technical Memorandum P 1223

Received for printing 24 February 1992

MEASUREMENTS AND COMPUTATIONS OF EXTERNAL HEATTRANSFER AND FILM COOLING IN TURBINES

by

S. P. HarasgamaC. D. Burton

K. S. Chana

SUMMARY

A review of recent work on turbine heat transfer performed at the RAE (Pyestock) ispresented. The work covers the effects of secondary flows on turbine nozzle guide vaneheat transfer with and without film cooling. It is shown that the heat load to the platforms(endwalls) are significantly affected by the secondary flow action. The platform filmcooling data has been well correlated with flat plate single row film cooling data to within±11%. A three-dimensional Navier-Stokes computational study of the effects of turbineinlet temperature distortion on the thermo-fluid mechanics within a rotating blade passage isgiven. It is shown that the temperature distortion is modified within the rotor blade and canlead to increased pressure side and over tip heat transfer.

A paper presented at the tenth International Symposium On Air Breathing Engines(ISABE) 1-6 September 1991, Nottingham, UK.

Copyright

Controller HMSO London1992

2

LIST OF CONTENTS

Page

INTRODUCTION 3

TEST FACILITY 4

TURBINE STATOR HEAT TRANSFER RESULTS 4

TURBIRE STATOR FILM COOLING RESULTS 6

HOT GAS MIGRATION WITHIN ROTOR BLADES 8

CONCLUSIONS AND OBSERVATIONS 9

Acknowledgments 10

References 10

Illustrations Figures 1-19

Report documentation page inside back cover

MEASURIENTS AND CXOPUTATIONS OF EXTERNAL HEATTRANSFER AND FILM COOLING IN TURBINES

S.P. Harasgama, C.D. Burton & K.S. ChanaRoyal Aerospace EstablishmentPropulsion Department, Pyestock,Farnborough, Hampshire, U.K.

INTR0DUCTIONABSTRACT Turbine heat transfer research has beenA review of recent work on turbine heat undertaken at the RAE in both experimentaltransfer performed at the RAE (Pyestock) and computational fields. In the main, theis presented. The work covers the effects work has concentrated on the heat transferof secondary flows on turbine nozzle guide resulting from the three-dimensional (3D)vane heat transfer with and without film nature of the flow field within highcooling. It is shown that the heat load to pressure turbine stators and rotors. Thesethe platforms (endwalls) are significantly secondary flows have been well catalnuedaffected by the secondary flow action. The in terms of their aerodynamic impact , .platform film cooling data has been well The endwall inlet boundary layers givecorrelated with flat plate single row film rise to two sets of secondary flows. Thecooling data to within ± 11% . A Three- passage vortex is caused by overturning ofDimensional Navier-Stokes computational the endwall boundary layers within thestudy of the effects of turbine inlet turbine passage. In addition, a horse-shoetemperature distortion on the thermo-fluid vortex is formed by the stagnation of themechanics within a rotating blade passage endwall boundary layers at the bladeis given. It is shown that the temperature leading edge. Both mechanisms cause adistortion is modified within the rotor migration of flow towards the suction sideblade and can lead to increased pressure of the passage and can lead to increasedside and over tip heat transfer. losses.

Heat transfer studies on the effects ofsecondary flow hIve also been reported by

NOMENCLATURE several authors -7. It is shown that thehorse shoe vortex has a major effect on

A Film cooling superposition parameter the platform heat transfer and also on theB Film cooling superposition parameter vane suction side. The horse shoe vortex-B/A Cooling Effectiveness strips off the incoming boudary layer andCt Vane tangential chord this gives rise to a new boundary layerd Cooling hole diameter starti.,g just downstream of the separationG Blowing rate (riUi/rgUg) line. This results in enhancing the heath Heat transfer coefficient load in the downstream region.I Momentum flux ratio (riui2 /rgug 2 )k Thermal conductivity of air The heat transfer effects due to radialM Mach number variations in inlet temperature areNu Nusselt number brought about by hot streaks genezatedQ Heat Flux within the combyttor and is essr-lially anRe Reynolds number unsteady effect . Early analytijal work 14

S Streamwise dimension has indicated that density sf:tificationS2 Pitchwise dimension in a rotating pasage can gi- e rise toT Temperature secondary flows due to Cor-olis andU Velocity Centripetal accelarations affecting theX Axial dimension flow. Recent two-dimenpional viscousr Density numerical simulations-, 6 show that thee Temp. Diff. Ratio (Tg-Ti)/(Tg-Tw) hot gas is transported within the blade0 Cooling function parameter passage and is deposited onto the bla $

pressure side.Vis.ous 3D computations'",18

Subscripts additionally sh-,w that hot gas migratescp Corrected for property variations radially leading to enhanced rotor tipg Gas path value heat transfei.i Injectant valueo Datum (uncooled) value This paper reviews the work done at thet Total (stagnation) value RAE on experimental and computational heatw Wall value transfer in turbines. Detailed results of

heat transfer on a fully annular geometry

TM P 1223

*4

of the aerofoils and platforms of turbine number of thin-film gauges were placed onnozzle guide vanes along with flow the vanes to give accurate measurements ofvisualisations which corroborate the heat the heat transfer distributions. In alltransfer distributions are presented. Full cases aerodynamic Jata were acquired atcoverage film cooling data on the endwall the same locations as the heat transferof a high pressure vane are also presented measurements. This was achieved by havingand correlated with axial pressure static pressure tappings on the vanes andgradient. A computational skudy using afully 3D Navier Stokes code on theeffects of combustor generated temperaturedistortion within a rotating turbine bladepassage is also reviewed.

TEST FACILITY

The Isentropic Light Piston Cascade (ILPC)as shown in Figure 1 is derived from theshort duration thermc-fluids facilitiespioneered at Oxford University . Thefacility operates by rapidly compressingair with a light piston which raises itspressure and temperature to a requiredlevel. Subsequently a fast acting valveis opened and the air is dischargedover the set of turbine vanes/blades. Thecompression time of the facility is 1.0sec Fig 2 Heat Transfer & Cooling Bladewith a run time of 0.5sec when operatingat a gas-to-wall temperature ratio of 1.5.Engine similarity conditions of Reynolds total pressure probes upstream of thenumber and Mach number prevail for the turbine module. The inlet free stream0.5sec run time. For film cooling studies, turbulence intensity was set to around 6%a mixture of air and carbon-dioxide (CO2 ) by a bar grid placed 4.5 chords upstreamwas used to obtain the correct of the cascade.coolant-to-gas density ratio whilstmaintaining the correct gas-to-wall TURBINE STATOR HEAT TRANSFER RESULTStemperature ratio.

The heat transfer to the aerofoils suctionGATE vAES(4 0" surface is shown on Figure 3. The NusseltORS":NG IKE" number (Nu) contours indicate the extent

'4OA ,£ I ... A,*of three-dimensionality of the heat loads.For all Reynolds numbers and Mach numbers

Sc A' NO ,s .S the secondary flow has transported lowenergy endwall boundary layer flow onto

______ _ '_ the suction surface. These boundary layersZ " E= are relatively thick and thereby reduce

'CP," the root and tip heat transfer rates. TheA- E .0flow visualisation of Figure 4 tends to

- 1 corroborate the heat transfer patterns. ItV can be seen that flow migrating accross

.. the suction surface appears to come fromthe endwall boundary layers.

j'pJ X,9f Platform Nusselt number contours are shownon Figure 5. The heat transfer patterns

VAC, are significantly non-uniform and this/ M again is attributed to the horse-shoe

vortex and the passage secondary flow. ItRoA ARE A,*.VAC' can be seen that as the Reynolds number is

increased the high heat transfer region

migrates accross the endwall and movesPI CLPC : T,,t Fxlity upstream. A similar effect is observed as

the Mach number is increased. In all casesthe increase in Reynolds number results in

The vanes under test are manufactured from an increase of Nusselt number as expected.machineable glass ceramic (MACOR ) which This trend is the opposite for increasinghas a low thermal diffusivity. Thin film Mach number. The high heat transfer ratesheat transfer gauges are deposited onto in the downstream portion of the passagethe surface of the vanes and monitor the are caused by the horse shoe vortex whichtemperature during the run time. The heat strips off the incoming boundary layer andtransfer coefficients are obtained by causes a thin new , high energy, boundarysolving the one-dimensional transient heat layer to form.conduction equation. A vane with typicalinstrumentation and platform film cooling The fli visualisation of Figure 6 showsholes is shown on Figure 2. A sufficient the nature of the secondary flow on the

TM P 1223

I 3.46 It 3.416i 14 3.416.094 I 1.29

M P3800 P424M030a 03200 033o 1M Li

12012000 . ..

1420 14U:j H2200

G? 900 lo

h* 17E6 IV: 3,46

H4 2200

E 600 to I0 400 120C1200 0 MO0)

P 30000 3200

Fig 3 Suction Side Nusselt Numbers

vane platform. The horse shoe vortex canbe clearly seen to start just upstream ofthe leading edge and traverse accross thepassage until it meets the suction surfaceat around 40% axial chord. A computationwas performed with the 3D Navier Stokescode to predict the secondary flow fieldand this is shown on Figure 7. The codewas run with 25 radial, 83 axial and 25blade-to-blade grid points with an inletboundary layer thickness of approximately5% annulus height on both endwalls.Computations were performed with the codein the fully turbulent mode throughout. Itcan be seen that the agreement betweenprediction and experiment is qualitatively

Fig 4 Suction Side Flow Vis. very good. These predictions yield furtherinsight into the nature of the secondaryflows and corroborate the heat transferresults.

NI 3416 Re 3.41

J400 J 00

1)"0/~ 3100

S3100 . 6 1300 40602200 11271

62600 :920060

02200 .Tr B' . 4l DJK i.i e= .l

010011.4 AIo '

t _ _ _ _ _ _ I_ _ _ _.I_ _ _ _ __"_ _ __I , , ~lS .7E6 *e3 416

1400

Fig 5 Hub Nusselt Numbers

TM P 1223

, -. 6

- In order to achieve these conditions aheavy gas, carbon-dioxide, was mixedwith air to simulate density ratio's upto 2.0 .The coolant was injected using aseparate reservoir with a fast actingvalve which fed a plenum just below thesurface of the platforms being cooled. Thecoolant was injected synchronously with,the onset of main stream flow.

The Nusselt number contours on the cooledplatform are shown on Figure 8. These dataare for increasing blowing rate (G) from0.53 to 1.11 for a fixed 8 of 1.62, alsoshown is the datum (uncooled) result. Itcan be seen that as G is increased theNusselt numbers decrease significantly.Figure 9 shows a similar Nusselt numberpattern for G = 1.11 and 8 increasingfrom 0.87 to 1.62. Here again Nu

..... ...... decreases with increasing e. In both.-. .... figures it is clear that the suction side

/ "of the platform is receiving more cooling. than the pressure side. This is because

the secondary flow convects the coolanttowards the suction side which is

H 200C evident from the flow visualisation of. °'1 .. •G 180F 1600 Figure 6. Figure 10 shows the ratio ofo 00 cooled-to-uncooled Nusselt numbers for

C 1000 G = 1.11 £ E = 1.44, it is clear that over" • " i ' ;8 801

A 600 most of the platform the heat load isreduced by 75% whilst at the pressure sidecorner the reduction is about 20%.

, ,q C .gs5 .., t.K s -oIt ,1 These data have been analysed using theAnd rtilt*1 N....lt . 10 + lm co lsuperposition approach to film cooling.All data have been normalised to constantproperty conditions1 using:

TURBINE STATOR FILM COOLING RESULTS(Nu/NuO)cp = (Nu/Nuo ) * (Tw/Tg)n 3

Experimental data on film cooling testswithin a fully annular turbine stator vane The superposition representation is:passage are reported. The tests were doneon the outer platform of the vane. These (Nu/Nuo)cp = A + Be 4tests were performed to corroborate adesign philosophy of placing cooling holes Where n = -0.25 for a turbulent boundaryon an iso-Mach line. Such a design will layer. A and B are constants ofgenerate a uniform blowing rate accross proportional~ty in the superpositionthe whole passage. Consider the Momentum formulation 1 -

Flux Ratio :Figure 11 shows the linearity of the

I = (i*Ui2)/(g*Ug 2 ) 1 energy equation when applied to these dataand indicates that the superposition

I = (Pti-Pi)/(Ptg-Pg) 2 formulation holds even in the presence ofstrong secondary flows. This enables the

In the turbine situation the free stream present data to be extrapolated to otherand cooling total pressures are fixed by engine operating conditions. Also showncompressor exit conditions, as is the are the platform cooling effectiveness oncooling static pressure. Hence the only Figure 12a&b. Figure 12a shows thevariable is Pg, which is constant along an streamwise variation of effectivenessiso-Mach line. Placing the film cooling which indicates that there is a gradualholes along such a line will yield a reduction of -B/A in the downstreamconstant momentum flux ratio, which for a direction as the coolant mixes with thefixed density ratio yields a uniform mainstream gas. It also indicates ablowing rate(G). reduction of -B/A from suction to pressure

side which is more explicitly presented inFor the present tests the vane Reynold; Figure 12b. This indicates that thenumber was set to approximately 2.6x 10b suction side effectiveness is greater thanbased on exit conditions and tangential the pressure side value for any axialchord. The Exit Mach number was 0.93 and distance due to secondary flow effects.the gas-to-wall temperature ratio (Tg/Tw)was 1.30. The engine cooling design The present data have been successfullyrequired a density ratio of 1.8 and a correlated with single row film coolingtemperature difference ratio (0) of 1.75 . results on a flat plate with zero axial

TM P 1223

Datum Uncooled G 0.52 )A 8008 1200o 1600D 2000E 2400F 2800G 3200

G =0.83 G =1.11

Fig 8 Casing Nusselt Numbers for 8=1.62, All G

A 800e 0.87 e =1.22 B 1200C 1600

_________________________D 2000

E 2400F 2800G 3200

0 1.44 e 1.62

Fig 9 Casing Nusselt Numbers for G=1.11, All 8

*MACOR -Trademark of Corning, U.S.A

TM P 1223

L8

pressure gradient. This was achieved bymodifying the present data with acorrection for the turbine pyessuregradient. It has been shown" that theratio Nu/Nu0 can be put as:

(Nu/Nuo)cp = (S/d) x (Ui/Ug)-4/3 5

A 00 A modification was made to the prestnJB 01 data by multiplying it by (U/Uinlet)

'

,co2 and plotting against Equation 5. The valueD 03 n=0.2 was chosen by arguing that, usually,E 04 heat transfer coefficients can be put asO.6 functions of Reynolds number (velocity)

0 raised to the power 1/5th or 4/5th. The10.6 results of this analysis are shown on

K 09 Figure 13a&b. Figure 13a shows the dataplotted against the discrete single rowinjection parameter (Eq-5). It is seen

rig 10 A.tio ot Coo. to unoold ft... t Nob*r that (Nu/Nuo)cp lies below the flat plateForT C *.11 £ 1.44 zero pressure gradient correlation. The

corrected data are shown on Figure 13b. Itcan be seen that Equation 5 now representsthe platform data very well. The scatterabout the line is only ± 11%, which isconsidered to be very good in the presentcase of large secondary flows.

0It Wn -t1.,.%*t ~

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

1.. (

.. Y

- ~ .V*hZ fo .... I. ,tI;¢UU

/ I i

.. .L, ., 0.ltl:""I

HOT GAS MIGRATION WITHIN ROTOR BLADES

The effect of inlet temperature distortionon the thermo-fluid mechnics in the rotorblade has been computed 1 using a steady

T stats three-dimensional Navier-Stokes" " ., t ,a -,-".. .,,- -.. .,e ,m-m,--. n *, " code . The inlet temperature distortion is

generated by the combustor burners andaffects the rotor blading as they passthrough these hot patches; it is thereforean unsteady effect. However, an "average"effect of the hot gas migration can be

TM P 1223

evaluated using a steady state 3D viscouscomputation in order to gain anunderstanding of the flow physics. . -47

The simulation of inlet temperaturedistortion was achieved with a parabolic ..... P RAlOl)

radial temperature profile as shown on ".Figure 14. The computational grid for therotor blade is shorn on Figure 15. The 3DNavier-Stokes code was run with 67-axial, f.25-radial and 25-blade-to-blade grid .. ,points. The predicted secondary flow

Figure 16 SOcondary Fl- Vector. at 30% Axial Chord

cContours of rotor relative totaltemperature are given in Figure 18. This

0.9- shows that the hot gas migrates from thei RO mid-span region at turbine inlet to the

Pb presssure side tip of the blade by 80%8axial chord. The hot gas subsequently

enters the tip gap and is entrained intothe tip vortex. The heat transferenhancement on the blade pressure side dueto hot gas migration is shown on Figure19. This indicates a 75% enhancement dueto hot gas migartion. The actual increase

0 20 40 50 s0 I00 will be periodic and fluctuate between 0%Root % BLADE HEIGHT Tp and 75% as the blade passes through the

hot gas patches.Fig 14 Inlet Radial Temperature CONCLUSIONS & OBSERVATIONS

Full three-dimensional heat transferdistributions have been measured onturbine vanes operating at enginerepresentative conditions. The effects ofsecondary flow on heat tranfer are large.Rotor blade heat transfer is significantlyaffected by inlet temperature distortion.The main conclusions are as follows:

- The secondary flow reduces the root andtip heat transfer on the vane suctionside due to low energy endwall fluidmigrating from the endwall regions.

- Endwall heat transfer is also affectedby secondary flow leading to high heattransfer near the pressure side trailingedge.

- Flow visualisations indicate that theFtgsl1.". coputtoi1 Grid horse-shoe vortex sweeps off the

incoming endwall boundary layer.This leads to high energy free streamgas forming a thin new boundary layer,

patterns are given on Figure 16. It can be giving high heat transfer.seen that the computations with radialtemperature distortion give higher - Platform film cooling data indicate goodsecondary flows. Figure 17 shows the over effectiveness when the injectant emergestip Mach number distributions which reach into a uniform pressure field. This dataa peak value of 1.80 and this will also has been successfully correlated to ±11%enhance the rotor tip heat transfer rates. of flat plate results.

TM P 1223

10

Pressure Sirface41 1, ' 0, 2--- -

G6P r0

6 . ..DA. .N

-A. 10SPA

0.000 20.000 40.000 60.000 00.000 tO0.0

I AXIAL Ct40 D

Fig 19 Ratio of Heat Flux With RTD1'9,r, , Mch N,,b~r. In Tip Isp. A I vj., To Without RTD

- Rotor blade pressure side heat transfer 3. Gaugler R.E. & Russell L.M. (1983),is increased due to hot gas migration. "Comparison of visualised turbine endwallThe maximum increase can reach 75% as secondary flows and measured eat transferthe blades pass periodically through hot patterns". NASA Tech. Memo. TM83016.gas patches generated by the combustor.

4. Boyle R.J. & Russell L.M. (1989),ACKNOWLEDGEMENTS "Experimental determination of stator

endwall heat transfer". ASME PaperThanks are due to Mr K.J. Walton who Number:89-GT-219.maintained the test facility. Theturbomachinery heat transfer group at 5. Wedlake E.T.,Brooks A.& Harasgama S.P.Oxford University led by Professor Terry (1989), "Aerodynamic and heat transferJones were instrumental in launching the measurements on a transonic nozzle guideRAE effort, their invaluable advice and vane", Trans. ASME, Jnl. of Turbomachineryencouragement are gratefully acknowledged. Vol 111, pp36-42.

REFERENCES 6. Harasgama S.P. & Wedlake E.T. (1990)

"Heat transfer and aerod. jamics of a high1. Langston L.S., Nice M.L. & Hooper R.M. rim speed turbine nozzle guide vane tested(1977), "Three-dimensional flow within a in the RAE Isentropic Light Piston Cascadeturbine blade passage", ASME Jnl. of Eng. (ILPC)". ASME Paper Number:90-GT-41, to befor Gas Turbines & Power, Vol-99,pp2l-28. published in ASME Jnl. of Turbomachinery.

2. Sieverding C.H. (1985),"Recenit progress 7. Jones T.V., Harvey N.W., Ireland P.T. &in the understanding of basic aspects of Wang Z (1989) "Deta-led heat transfersecondary flows in turbine blade passages" measurements in nozzle guide vane passagesASME Jnl of Engineering for Gas Turbines in linear and annular cascades in theand Power, April, Voi-107, pp248- 2 5 7. presence of secondary flows".AGARD-CP-422

PS 30% ea 140s: 80 "SS5 50%

oo 'I80

?,w,. I* Total ?m~paratu, Coftors

TM P 1223

8. Dawes W.N. (1986). "A numerical methodfor the analysis of 3D viscouscompressible flow in turbine cascades;Application to secondary flow developmentin a cascade with and without dihedral",ASME paper number: 86-GT-145

9. Schultz D.L. & Jones T.V. (1973), "Onthe flow in an isentropic light pistontunnel". ARC, R&M-3731.

10. Metzger D.E., Carper H.J. & Swank L.R.(1968), "Heat transfer with film coolingnear non-tangential injection slots", ASMEJnl. of Eng. for Power, April.

11. Jones T.V., Forth C.J.P., Fitt A.D. &Robertson B.A. (1986), "Temperature ratioeffects in compressible boundary layers",Irt. Jnl. of Heat & Mass Transfer, Vol 29

12. Jones T.V. (1986), "The scaling offilm cooling, theoretical & experimentalresults", VKI-LS-1986-06.

13. Butler T.L., Sharma O.P., Joslyn H.D.& Dring R.P. (1986), "Redistribution of aninlet temperature distortion in an axialflow turbine". AIAA Paper: AIAA-86-1468.

14. Hawthorne W.R. (1974), "Secondaryvorticity in stratified compressiblefluids in rotating systems", Univ. ofCambridge, CUED/A-TURBO/TR63.

15. Rai M.M & Dring R.P. (1987), "Navier-Stokes analyses of the redistribution ofinlet temperature distortion in a turbine"AIAA Paper: AIAA-87-2146.

16. Krouthen B. & Giles M.B. (1988),"Numerical investigation of hot streaks inturbines". AIAA Paper: AIAA-88-3015.

17. Harasgama S.P. (1990), "Combustor exittemperature distortion effects on heattransfer and aerodyamics within a rotatingturbine blade passage", ASME No:90-GT-174

18. Sharma O.P., Pickett G.F & Ni R.H.(1990), "Assessment of unsteady flows inturbines", ASME Paper No: 90-GT-150

TM P 1223

%0

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